Handbook of Composites, 2nd Edition

HANDBOOK OF COMPOSITES SECOND EDITION Edited by S.T.Peters Process Research, Mountain View, Calfornia, USA CHAPMAN & H...

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HANDBOOK OF COMPOSITES SECOND EDITION Edited by

S.T.Peters Process Research, Mountain View, Calfornia, USA

CHAPMAN & HALL

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London Weinheim . New York Tokyo Melbourne Madras

Published by Chapman & Hall, an imprint of Thomson Science, 2-6 Boundary Row, London SE18HN, UK Thomson Science, 2-6 Boundary Row, London SE18HN, UK Thomson Science, 115 Fifth Avenue, New York, NY 10003, USA Thomson Science, Suite 750,400 Market Street, Philadelphia, PA 19106, USA Thomson Science, Pappelallee 3,69469 Weinheim, Germany

First edition 1982 Second edition 1998

0 1998 Chapman & Hall Thomson Science is a division of International Thomson Publishing Typeset in 10/12 pt Palatino by GreenGate Publishing Services, Tonbridge, England Printed in Great Britain by Cambridge University Press ISBN 0 412 54020 7 All rights reserved. No part of this publication may be reproduced, stored in a retrieval system or transmitted in any form or by any means, electronic,,mechanical, photocopying, recording or otherwise, without the prior written permission of the publishers. Applications for permission should be addressed to the rights manager at the London address of the publisher. The publisher makes no representation, express or implied, with regard to the accuracy of the information contained in this book and cannot accept any legal responsibility or liability for any errors or omissions that may be made. A catalogue record for this book is available from the British Library

CONTRIBUTORS

SURESH G. ADVANI Department of Mechanical Engineering, University of Delaware, Spencer Laboratory, Newark, DE 19716, USA MAURICE E AMATEAU Applied Research Laboratory, Pennsylvania State University, PO Box 30, State College, PA 16804, USA

EVER J. BARBER0 315 Engineering Science Building, West Virginia University, Morgantown, WV 26506-6106, USA A.I. BEIL' Institute of Polymer Mechanics, Latvian Academy of Sciences, 23 Aizkraukles Street, Riga LV-1006, Latvia JEROME S. BERG True Temper Sports, 5421 Avenida Encinas, Suite G, Carlsbad, CA 92008, USA

KENNETH R. BERG Riggs Corporation, 837 Agate Street, Medford, OR 97501, USA LARS A. BERGLUND LuleH University of Technology, SE-97187 LuleH, Sweden D. BROWN Boeing Commercial Airplane Group, Douglas Products Division, Mail Stop D001-0018, Long Beach, CA 90846, USA JOHN D. BUCKLEY 23 East Governor Drive, Newport News, VA 23602, USA JERRY L. CADDEN C & S Technologies, 42759 Mountain Shadow, Murrieta, CA 92562, USA ZHONG CAI(deceased) 4180 Berkeley Creek Drive, Duluth, GA 30136, USA

x

Handbook of composites

FRANK A. CASSIS FAC Associates, 1150 N. Mountain, Suite 1028, Upland, CA 91786, USA

MIRIA M. FINCKENOR EH12 Bldg 4711, Marshall Space Flight Center, AL 35812, USA

LINDA L. CLEMENTS C & C Technologies, PO Box 1089, Dayton, NV 89403, USA

LIHWA FONG BLK G 5, Nanyang Avenue, Singapore 63616

DOUGLAS L. DENTON Chrysler Corporation, CIMS 482-00-13, 800 Chrysler Drive, Auburn Hills, MI 48326-2757, USA EDDY A. DERBY Composite Optics, 9617 Distribution Ave, San Diego, CA 92121, USA GEORGE W. DU Principal Engineer, 16331 Bay Vista Drive Cleanvater, FL 34620, USA HARRY W. DURSCH Boeing Defense and Space Group, PO Box 3999, Mail Stop 73-09, Seattle, WA 98124-2846, USA DON 0. EVANS Cincinnati Milacron, 4701 Marburg Avenue, Cincinnati, Ohio 45209, USA

HUGH H. GIBBS Polycomp Consulting, Inc., 25 Crestfield Road, Wilmington, DE 19810, USA TIMOTHY GUTOWSKI Department of Mechanical Engineering, Massachusetts Institute of Technology, Bldg 35-234, Cambridge, MA 02139, USA RICHARD N. HADCOCK 6 Sue Circle, Huntington, NY 11743, USA ENAMUL HAQUE Azdel, Inc., Technology Center, 658 Washburn Switch Road, Shelby, NC 28151-2284, USA

L.J. HART-SMITH Boeing Commercial Airplane Group, Douglas Products Division, Mail Stop D800-0019, 4000 Lakewood Boulevard, Long Beach, CA 90846, USA

Confributors xi JENNIFER HETH Cytec Fiberite, 501 W. Third Street, Winona, MN 55987-2854, USA

VALERY I. KOSTIKOV Niigrafit Institute, 2 Electrodonaya Street, Moscow, 111524, Russia

THOMAS S. JONES Industrial Quality, Inc., 640 E. Diamond Ave., Suite C, Gaithersburg, MD 20877, USA

GARY C. KRUMWEIDE Composite Optics, 9617 Distribution Avenue, San Diego, CA 92121, USA

THOMAS JUSKA Naval Surface Warfare Center, Carderock Division, Structures and Composites Department, Bethesda, MD 20084-5000, USA

V.L. KULAKOV Institute of Polymer Mechanics, Latvian Academy of Sciences, 23 Aizkraukles Street, Riga LV-1006, Latvia

JOHN T. KANNE 2201 Johnson Road, Memphis, TN 38139, USA

KHALID LAFDI Center for Advanced Friction Studies, Southern Illinois University at Carbondale, Carbondale, IL 62901-4343, USA

HARRY S. KATZ Utility Development Corporation, 112 Naylon Avenue, Livingston, NJ 07039, USA

CHRISTY KIRCHNER LAPP 1412 Bellingham Way, Sunnyvale, CA 94087, USA

V.S. KILIN Niigrafit Institute, 2 Electrodonaya Street, Moscow, 111524, Russia FRANK K. KO Drexel University, Fibrous Materials Research Laboratory 27-439, Philadelphia, PA 19104, USA KENT E. KOHKONEN Brigham Young University, 435 CTB Technology Department, Provo, UT 84602, USA

ROBERT A. LATOUR Clemson University, Clemson, SC 29634, USA BURR L. LEACH Cambridge Industries, 1700 Factory Avenue, Marion, IN 46952, USA STEWART N. LOUD Composites Worldwide Inc., 991 Lomas Santa Fe Drive, C469, Solana Beach, CA 92075-2125, USA

xii Handbook of composites VICKI P. MCCONNELL Ray Publishing, Independence Street, Suite 270, Wheat Ridge, CO 80033, USA

NITIN POTDAR Brigham Young University, 435 CTB Technology Department, Provo, UT 84602, USA

ANDREW C. MARSHALL Marshall Consulting, 720 Appaloosa Drive, Walnut Creek, CA 94596, USA

KENNETH REIFSNIDER Virginia Polytechnic Institute and State University, Patton Hall 120, Blacksburg, VA 24061-0219, USA

ANTHONY MARZULLO 39 Harold Street, COSCob, CT 06807-2132, USA

THEODORE J. REINHART 345 Forrer Boulevard, Dayton, OH 45419-3238, USA

DONALD W. OPLINGER Federal Aviation Administration, Wm. J. Hughes Technical Center AAR-431, Atlantic City, International Airport, NJ 08405, USA

PAUL E SADESKY C & S Technologies, 23547 Mountain Court, Murrieta, CA 92562, USA

HARRY E. PEBLY 198 Center Grove Road, Randolph, NJ 07869, USA LYNN S. PENN Department of Chemical and Materials Engineering, 177 Anderson Hall, University of Kentucky, Lexington, KY 40506-0046, USA S.T. PETERS Process Research, 925 Sladky Avenue, Mountain View, CA 94040-3625, USA

FRANK J. SCHWAN 36671 Montecito Drive, Fremont, CA 94536, USA ANTON L. SEIDL 18941 Mellon Drive, Saratoga, CA 95070, USA JOCELYN M. SENG Owens Corning Science and Technology Center, 2790 Columbus Road, Granville, OH 43023-1200, USA SHALABY W. SHALABY Clemson University, 301 Rhodes Res., Clemson, SC 29634, USA

Contributors xiii DAVID A. SHIMP PO Box 974, Prospect, KY 40059, USA DONALD R. SIDWELL 44609 Grove Lane, Lancaster, CA 93534-2833, USA BRIAN E. SPENCER Spencer Composite Corporation, 3220 Superior Street, PO Box 4377, Lincoln, NE 68504-0377, USA ROBERT C. TALBOT 7199 Lorine Court, Columbus, OH 43235-5125, USA YU.M. TARNOPOL'SKII Institute of Polymer Mechanics, Latvian Academy of Sciences, 23 Aizkraukles Street, Riga LV-1006, Latvia R.C. TENNYSON University of Toronto, Institute for Aerospace Studies, 4925 Dufferin Street, Downsview, Ontario, Canada M3H 5T6 JAMES L. THRONE Shenvood Technologies, Inc., 158 Brookside Boulevard, Hinckley, OH 44233-9676, USA FRANK TRACESKI Department of Defense, 5203 Leesburg Pike Suite 1403, Falls Church, VA 22041, USA

WAYNE C. TUCKER Naval Undersea Warfare Center, PO Box 86, Exeter, RI 02822, USA V. V. VASILIEV Moscow State University, 14-1-110 Podolskih Kursantov Street, Moscow 113545, Russia DENNIS J. VAUGHAN 146 Longview Drive, Anderson, SC 29621, USA H. WANG Department of Chemical and Materials Engineering, 177 Anderson Hall, University of Kentucky, Lexington, KY 40506-0046, USA ANN E WHITAKER EHOl Bldg 4612, Marshall Space Flight Center, AL35812, USA BRIAN A. WILSON Wilson Composite Group, 6611 Folsom-Auburn Road, Suite C, Folsom, CA 95630, USA

S. WONG Boeing Commercial Airplane Group, Douglas Products Division, Mail Stop D001-0018, Long Beach, CA 90846, USA

xiv Handbook of composites MAURICE A. WRIGHT Center for Advanced Friction Studies, Southern Illinois University at Carbondale, Carbondale, IL 62901-4343, USA

PHILIP R. YOUNG Emory & Henry College, Department of Chemistry, Emory, VA 24327, USA

ABOUT THE EDITOR

S.T. Peters was previously a fellow engineer with Westinghouse Electric Corporation, Marine Division prior to devoting full time to composite and materials and processing consulting for his own company, Process Research, in Mountain View, CA. He has written many articles on composites and filament winding, a book on filament winding, edited one previous book and holds several patents on winding techniques and composite joints.

He is a private consultant with worldwide clients and has presented tutorials on composites to many audiences, including the US Navy and NASA, several technical societies and two universities. He is a licensed professional engineer in the state of California, a member of ASM, and the composites division of SME and has been elected a fellow of SAMPE.

ACKNOWLEDGEMENTS

As with any large undertaking there is a supporting group of people without whose help the objective would not be met. I wish to acknowledge my wife, Lynn, for her help in deciphering and rewriting some of the articles and for enduing my sometimes uncivil

approach to resolving problems. Thanks also go to Mr Frank Heil and Dr Alvin Nakagawa of Westinghouse Electric, Marine Division (now Norton Grumman) for their editorial and review help. I also wish to thank Dr Linda Clements for her advice and support.

PREFACE

Today, fiber reinforced composites are in use in a variety of structures, ranging from spacecraft and aircraft to buildings and bridges. This wide use of composites has been facilitated by the introduction of new materials, improvements in manufacturing processes and developments of new analytical and testing methods. Unfortunately, information on these topics is scattered in journal articles, in conference and symposium proceedings, in workshop notes, and in government and company reports. This proliferation of the source material, coupled with the fact that some of the relevant publications are hard to find or are restricted, makes it difficult to identify and obtain the up-to-date knowledge needed to utilize composites to their full advantage. This book intends to overcome these difficulties by presenting, in a single volume, many of the recent advances in the field of composite materials. The main focus of this book is on polymeric matrix, metal matrix, and ceramic matrix composites. The book treats a wide range of subjects. The topics, presented in 49 chapters and two appendices include: 0

overview of composite material systems and products;

0 0

0 0 0 0 0 0

0 0

properties of different component (fiber, matrix, filler) materials; manufacturing techniques; analysis and design; testing; mechanically fastened and bonded joints; repair; damage tolerance; environmental effects; health, safety, reuse, and disposal; applications in: aircraft and spacecraft; land transportation; marine environments; biotechnology; construction and infrastructure; sporting goods.

Each chapter, written by a recognized expert, is self-contained, and contains many of the 'state-of-the-art' techniques required for practical applications of composites. Thus, this book should serve as a useful source of information for practicing engineers and specialists, as well as for workers new to this field. George S. Springer

CONTENTS

Contributors

ix

Preface

xv

About the editor

xvi

Foreword

xvii

Acknowledgements

xviii

Introduction, composite basics and road map S.T. Peters 1 Overview of composite materials

1

21

Theodore J. Reinhart PART ONE: BASIC MATERIALS Polymeric matrix systems 2 Polyester and vinyl ester resin Frank A. Cassis and Robert C. Talbot

34

3 Epoxyresins L.S. Penn and H . Wang

48

4 High temperature resins

75

Hugh H . Gibbs 5 Speciality matrix resins David A . Shimp 6 Thermoplastic resins Lars A. Berglund

99 115

Reinforcements and composites 7 Fiberglass reinforcement Dennis J. Vaughan

131

vi Handbook of composites 8 Boron, high silica, quartz and ceramic fibers Anthony Marzullo

156

9 Carbon fibers Khalid Lafdi and Maurice A. Wright

169

10 Organic fibers Linda L. Clements

202

11 Particulate fillers Harry S. Katz

242

12 Sandwich construction Andrew C. Marshall

254

13 Metal matrix composites V l . Kostikov and V S . Kilin

29 1

14 Ceramic composites M.E Amateau

307

15 Carbon-carbon composites John D. Buckley

333

PART TWO: PROCESSING METHODS General composites and reinforced plastics 16 Hand lay-up and bag molding D.R. Sidwell

352

17 Matched metal compression molding of polymer composites Enamul Haque and Burr (Bud) L. Leach

378

18 Textile preforming Frank K. KO and George W. Du

397

19 Table rolling of composite tubes John T. Kanne and Jerome S. Berg

425

20 Resin transfer molding Lihwa Fong and S.G. Advani

433

21 Filament winding Yu.M. Tarnopol’skii, S.T. Peters, A.I. Beil’

456

22 Fiber placement Don 0. Evans

476

23 Pultrusion Brian A. Wilson

488

24 Processing thermoplastic composites James L. Throne

525

Contents vii Advanced composites 25 Tooling for composites Jerry L. Cadden and Paul F. Sadesky

556

26 Consolidation techniques and cure control Zhong Cui and Timothy Gutowski

576

27 Composite machining Kent E. Kohkonen and Nitin Potdar

596

28 Mechanical fastening and adhesive bonding D. W. Oplinger

610

29 Surface preparations for ensuring that the glue will stick in bonded composite structures L.J. Hart-Smith, D. Brown and S. Wong

667

PART THREE: DESIGN AND ANALYSIS 30 Laminate design Jocelyn M . Seng

686

31 Design of structure with composites F.J. Schwan

709

32 Analysis methods V.V. Vasiliev

736

33 Design allowables substantiation Christy Kirchner Lapp

758

34 Mechanical tests Yu.M. Tarnopol'skii and V.L. Kulakov

778

PART FOUR. ENVIRONMENTAL EFFECTS 35 Durability and damage tolerance of fibrous composite systems Ken Reifsnider

794

36 Environmental effects on composites A n n F. Whitaker, Miria M . Finckenor, Harry W. Dursch, R.C. Tennyson and Philip R. Young

810

37 Safety and health issues

822

Jennifer A. Heth 38 Nondestructive evaluation methods for composites Thomas S. Jones

838

39 Repair aspects of composite and adhesively bonded aircraft structures Anton L. Seidl

857

40 Reuse and disposal Harry E , Pebly

883

viii Handbook of composites

PART FIVE APPLICATIONS 41 Land transportation applications Douglas L. Denton

905

42 Marine applications Wayne C. Tucker and Thomas Juska

916

43 Commercial and industrial applications of composites Stewart N. Loud

931

44 Composite biomaterials Shalaby W. Shalaby and Robert A. Latour

957

45 Scientificapplications of composites Vicki I? McConnell

967

46 Construction Ever J. Barber0

982

47 Aerospace equipment and instrument structure G a y C. Krumweide and Eddy A. Derby

1004

48 Aircraft applications Richard N. Hadcock

1022

49 Composites in the sporting goods industry Brian E. Spencer

1044

APPENDICES Appendix A Typical properties for advanced composites Kenneth R. Berg

1053

Appendix B Specifications and standards for polymer composites Frank T. Traceski

1059

Index

1069

INTRODUCTION, COMPOSITE BASICS AND ROAD MAP* S.T. Peters

This is an introduction to composites and will encourage the reader to obtain more information. Only the basic concepts will be covered here; reference will be made to the chapters in the book that expand or follow up and elaborate on these basics. The reader will see that the subjects of this book cover the spectrum of composites and range from the basic and simple to the complex. Thus, there are complicated equations because they are the tools that are used every day to describe real structures; and there will also be the more general, less complicated approaches that are limited in analysis power. These chapters have been developed by the most knowledgeable composite professionals in the world; a blend of academicians and the engineers who fabricate real composite structures. Modern structural composites, frequently referred to as ’Advanced Composites’, are a blend of two or more components, one of which is made up of stiff, long fibers, and the other, a binder or ’matrix’ which holds the fibers in place. The fibers are strong and stiff relative to the matrix and are generally orthotropic (having different properties in two different directions). The fiber, for advanced structural composites, is long, with length to diameter ratios of over 100. The fiber’s strength and stiffness are usually much greater, perhaps several times more, than the matrix material. The matrix material can by polymeric (e.g. polyester resins, epoxies), Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

metallic, ceramic or carbon. When the fiber and the matrix are joined to form a composite they retain their individual identities and both directly influence the composite’s final properties. The resulting composite will generally be composed of layers (laminae) of the fibers and matrix stacked to achieve the desired properties in one or more directions. The high strength or stiffness to weight ratios of advanced composites are well known, but there are other advantages also (Table 1.1). These advantages translate not only into aircraft, but into everyday activities, such as longer drives with a graphite-shafted golf club (because more of the mass is concentrated at the clubhead) or less fatigue and pain because a graphite composite tennis racquet has mherent damping. Generally, the advantages accrue for any fiber/composite combination and disadvantages are more obvious with some. These advantages have now resulted in many more reasons for composite use as shown in Table 1.2. Proper design and material selection can circumvent many of the disadvantages. 1.1 MATERIAL SYSTEMS

An advanced composite laminate can be tailored so that the directional dependence of strength and stiffness matches that of the loading environment. To do that, layers of unidirectional material called laminae are ori* This chapter has been adapted from S.T. Peters, in Handbook of Plastics Elastomers and Composites, 3rd edn, (ed. C.A. Harper). McGraw-Hill, New York, 1996, and is used with permission of the McGraw-Hill companies.

2 Introduction, composite basics and road map Table 1.1 Advantages/disadvantages of advanced composites Advantages

Disadvantages

Weight reduction High strength or stiffness to weight ratio

Cost of raw materials and fabrication

Tailorable properties Can tailor strength or stiffness to be in the load direction

Transverse properties may be weak

Redundant load paths (fiber to fiber)

Matrix is weak, low toughness

Longer life (no corrosion)

Reuse and disposal may be difficult

Lower manufacturing costs because of less part count

Difficult to attach

Inherent damping

Analysis is difficult

Increased (or decreased) thermal or electrical conductivity

Matrix subject to environmentaldegradation

___-

Carbon/graphite fibers (Chapter 9) have demonstrated the widest variety of strengths and modulii and have the greatest number of suppliers. The fibers begin as an organic fiber, rayon, polyacrylonitrile or pitch which is called the precursor. The precursor is then stretched, oxidized, carbonized and graphitized. There are many ways to produce these fibers, but the relative amount of exposure at temperatures from 2500-3000°C results in greater or less graphitization of the fiber. Higher degrees of graphitization usually result in a stiffer fiber (higher modulus) with 0 fiberglass; greater electrical and thermal conductivities 0 graphite; and usually higher cost. 0 aramid; The organic fiber Kevlar 49, (Chapter 10) 0 polyethylene; also called aramid, essentially revolutionized 0 boron; pressure vessel technology because of its great 0 silicon carbide; tensile strength and consistency coupled with 0 silicon nitride, silica, alumina, alumina silica. low density, resulting in much more weight The advantages of fiberglass (Chapter 7) are its effective designs for rocket motors. Aramid high tensile strength and strain to failure, but composites are still widely used for pressure heat and fire resistance, chemical resistance, vessels but have been largely supplanted by moisture resistance and thermal and electrical the very high strength graphite fibers. Aramid properties are also cited as reasons for its use. composites have relatively poor shear and It is by far the most widely used fiber, primar- compression properties; careful design is ily because of its low cost; but its mechanical requires for their use in structural applications properties are not comparable with other that involve bending or compression. structural fibers.

ented to satisfy the loading requirements. These laminae contain fibers and a matrix. Because of the use of directional laminae, the tensile, flexural and torsional shear properties of a structure can be disassociated from one another to some extent and a golf shaft, for example, can be changed in torsional stiffness without changing the flexural or tensile stiffness. Fibers can be of the same material within a lamina or several fibers mixed (hybrid). The common commercially available fibers are as follows:

Material systems 3 Table 1.2 The reasons for using composites

Reason for use

Material selected

___

Appl ica t ion/driver

Lighter, stiffer stronger

Boron, all carbodgraphites, some aramid

Military aircraft, better performance Commercial aircraft, operating costs

Controlled or zero thermal expansion

Very high modulus carbon/graphite

Spacecraft with high positional accuracy requirements for optical sensors

Environmental resistance

Fiberglass, vinyl esters, bisphenol A fumarates, chlorendic resins

Tanks and piping, corrosion resistance to industrial chemicals, crude oil, gasoline at elevated temperatures

Lower inertia, faster startups, less deflection

High strength carbon/graphite, epoxy

Industrial rolls, for paper, films

Lightweight, damage tolerance

High strength carbon/graphite, CNG tanks for ’green’cars, trucks fiberglass, (hybrids), epoxy and busses to reduce environmental pollution

More reproducible complex surfaces

High strength or high modulus carbon graphite/ epoxy

High-speed aircraft. Metal skins cannot be formed accurately

Less pain and fatigue

Carbon/graphite/epoxy

Tennis, squash and racquetball racquets. Metallic racquets are no longer available

Reduces logging in ‘old growth’ forests

Aramid, carbon/graphite

Laminated ‘new’ growth wooden support beams with high modulus fibers incorporated

Reduces need for intermediate support and resists constant 100% humidity atmosphere

High strength carbon/graphite-epoxy

Cooling tower driveshafts

Tailorability of bending and twisting response

Carbon/graphite-epoxy

Golf shafts, fishing rods

Transparency to radiation

Carbon/ graphite-epoxy

X-ray tables

Crashworthiness

Carbon/ graphite-epoxy

Racing cars

Higher natural frequency, lighter Carbon/ graphite-epoxy

Automotive and industrial driveshafts

Water resistance

Fiberglass (woven fabric), polyester or isopolyester

Commercial boats

Carbon/graphite, fiberglass

Freeway support structure repair after earthquake

Ease of field application

- epoxy, tape and fabric

The polyethylene fibers have the same property drawbacks as aramids, but also suffer from low melting temperature which limits

their use to composites that cure or operate below 149°C (300°F) and a susceptibility to degradation by ultraviolet light exposure.

4 lntvodmction, composite basics and road map

Both of these types of fibers have wide usage in personal protective armor. In spite of the drawbacks, production of both of these fibers is enjoying strong worldwide growth. Boron fibers (Chapter 8), the first advanced composite fibers to be used on production aircraft, are produced as individual monofilaments upon a tungsten or carbon substrate by pyrolytic reduction of boron trichloride (BC1,) in a sealed glass chamber. The relatively large cross section fiber is used today primarily in metal matrix composites which are processed at temperatures which would attack carbon/graphite fibers.

1.2 MATRIX SYSTEMS

If parallel and continuous fibers are combined with a suitable matrix and cured properly, unidirectional composite properties such as those shown on Table 1.3 are the result. The functions and requirements of the matrix are to: 0

0 0

0

0

keep the fibers in place in the structure; help to distribute or transfer loads; protect the filaments, both in the structure and before and during fabrication; control the electrical and chemical properties of the composite; carry interlaminar shear.

Table 1.3 Properties of typical unidirectional graphite/epoxy composites (Fiber volume fraction, V , = 0.62)

Elastic constants

High strength

High modulus

GPa (psi x I O 6 )

GPa (psi x IO6)

145 (21) 9.6 (1.4) 5.8 (0.85) 0.30

220 (32) 6.9 (1.0) 4.8 (0.7) 0.30

~~

Longitudinal modulus, E, Transverse modulus, E , Shear modulus, G , Poisson’s ratio (dimensionless)u ~ , ~~~

.~

~~

~

~~

~~

~~

Strength properties ~

Longitudinal tension, Ft”, Transverse tension, FtUT Longitudinal compression, FCUL Transverse compression, FCUT Inplane shear, PLT Interlaminar shear, F’,””

~~~

~~~~

~

MPa ( Z 0 3 psi)

MPa (lo3psi)

2139 (310) 54 (7.8) 1724 (250) 76 (11) 87 (12.6) 128 (18.5)

760 (110) 28 (4) 690 (100) 170 (25) 70 (10) 70 (10)

~

Ultimate strains -

%

-~

Longitudinal tension, Transverse tension, Longitudinal compression, ECUL Transverse compression, EC1lT Inplane shear

1.4 0.67 0.9 3.6 2.0

0.3 0.4 0.3 2.8

1600 (0.056)

1700 (0.058)

-0.079 (-0.044)

-0.54 (-0.3)

21.6 (12)

58 (32)

-

Physical properties Density, kg/m3 (Ib/in3) Longitudinal CTE, ye/K (pe/OF) Transverse CTE ~ E / (pe/OF) K

From References 1, 2 and 3; CTE = coefficient of thermal expansion

Matrix systems 5 The needs, or desired properties of the matrix, that depend on the purpose of the structure are: 0 0 0 0 0

0 0

0

0

0

0

minimize moisture absorption; have low shrinkage; Must wet and bond to fiber; low coefficient of thermal expansion; must flow to penetrate the fiber bundles completely and eliminate voids during the compacting/curing process; have reasonable strength, modulus and elongation (elongation should be greater than fiber); must be elastic to transfer load to fibers; have strength at elevated temperature (depending on application); have low temperature capability (depending on application); have excellent chemical resistance (depending on application); be easily processable into the final composite shape; have dimensional stability (maintain its shape).

There are many matrix choices available; each type has impact o n the processing technique, physical and mechanical properties and environmental resistance of the finished

composite. The common thermoset matrices for composites include the following: 0

0 0 0

polyester and vinylesters (Chapter 2); epoxy (Chapter 3); bismaleimide (BMI) (Chapter 4); polyimide (Chapter 4); cyanate ester and phenolic triazine (Chapter 5).

Each of the resin systems has some drawbacks, which must be accounted for in design and manufacturing plans. Polyester matrices have been in use for the longest period, and are used in the widest range and greatest number of structures. The usable polymers may contain up to 50% by weight of unsaturated monomers and solvents such as styrene. Polyesters cure via a catalyst (usually a peroxide) resulting in an exothermic reaction, which can be initiated at room temperature. The most widely used matrices for advanced composites have been the epoxy resins. These resins cost more than polyesters and do not have the high temperature capability of the bismaleimides or polyimides, but because of the advantages shown in Table 1.4 they are widely used.

Table 1.4 Selection criteria for epoxy resin systems

Advantages

Disadvantages

Adhesion to fibers and to resin

Resins and curatives somewhat toxic in uncured form

No by-products formed during cure

Absorb moisture Heat distortion point lowered by moisture absorption

Low shrinkage during cure High or low strength and flexibility Solvent and chemical resistance Resistance to creep and fatigue Solid or liquid resins in uncured state Wide range of curative options Adjustable curing rate Good electrical properties

Change in dimensions and physical properties due to moisture absorption Limited to about 200°C upper temperature use (dry) Difficult to combine toughness and high temperature resistance High thermal coefficient of expansion High degree of smoke liberation in a fire May be sensitive to ultraviolet light degradation Slow curing

6 Introduction, composite basics and road map There are two resin systems in common use for higher temperatures, bismaleimides and polyimides. New designs for aircraft demand a 177°C (350°F) operating temperature not met by the other common structural resin systems. The primary bismaleimide in use is based on the reaction product from methylene dianiline (MDA) and maleic anhydride: bis (4-maleimidophenyl) methane (MDA BMI). Two newer resin systems have been developed and have found applications in widely diverse areas. The cyanate ester resins, marketed by Ciba-Geigy, have shown superior dielectric properties and much lower moisture absorption than any other structural resin for composites. The dielectric properties have enabled their use as adhesives in multilayer microwave printed circuit boards, and the low moisture absorbance have caused them to be the resin of universal choice for structurallystable spacecraft components. The phenolic triazine (PT) resins also have superior elevated temperature properties, along with excellent properties at cryogenic temperatures. Their resistance to proton

ROVING

radiation under cryogenic conditions was a prime cause for their choice for use in the superconducting supercollider, subsequently canceled by the US Congress. Polyimides are the highest temperature polymer in general advanced composite use with a long term upper temperature limit of 232°C (450°F) or 316°C (600°F). Two general types are: condensation polyimides, that release water during the curing reaction, and addition type polyimides with somewhat easier process requirements. 1.3 FIBER MATRIX SYSTEMS

The end user sees a composite structure. Someone else, probably a prepregger, combined the fiber and the resin system and someone else caused the cure and compaction to result in a laminated structure. A schematic of the steps to arrive at a finished composite from the initial fiber is shown in Fig. 1.1. In many cases, the end user of the structure has fabricated the composite from prepreg, which is a low-temperature-stable combination

WEAUE?

N?

COLLIMRTE

UNI TRPE

A

Fig. 1.1 Manufacturing steps in composite structure.

Fiber matrix systems 7 of the resin, its curing agents and the fiber. The three types of continuous fibers, roving, tape and woven fabric available as prepregs give the end user many options in terms of design and manufacture of a composite structure. Although the use of dry fibers and impregnation at the work (i.e. filament winding, pultrusion or hand lay-up) is very advantageous in terms of costs; there are many advantages to the use of prepregs as shown in Table 1.5, particularly for the manufacture of modem composites.

The prepreg process for thermoset matrices can be accomplished by feeding the fiber continuous tape, woven fabric or roving through a resin-rich solvent solution and then removing the solvent by hot tower drying. The excess resin is removed via a doctor blade or metering rolls and then the product is staged to the cold-stable prepreg form, (B stage) (Fig. 1.2). The newer hot melt procedure for prepregs is gradually replacing the solvent method because of environmental concerns. A film of resin that has been cast hot onto release paper

Table 1.5 Advantages of prepregs over wet impregnation

Prepregs reduce the handling damage to dry fibers Improve laminate properties by better dispersion of short fibers Prepregs allow the use of hard-to-mix or proprietary resin systems Allow more consistency because there is a chance for inspection before use Heat curing provides more time for the proper laydown of fibers and for the resin to move and degas before cure Increased curing pressure reduces voids and improves fiber wetting Most prepregs have been optimized as individual systems to improve processing

Release Poly

Unwind

Prepreg Wind

Pump and Reservoir

Unwind

Fig. 1.2 Schematic of typical solvent prepregging process. (Adapted from Reference 2.)

8 Introduction, composite basics and road map

0

0

Pauer Paper

T

Doctor Plate 1 Impregnation Zone Creel

Paper

Plate 2

Take-up Prepreg Windup

Chill Plate

Fig. 1.3 Schematic of typical film impregnating process. (Adapted from Reference 2.)

is fed, along with the reinforcement, through a series of heaters and rollers to force the resin into the reinforcement. Two layers of resin are commonly used so that a resin film is on both sides of the reinforcement; one of the release papers is removed and the prepreg is then trimmed, rolled and frozen (Fig. I.3)2.The solvent technique has been largely replaced for advanced fibers because of environmental pollution concerns and a need to exert better control over the amount of resin on the fiber. 1.3.1 UNIDIRECTIONAL PLY PROPERTIES

The manufacturer of the prepreg reports an areal weight for the prepreg and a resin percentage, by weight. Each of the different fibers has a different density, resulting in a composite of different density at the same fiber volume percentage. Since fiber volume is used to relate the properties of the manufactured composites, the following equations can be used to convert between weight fraction and fiber volume.

where:

Wf = weight fraction of fiber wf = weight of fiber wc= weight of composite pf = density of fiber p, = density of composite uf = volume of fiber u, = volume of composite Vf = volume fraction of fiber V, = volume fraction of matrix p, = density of matrix. A percentage fiber that is easily achievable and repeatable in a composite and convenient for reporting mechanical and physical properties for several fibers is 60%. The properties of unidirectional fiber laminates are shown in Table 1.3 for carbon/graphite/epoxy. Values for the other fibers can be seen in their respective chapters. These values are for individual lamina or for a unidirectional composite, and they represent the theoretical maximum (for that fiber volume) for longitudinal in plane properties. Transverse, shear and compression properties will show maxima at different fiber volumes and for different fibers, depending on how the matrix and fiber interact. These values can be used to calculate the properties of a laminate which has fibers oriented in several directions. To do that, the methods of description for ply orientation must be introduced.

Quasi-isotropic laminate 9 1.4 PLY ORIENTATIONS, SYMMETRY AND

BALANCE

lined to indicate that half of it lies on either side of the plane of symmetry (Fig. 1.4(f)).

1.4.1 PLY ORIENTATIONS

One of the advantages of using a modern composite is the potential to orient the fibers to respond the load requirements. This means that the composite designer must show the material, the fiber orientations in each ply, and how the plies are arranged (ply stackup). A 'shorthand' code for ply fiber orientations has been adapted for use in layouts and studies. Each ply (lamina)is shown by a number representing the direction of the fibers in degrees, with respect to a reference ( x ) axis. 0" fibers of both tape and fabric are normally aligned with the largest axial load (axis) (Fig. 1.4(a)). Individual adjacent plies are separated by a slash in the code if their angles are different (Fig. 1.4@)). The plies are listed in sequence, from one laminate face to the other, starting with the ply first on the tool and indicated by the code arrow with brackets indicating the beginning and end of the code. Adjacent plies of the same angle of orientation are shown by a numerical subscript (Fig. 1.4(c)). When tape plies are oriented at angles equal in magnitude but opposite in sign, (+) and (-) are used. Each (+) or (-) sign represents one ply. A numerical subscript is used only when there are repeating angles of the same sign. Positive and negative angles should be consistent with the coordinate system chosen. An orientation shown as positive in one right handed coordinate system may be negative in another. If the y and z axis directions are reversed, the f 45 plies are reversed (Fig. 1.4(d)). Symmetric laminates with an even number of plies are listed in sequence, stating at one face and stopping at the midpoint. A subscript 'S' following the bracket indicates only one half of the code is shown (Fig. 1.4(e)). Symmetric laminates with an odd number of plies are coded as a symmetric laminate except that the center ply, listed last, is over-

1.4.2 SYMMETRY

The geometric midplane is the reference surface for determining if a laminate is symmetrical. In general, to reduce out-ofplane strains, coupled bending and stretching of the laminate and complexity of analysis, symmetric laminates should be used. However, some composite structures (e.g. filament wound pressure vessels) can achieve geometric symmetry so that symmetry through a single laminate wall is not necessary, if it constrains manufacture. To construct a midplane symmetric laminate, for each layer above the midplane there must exist an identical layer (same thickness, material properties, and angular orientation) below the midplane (Fig. 1.4(e)). 1.4.3 BALANCE

All laminates should be balanced to achieve inplane orthotropic behavior. To achieve balance, for every layer centered at some positive angle +e there must exist an identical layer oriented at -8 with the same thickness and material properties. If the laminate contains only 0" and/or 90" layers it satisfies the requirements for balance. Laminates may be midplane s p metic but not balanced and vice versa. Figure 1.4(e) is symmetric and balanced whereas Fig. 1.4(g)is balanced but unsymmetric . 1.5 QUASI-ISOTROPICLAMINATE

The goal of composite design is to achieve the lightest, most efficient structure by aligning most of the fibers in the direction of the load. Many times there is a need, however, to produce a composite which has some isotropic properties, similar to metal, because of multiple or undefined load paths. A 'quasi-isotropic' laminate lay-up accomplishes this for the x and y planes only; the z or through-the-laminate-

-

10 Introduction, composite basics and road map 90"

Reference Axis

lz;

90"

Tool side

,/

.-,

Tape Laminate

0"

1

I I P

I

45'

90' -45'

-450

w

\

\

0"

+450

[0/903/0]

P

I

90" +45" -45" -45" +45" 90" V

0"

[0/9O]s Typical Callout

T

[0/90/*45]s

Typical Callout

Line of Symmetry

I

Tape and Fabric Laminate [ 0/f45/To1 s. Typical Callout

Line of Symmetry

Fig. 1.4 Ply orientations, symmetry and balance. (Continued on next page)

0

0"

L

Methods of analysis 11 Tape Laminate

p, +45" -45"

[0/90/f45/i452/9 0/ 01 Typical Callout

-45"

+45"

Fabric Laminate

I

0".90" I

j

0",90"

I

[(0,90)/(~45)/(0,90)] Typical Callout

h)

Fig. 1.4 Ply orientations, symmetry and balance. (Continued)

thickness plane is quite different and lower. 1. arrive at quick values to determine if a comMost laminates produced for aircraft applicaposite is feasible; tions have been, with few exceptions, 2. arrive at values for insertion into computer 'quasi-isotropic'. As designers become more programs for laminate analysis or finite eleconfident and have access to a greater database ment analysis; with fiber-based structures, more applications 3. check on the results of computer analysis. will evolve. For a quasi-isotropic laminate, the The rule of mixtures holds for composites. The following are requirements: micromechanics formula to arrive at the 0 It must have three layers or more. Young's modulus for a given composite is: 0 Individual layers must have identical stiffEc = V,E, + Vm Em ness matrices and thicknesses. 0 The layers must be oriented at equal angles. and v,+ vm= 1 For example, if the total number of layers is = V ,E , + Em (1- V,) (1.3) M , the angle between two adjacent layers should be 360"ln. If a laminate is con- where structed from identical sets of three or more Ec = composite or ply Young's modulus in layers each, the condition on orientation tension for fibers oriented in direction of must be satisfied by the layers in each set, applied load for example: ( O o / + 60"), or ( O o / + 45"/90)s. V = volume fraction of fiber ( f ) or matrix (m) E = Young's modulus of fiber ( f ) or matrix 1.6 METHODS OF ANALYSIS (m). There are a number of methods in common But, since the fiber has much higher use for the analysis of composite laminates. Young's modulus than the matrix, the second The use of micromechanics, i.e. the application part of the equation can be ignored. of the properties of the constituents to arrive at E, >> Em the properties of the composite ply can be used to: Ec = E,V, (1.4)

12 Introduction, composite basics and road map appropriate for a particular application. Figure 1.5 shows the progression of physical properties for Young’s modulus in tension, E, (fiber), E, (lamina) and Ex,, (laminate), longitudinal tensile strength, and coefficient of thermal expansion a where the subscripts L and X stand for in-plane in the principal fiber direction and t and Y stand for the transverse direction for a theoretical high strength (from Ec = (3/8) E,V, (1.5) Table 1.3) carbon/graphite fiber composite The quasi-isotropic modulus, E, of a composite from the fiber to the laminate. The values laminate is (3/8)E,+(5/8)EZ where E,, is the decrease or are ’translated’in a logical fashion modulus of the lamina in the fiber direction and and reflect the law of mixtures. The analysis is E, is the transverse modulus of the lamina3. relatively simple for modulus dominated The transverse modulus for polymeric-based properties but strength-dominated values composites is a small fraction of the longitudinal must be treated in light of one of several failmodulus (see E, in Table 1.3)and can be ignored, ure theories and changes in the thermal for preliminary estimates, resulting in a slightly coefficient of expansion are not predictable lower-than-theoretical value for Ec for a quasi- from laws of mixtures. Other factors which isotropic laminate. This approximate value for enter into the translation efficiency are: comthe quasi-isotropic modulus represents the patibility of the resin system with the fiber and lower limit of composite modulus. It is useful in the fiber finish, strain-to-failure of the resin comparing of composite properties to those of system and the damage the fiber undergoes metals and in establishing if a composite is during impregnation, laydown and cure.

This is the basic rule of mixture and represents the highest Young’s modulus composite, where all fibers are aligned in the direction of load. The minimum Young’s modulus for a reasonable design (other than a preponderance of fibers being orientated transverse to the load direction) is the quasi-isotropic composite and can be approximated by:

.6 GPa, FT‘“ =54 MPa

E x = 76 GPa a x = 4.98peK

r

E y = 76 GPa a y = 4.98~ E K

a2

>ay

Fig. 1.5 The anatomy of a composite laminate.

Composite fabrication techniques Table 1.6 High-strength carbon/graphite laminate

properties Laminate

(0/90,/0) (90/0,/90) (02/902/OJ (0,/~45,/0,) (0/+45/90)>

Aluminum

Longitudinal modulus E,, (GPa) 76.5 76.5 98.5 81.3 55.0 41.34

Bending modulus, E , (GPa) 126.8 26.3 137.8 127.5 89.6 41.34

Shear modulus, G,, (GPa) 5.24 5.24 5.24 21.0 21.0 27.56

Table 1.6 shows mechanical values for several composite laminates with the fiber of Table 1.3 and a typical resin system. The first and second entries are for simple 0/90 laminates and show the effect of changing the position of the plies. The effect of increasing the number of 0 plies is shown next and the final two laminates demonstrate the effect of +45 plies on mechanical properties, particularly the shear modulus. The last entry is a quasi-isotropic laminate. These laminates are then compared to a typical aluminum alloy. When employing the data extracted from tables, some caution should be observed by the reader. The values seen in many tables of data may not always be consistent for the same materials or the same group of materials from several sources for the following reasons: 1. Manufacturers have been refining their production processes so that newer fibers may have greater strength or stiffness. These new data may not be reflected in the compiled data. 2. The manufacturer may not be able to change the value quoted for the fiber because of government or commercial restrictions imposed by the specification process of his customers. 3. There are many different high-strength fibers commercially available. Each manufacturer has optimized their process to maximize their mechanical properties and each process may differ from that of the

13

competitor, so vendor values in a generic class may differ widely. 4. Most tables of values are presented as 'typical values'. Those values and the values that are part of the menu of many computer analysis programs should be used with care. Each user must find their own set of values for design, develop useful design allowables, and apply appropriate 'knock down' factors, based on the operating environments expected in service. (Chapter 33 and Appendix A give guidelines.) 1.7 COMPOSITE FABRICATION TECHNIQUES

The goals of the composite manufacturing process are to: 0

achieve a consistent product by controlling fiber thickness; - fiber volume; - fiber directions; minimize voids; reduce internal residual stresses; process in the least costly manner. -

0 0 0

The procedures to reach these goals involve iterative processes to select the three key components: 0 0 0

composite material and its configuration; tooling; process.

Once material selection has been completed, the first step leading to the acceptable composite structure is the selection of tooling, which is intimately tied to process and material. For all curing techniques the tool must be: 0

0

0

0

strong and stiff enough to resist the pressure exerted during cure; dimensionally stable through repeated heating and cooling cycles; light enough to respond reasonably quickly to the changes in cure cycle temperature and to be moved in the shop; leakproof so that the vacuum and pressure cycles are consistent.


The tool face is commonly the surface imparted to the outer surface of the composite and must be smooth, particularly for aerodynamic surfaces. The other surface frequently may be of lower finish quality and is imparted by the disposable or reusable vacuum bag. This surface can be improved by the use of a supplemental metal tool known as a caul plate. (Press curing, resin transfer molding, injection molding and pultrusion require a fully closed or two sided mold). Figure I. 6 shows the basic components of the tooling for vacuum bag or autoclave processed components and Table 1.7 shows the function of each part of the system. Tooling options have been augmented by 3

2

12

13

4

14

the introduction of elastomeric tooling wherein the thermal expansion of an elastomer provides some or all of the pressure curing cure, or a rubber blanket is used as a reusable vacuum bag. The volumetric expansion of an elastomer can be used to fill a cavity between the uncured composite and an outer mold. The use of elastomeric tooling can provide the means for fabricating complex box-like structures such as integrally stiffened skins with a co-cured substructure in a single curing operation. Tooling (Chapter 25) and the configuration of the reinforcement have a great influence on the curing process selected and vice-versa. The 5

6

9

8

7

10

9

11

Fig. 1.6 Typical vacuum bag lay-up components. Table 1.7 Functions of vacuum bag components

Component *

Functions

-

1 2 3 4 5 6 7 8 9 10 11 12

13 14

Bag sealant Vacuum fitting and hardware Bagging film Open weave breather mat Polyester tape (wide) Polyester tape (narrow) Caul sheet Perforated release film Non-perforated release film Peel ply Laminate 1581-styleglass breather manifold 1581 style glass bleeder ply Stacked silicone edge dam

* numbers refer to Fig. 1.6

Temporarily bonds vacuum bag to tool Exhausts air, provides convenient connection to vacuum pump Encloses part, allows for vacuum and pressure Allows air or vacuum transfer to all of part Holds other components of bag in place Holds components in place Imparts desired contour and surface finish to composite Allows flow of resin or air without adhesion Prevents adhesion of laminate resin to tool surface Imparts a bondable surface to cured laminate Allows transfer of air or vacuum Soaks up excess resin Forces excess resin to flow vertically, increasing fluid pressure

Composite fabrication techniques 15 probable reinforcement configuration that facilitates the completion of the finished composite is shown on Table 1.8. The choice between unidirectional tape and woven fabric has frequently been made on the basis of the greater strength and modulus attainable with the tape particularly in applications which compression strength is important. There are other factors that should be included in the trade, as shown in Table 1.9. 1.7.1 LAY-UP TECHNIQUE

Lay-up techniques along with composite cure control have received the greatest attention for processing. In efforts to reduce labor costs of composite fabrication, to which lay-up (Chapter 16) has traditionally been the largest contributor, mechanically assisted, controlled tape laying and automated integrated manu-

Table 1.8 Common reinforcement configuration for the manufacturing process

Reinforcement Prepreg Prepreg Prepreg Other, configuration tape or (dry) or (dry) woven tow woven preforms, or non- chopped woven fibers fabric Handlay-up Automatic tape laydown

X X

x, (XI

x

Filament winding

x, (X)

xm

xm

Resin transfer molding

(XI

(X)

X

Pultrusion

(X)

Fiber placement

X

X

X

Table 1.9 Fabric compared with tape reinforcement

Tape advantages

_

_

_

_

_

~

-

~

Tape disadvantages

Best modulus and strength efficiency

Poor drape on complex shapes

High fiber volume achievable

Cured composite more difficult to machine

Low scrap rate

Lower impact resistance

No discontinuities

Multiple plies required for balance and symmetry

Automated lay-up possible

Higher labor costs for hand lay-up

Available in thin plies Lowest cost prepreg form Less tendency to trap volatiles

Fabric advantages

Fabric disadvantages

~

_ _ ~ _ _ _ - - ~ _ _

Better drape for complex shapes

Fiber discontinuities (splices)

Single ply is balanced and may be essentially symmetric

Less strength and modulus efficient

Can be laid up without resin

Lower fiber volume than tape

Plys stay in line better during cure

More costly than tape

Cured parts easier to machine

Greater scrap rates

Better impact resistance

Warp and fill properties differ

Many forms available

Fabric distortion can cause part warping

16 Introduction, composite basics and road m a p facturing systems have been developed. Table Generally, the percent matrix weight is higher 1.10 shows some of the considerations for before cure initiation; the matrix flows out of choosing a lay-up technique. the laminate and takes the excess resin with In addition to any cost savings by the use of the potential voids. An arbitrary 1%void limit an automated technique for long production has been adopted for most autoclaved comruns, there are two key quality assurance fac- posites; filament wound and pultruded tors which validate the automated techniques. composites will have higher void volumes They are: greatly reduced chance that release depending upon the application. An autoclave is essentially a closed, prespaper or film could be retained, which would destroy shear and compressive strength if surized oven; many common epoxy laminates undetected, and reduced probability of the are cured at an upper temperature of 177°C addition or loss of an angle ply which would (350°F) and 6 MPa (100 psi). Autoclaves are cause warping due to the laminate’s lack of still the primary tool in advanced composite symmetry and balance. processing and have been built up to 16 m (55 All curing techniques use heat and pressure feet) long at 6.1 m (20 feet) diameter. Since to cause the matrix to flow and wet out all the autoclaves are expensive to build and operate, fibers before the matrix solidifies (Chapter 26). many other methods of curing, compacting Table 1.10 Considerations in composite lay-up technique

Considerat ion

Manual

Flat tape

Contoured tape

Orientation accuracy

Least accurate

Automatic

Somewhat dependent on tape accuracy and computer program

Ply count

Dependent on operator, count Mylars

Dependent on operator

Program records

Release film retention

Up to operator

Automatic

Automatic removal

Labor costs

High

86% improvement quoted

Additional improvement

Machine costs

N/A

Some costs

Approximately 1M$ or greater

Production rate

Low (1.5 Ib/h)

10 lb/h

Approximately same as flat tape

Machine ’up’ time

N/A

Not a consideration

Complex program and machine make this a consideration

Varying tape widths

Not a concern

Easily changed

Difficulty in changing

Tape lengths

Longer tapes more difficult

Longer is more economical

Longer tape is more economical

Cutting waste

Scrap on cutting

Less scrap

Least scrap due to back and forth laydown

Compaction pressure

No pressure

Less voids

Least voids

Programming

N/A

N/A

Necessary

Compositefabrication techniques 17 composites have been developed. The two newest and most attractive methods are fiber placement and resin transfer molding. 1.7.2 RESIN TRANSFER MOLDING

Previous discussions have centered on moving resin out of the laminate to reduce voids. Resin transfer involves the placement of dry fiber reinforcement into a closed mold and then injecting a catalyzed resin into the mold to encapsulate the reinforcement and form a composite (Chapter 20). The impetus for the use of this process comes from the large cost reductions that can be realized in raw materials and lay-up. The process can utilize low injection pressures i.e. 55 MPa (80 psi), therefore, the tooling can be lower cost plastic or a vacuum bag rather than metal.

a wind eye at speeds synchronized with the mandrel rotation, control winding angle of the reinforcement and the fiber lay-down rate. The reinforcement may be wrapped in adjacent bands or in repeating bands that are stepped the width of the band and that eventually cover the mandrel surface. Local reinforcement can be added to the structure using circumferential windings, local helical bands, or by the use of woven or unidirectional cloth. The wrap angle can be varied from low angle helical to high angle circumferential or 'hoop', which allows winding from about 4"-90" relative to the mandrel axis; newer machines can 'place' fiber at 0". 1.7.4 FIBER PLACEMENT

Fiber placement, initially developed by Hercules Aerospace Co., is a cross between filament winding and automatic tape laydown, 1.7.3 FILAMENT WINDING retaining many of the advantages of both. The Filament winding is a process by which con- natural outgrowth of adding multiple axes of tinuous reinforcements in the form of rovings control to filament winding machines results or tows (gathered, untwisted strands of fiber) in control of the fiber laydown so that non axiare wound over a rotating mandrel. The man- symmetric surfaces can be wound. This drel can be cylindrical, round or any other involves the addition of a modified tape layshape as long as it does not have re-entrant down head to the filament winding machine curvature. Special machines (Fig. and much more. The Cincinnati-Milacron - 1.7) traversing machine additions include in-process compaction, individual tow cut/start capabilities, a resin tack control system, differential tow payout, low tension on fiber and enhanced offf 1 iine programming (Chapter 22).

I/ 11

1.7.5 PULTRUSION

Pultrusion is an automated process for the I manufacture of constant volume/shape profiles from composite materials (Chapter 23). The composite reinforcements are continuously pulled through a heated die and shaped and cured simultaneously. If the cross-sectional shape is conducive to the process, it is p f r 4 the fastest and most economical method of Fig. 1.7 The helical filament wound ply. (Courtesy composite production. Straight and cured conof Westinghouse Electric Co., Marine Division.) figurations can be fabricated with square,

cF

I

18 Introduction, composite basics and road map round, hat-shaped, angled 'I' or 'T'-shaped cross-sections from vinylester, polyester, or epoxy matrices with E and S-glass, Kevlar and graphite reinforcements.. The curing is effected by combinations of dielectric preheating and microwave or induction (with conductive reinforcements like carbon graphite) while the shape traverses the die. 1.7.6 BRAIDING, WEAVING AND OTHER PREFORM TECHNIQUES

I

3

Fig. 1.8 The unidirectional ply.

Braiding, weaving, knitting and stitching represent methods of forming a shape, generally be the same in any transverse direction. This is referred to as preforming, with the composite the transverse isotropy assumption; it is fibers before impregnation (Chapter 18). The approximately satisfied for most unidirecshape may be the final product or some inter- tional composite plies. These properties are typically modified by mediate form such as a woven fabric. The braiding process is continuous and is transformation relative to the laminate axis amenable to round or rectangular shapes or where these may not be the same as the ply smooth curved surfaces and can transition axes. In a multidirectional laminate there can be easily from one shape to another. The other fabric preforming techniques are as many as 21 stiffness constants. Strength preweaving, knitting and the non-structural dictions are equally as complicated because of stitching of unidirectional tapes. Stitching sim- directional differences, i.e. compression is not ply uses a non-structural thread, such as nylon always equal to tension, and because the sevor Dacron, to hold dry tapes at selected fiber eral failure theories are complex. As the angles. Preforming in this manner results in a complexity of the matrix calculations increase, higher-cost raw material but saves labor costs it becomes evident that errorless mathematical for orientation of individual lamina. The manipulations are impossible without the aid stitched preform has known, stable fiber ori- of computers. Chapters 30 and 32 elaborate on entations similar to woven fabric, without the the techniques of laminate analysis and the crossovers which could reduce compressive applications of laminates to structures strength. 1.9 DESIGN OF COMPOSITES 1.8 MECHANICS OF COMPOSITE MATERIALS

The 1,2,3 axes in Fig. 1.8 are special and are called the ply axes, or material axes. The 1 axis is in the direction of the fibers, and is called the longitudinal axis or the fiber axis. The longitudinal axis is typically the highest stiffness and strength direction. Any direction perpendicular to the fibers (in the 2,3 plane) is called a transverse direction. Sometimes, to simplify analysis and test requirements, ply properties are assumed to

The design process for composites involves both laminate design and component design and must also include considerations of manufacturing process and eventual environmental exposure. These steps are all interdependent with composites and the most efficient design must involve true concurrent engineering. Figure 1.9 shows the various concerns that should be a part of the composite design process at the initiation of the design process, and continuously from there on.

Design of composites 19 1.9.1 LAMINATE DESIGN RECOMMENDATIONS

1. Take advantage of the orthotropic nature of the fiber composite ply. 0 To carry in-plane tensile or compressive loads align the fibers in the directions of these loads. 0 For in-plane shear loads, align most fibers at -c 45" to these shear loads. 0 For combined normal and shear in-plane loading provide multiple or intermediate ply angles for a combined load capability. 2. Intersperse the ply orientations. 0 If a design requires a laminate with 16 plies at *45", 16 plies at 0", and 16 plies at 90°, use the interspersed design (90,/ -c 45,/0,),s rather than (90,/ .+ 45,/10,)s. Concentrating plies at nearly the same angle (0" and 90" in the above example) provides the opportunity for large matrix cracks to form. These produce lower laminate allowables, probably because large cracks are more injurious to the fibers, and more readily form delaminations than the finer cracks occurring in interspersed laminates. 0 If a design requires all 0" plies, some 90" plies (and perhaps some off-angle plies ) should be interspersed in the laminate to provide some biaxial strength and stability and to accommodate unplanned Composite Material Environmental Considerations

Component

Fig. 1.9 Design considerations for composites.

0

loads. This improves handling characteristics, and serves to prevent large matrix cracks from forming. Locally reinforce with fabric or mat in areas of concentrated loading. (This technique is used to locally reinforce pressure vessel domes). Use fabric, particularly fiberglass or Kevlar, as a surface ply to restrict surface (handling) damage. Ensure that the laminate has sufficient fiber orientations to avoid dependence on the matrix for stability. A minimum coverage of 6 to 10% of total thickness in 0, ?45", 90" directions is recommended.

3. Select the lay-up to avoid mismatch of properties of the laminate with those of the adjoining structures - or provide a shear/separator ply. Poisson's ratio: if the transverse strain of a laminate greatly differs from that of adjoining structure, large interlaminar stresses are produced under load. Coefficient of thermal expansion: temperature change can produce large interlaminar stresses if coefficient of thermal expansion of the laminate differs greatly from that of adjoining structure. 0 The ply layer adjacent to most bonded joints should not be perpendicular to the direction of loading. Thicken the composite in the joint area, soften the composite by adding fiberglass or angle plies and select the highest strain-capability adhesive. 4. Use multiple ply angles. Typical composite laminates are constructed from multiple unidirectional or fabric layers which are positioned at angular orientations in a specified stacking sequence. From many choices, experience suggests a rather narrow range of practical construction from which the final laminate configuration is usually selected. The multiple layers are usually oriented in at least two different angles, and possibly three or four; (go, O0/&", or


attempt to standardize the raw materials and their test methods by publication of specifications (Appendix A). However, these standards have not reached the level of use to allow complete dependence upon them without supplier-user interaction and user testing. The fabricators of composites will rely on specifications for control of fiber, resin and/or the prepreg. Many prepreg resin and fiber Further suggestions can be seen in Chapter 31. vendors will certify only to their own specifications which may differ from those shown; users should consult the vendors to determine 1.10 COMPOSITE TESTING what certification limits exist before commitTo ensure consistent, reproducible compo- ting to specification control. nents, three levels of testing are employed: As part of raw materials verification, comincoming materials testing, in-process testing posite design effort and final product and control and final structure verification. verification mechanical testing of composite test specimens will be performed. The testing of composite materials offers unique chal1.10.1 INCOMING MATERIALS TESTING lenges because of the special characteristics of Incoming materials testing seeks to verify the composites. Factors not considered important conformance of the raw materials to specifica- in metals testing are very important in testing tions and to insure processibility. The levels of composites (Chapters 34,39). knowledge of composite raw materials do not approach those for metals, which can be bought to several consensus specifications and REFERENCES will appear generally identical although pur- 1. Foral, R.F. and Peters, S.T., Composite chased from many manufacturers. Although Structures and Technology Seminar Notes, 1989 there are fewer suppliers for composite raw 2. Hercules Data Sheet for AS-4/3901-6 prepreg H050-377/GF Prod Hdbk (4)/jc/2 materials, the numbers of permutations of resins, fibers and manufacturers prevents the 3. Agarwal, B.D. and Broutman, L.J., Analysis and Performance of Fiber Composites 2nd edn, John kind of standardization necessary to be able to Wiley and Sons, New York, 1990 p. 103 buy composite raw materials as if they were 4. Mayorga, G.D. in International Encyclopedia of alloys. ASTM (American Society for Testing Composites, (ed. S.M. Lee) Vol 4, VCH and Materials), SAE/AMS/NOMETCOM Publishers, N.Y., N.Y., 1991 (Society of Automotive Engineers, Aeronautical 5 . Tsai, S.W. and Pagano, N.J. in Composite Materials Workshop, (eds. S.W. Tsai, J.C. Halpin Materials Standards/ Nonmetallic Materials and N.J. Pagano), Technomic Publishing Co., Committee) and SACMA (Suppliers of Lancaster, PA, 1978, p. 249 Advanced Composite Materials Association) 0 ° / ~ 0 / 9 0 cover 0 most applications, with 0 between 30 and 60 degrees). Unidirectional laminates are rarely used except when the basic composite material is only mildly orthotropic (e.g. certain metal matrix applications) or when the load path is absolutely known or carefully oriented parallel to the reinforcement (e.g. stiffener caps).

2

POLYESTER AND VINYL ESTER RESINS Frank A. Cassis and Robert C. Talbot

2.1 INTRODUCTION AND HISTORY

Organic polymers are divided into two types, reinforced-thermoplastic and thermoset. With thermoset polymers such as unsaturated polyesters and vinyl esters, a chemical reaction cross links the material so that it cannot be returned to liquid form. Other common thermosetting polymers include epoxy and phenolic resins. Thermoset plastics made with polyester and vinyl ester resins represent the major portion of the reinforced plastic composites industry today. Early workers on unsaturated polyesters soon learned that despite the possession of reactive double bonds, these resins were sluggish in reacting with themselves. Even with effective catalysts, they still required high temperatures and lengthy cure times to complete the cross linking reaction. The key to modern day application of unsaturated polyesters was the discovery by Carlton Ellis in 1937l that the addition of reactive monomers, such as styrene, gave mixtures that would copolymerize many times faster than homopolymerization. The styrene addition produced the added benefit of an easily handled liquid material that could be pumped, transported and fabricated into a finished plastic by a myriad of molding processes. Developments during the 1940s accelerated the commercial applicability of unsaturated polyesters to the position they hold today. Styrene became readily available and lower in cost as a result of the US Government's sponHandbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

sored production of styrene-butadiene rubber. At the same time, scientists found that styrenated polyesters could yield high strength, light weight structures when reinforced with glass fibers. They also learned that fiberglassreinforced polyesters had excellent electrical properties and that large structures could be molded at low pressures with low cost tooling. As a result, commercial development proceeded rapidly after World War I1 with materials and molding research moving in many directions. In the 20 years that followed, polyester and vinyl ester resins matured rapidly and by the mid-l970s, the composites fabricator and end user had numerous options with these matrix systems to achieve the desired properties in the finished part. 2.2 POLYESTER RESINS

The reaction of an organic acid with an alcohol results in the formation of an ester. By using a difunctional acid and a difunctional alcohol (glycol)a linear polyester is produced (Fig. 2.1). 0 0 II II H-(-0 - C - R - C - 0

- R -)" -OH

Fig. 2.1

Properties of polyesters can be varied by using different combinations of diacids and glycols. These products are thermoplastic polyesters and they are made with various acids and

Polyester resins 35 glycols such as the following: -

Acids Phthalic anhydride Isophthalic acid Terephthalic acid Adipic acid

GZycols Ethylene glycol Propylene glycol Neopentyl glycol Diethylene glycol

In the esterification reaction with maleic anhvdride, the unsaturated acid isomerizes to the fumarate structure which copolymerizes with styrene much faster than the maleate form. A high degree of isomerization to the fumarate structure is essential to produce an unsaturated polyester with high reactivity. Although the isomerization of maleic anhydride is usually from 65-95% in the esterification reaction, some commercial resins are deliberately formulated with the more expensive fumaric acid to obtain maximum reactivity with the monomer employed.

The reaction product of terephthalic acid and ethylene glycol is the well known polyethylene terephthalate (PET) which is used to make polyester fibers and polyester plastics such as clear plastic bottles for soft drinks. Unsaturated polyesters are produced by replacing part of the saturated diacid with an 2.2.1 UNSATURATED POLYESTER CLASSES unsaturated diacid such as maleic anhydride or fumaric acid (Fig. 2.2). The former is vastly Unsaturated polyesters are divided into types preferred since it is lower in cost, easily han- or classes depending on the structure of the dled and produces only half the water that basic building block. These are orthophthalic, would be generated in the reaction when isophthalic, terephthalic, bisphenol-fumarate, fumaric acid is used. chlorendic and dicyclopentadiene.

CH = CH

t

\

0 = c-0-c = 0 Maleic anhydride H

H II

I1

HOOC- C = C

- COOH

Fumaric acid

Fig. 2.2

The resultant polyester contains reactive double bonds (unsaturation) along the entire polyester chain, which becomes the site for the eventual cross linking to produce the cured plastic (Fig. 2.3). 0

0

0

HO

It II II I II HO ( C-R-C-O-R'-O-C-C=C-C-O-R'-O)n H I

H

Fig. 2.3

Orthophthalic resins These are commonly referred to as ortho or general purpose resins and are usually based on phthalic anhydride, maleic anhydride and propylene glycol. Since the acid groups in phthalic anhydride are on adjacent carbons of the benzene ring, it is very difficult to produce resin molecular weights as high as those achievable with isophthalic and terephthalic acid. Accordingly, resins made from phthalic anhydride have poorer thermal stability and chemical resistance than their iso/tere counterparts.

Isophthalic resins These resins are produced from isophthalic acid and are characterized by greater strength, heat resistance, toughness and flexibility than their ortho cousins. In isophthalic acid, the acid groups are separated by one carbon of the benzene ring which increases the opportunity to produce polymers with greater linearity and higher molecular weight in the esterification reaction (Fig. 2.4).

36 Polyester and viny ester resins

Phthalic anhydride

ester can be three times longer than for an is0 resin. As a result of this, researchers have turned to polyethylene terephthalate scrap from the previously mentioned fiber and plastic operations to develop an economical source of terephthalic polyesters. This scrap can be effectively depolymerized by using different amounts of propylene glycol at elevated temperatures. The glycolyzed product is then reacted with maleic anhydride and diluted with styrene monomer to produce a cost effective terephthalic polyester.

lsophthalic acid

0 c=o ‘OH

Bisphenol A fumarate resins

Terephthalic acid

Fig. 2.4

Therephthalic resins Unsaturated polyesters can be produced from terephthalic acid with the expectation that the resin property improvement obtained in going from phthalic anhydride to isophthalic acid will be matched in going from isophthalic to terephthalic acid. This, however, is not the case and terephthalic resins appear to offer only a slight advantage in heat distortion temperature over their isophthalic counterparts. Other important resin properties such as modulus, hardness and overall chemical resistance favor the is0 resins. Because of its lower solubility and poorer reactivity, therephthalic acid requires the use of esterification catalysts or pressure processing to produce a resin economically. Without these, processing time for a terephthalic poly-

H HO - C I - CH2 - O

o

l

@

These resins are unsaturated rigid polyesters made by reacting bisphenol A with propylene oxide to produce the glycol shown in Fig. 2.5. This propoxylated bisphenol A is then reacted with fumaric acid to form an unsaturated polyester. The bisphenol structure illustrated above imparts a high degree of hardness, rigidity and thermal stability to this particular resin. Chlorendic resins These unique polyester resins are based on HET acid (hexachlorocyclopentadiene) or the anhydride shown in Fig. 2.6. When reacted with an unsaturated acid and a stable glycol such as neopentyl, an extremely rigid unsaturated polyester results with outstanding thermal stability and resistance to oxidizing environments. The inherent chlorine in the resin chain imparts some fire retardancy as well.

H .. 0 - CH2 - C I - OH

I

I

CH3

CH3

Fig. 2.5

CI CQ ;?!

CI-c-CI

Fig. 2.6

-c=o

37

Hydrocarbon Solvents Table 2.72: Hexene-1 (4) cn2 = cn-cn2-cn2-cn2-cn,

FORMULA RESEARCH GRADE

PROPERTIES

PURE GRADE

TECHNICAL GRADE

'Literature values.

Table 2.73: cis-Hexene-2 (4)

FORMULA

cn3-c

Composition. weight percent ~

-

~

_

~

~

0.1 0.2 99.6

~

.

Purity by freezing point, mol % Freezing point, F Boiling point, F _ _ _ -~ -~ __Distillation range, F Initial boiling point Dry point 0 7 6 m Specific gravity of liquid at 6 at 2014 C API gravity at 60 F Density of liquid at 60 F. m a l Vapor pressure at 70 F. psia 100 F, psia -___ 130 F, psia index. 20/0 _ _ Refractive _ ._ Color. Saybolt Acidity, distillation residue Nonvolatile matter. gramdl00 ml -__Flash point. appoximate. F *Literature values.

FORMULA

0.1

-~

99.28 -222.04' 156.00'

0.6 920* _________ 0.68720' 5.760' 2.4' 4.9' 9.1' 1.39761. +30

1

-

cH3-cH cn-cH2-cn2-cn3 PURE GRADE

PROPERTIES

~

Hexene-1 ___-_-.___~_-tranr-Hexene-2 cis-Hexene.2 Hexe!Er-?p.Normal Hexane lsoolefins _ _ _ . _ __Heptene-l tranr-Heptene-3 cis.Heptene.3 tranr-Heptene.2 cis.Heptene2

_ _ _ _ _ _ ~

- c-cn2-cn2-cn, RESEARCH GRADE

PROPERTIES

~~

Table 2.74: Mixed 2-Hexenes (4)

~-

Composition, weight percent .. Hexenel . tiant-Hexene-2 cis Hexene-2 Hexener-3 Normal Hexane lsoolefins -_____ Heptene-l

Purity by freezing point. mol % Freezing point. F Boiling point, F Distillation range. F Initial boiling point Dry point Specific gravity of liquid at 60160 F at2014 c API gravity at 60 F Density of liquid a t 60 F. Ibdgal Vapor pressure a t 70 F. psis 100 F, psia 130 F, psia Refractive index, 2010 Color, Saybolt Acidity. distillation residue Nonvolatile matter, gramdl00 ml Flash point, approximate, F

-

-

trxe

0.8

21 ~ -

-.

_

.

-

~ -- - - -

-

-

03

.-

i:it99Omin

__

tranr-Heptene-3 cis-Heptene3 tranr-Heptene 2 ca-Heptene2

TECHNICAL GRADE

--

-

--__-

-

..

_

..

-

-

-

~

.~

.

~-

~~~

~~

~

____

__-.

155.0 155.1 0.686

155.0 __.___.__ 155.1 -. 0.684 ~

- . . ~

.

!.2._.

1.396 +30 _ -~ neutral _. 0.0005 -5

. ~

~ _ _ _ - . ~ -

75.4 5.69 2.4 ____ 5.0 -~ ... . -~ .~~

~

~~-

74.8 5.71 2.4

~-

5.0

- -~ .-

~~

.. ..

9.2 1.396 +30 neutral -. 0.0005 -5

-..- ~-.

__

.

-

~

__ ~

38 Polyester and viny ester resins In addition to tailoring the resin for specific applications by varying the building blocks, the properties of unsaturated polyesters can often be altered significantly by selection of the esterification process. This is particularly true with isophthalic/terephthalic polyesters which are slower reacting than phthalic anhydride. By using a two stage or modified two stage reaction with these aromatic diacids, the molecular structure of the resultant polyester can be changed to markedly improve heat distortion temperature, hydrolytic stability and chemical resistance2.In the two stage process the aromatic acid and glycol are fully or partially reacted before the faster reacting unsaturated acid is added to the cook. This processing method, compared to charging all ingredients at once (one stage method), also leads to a more random distribution of the unsaturation in the polymer chain which changes the character of the final cross linked network in the cured resin. Cure plays one of the most important roles in the chemical resistance developed by unsaturated polyester resins. Theoretically, the curing reaction should go to completion at room temperature with all the double bonds converted to single bonds in the three-dimensional network. However, complete cross linking is rarely achieved at ambient temperatures. This then will result in reduced chemical resistance and, quite often, poorer than expected mechanical properties. In addition, unreacted diluent (styrene ) can remain in the not-so-well cured polymer leading to major problems when the polyester is used for food grade applications. Accordingly, maximum chemical resistance and certain other property improvements can most often be achieved by utilizing elevated temperatures for ‘post cure’ of the polyester resin finished product. Unsaturated polyester resins are used in the manufacture of a broad range of plastic products. A high percentage of these products utilize reinforcing materials, particularly fiberglass. It is estimated that less than 20% of the polyester resins produced are utilized in appli-

cations which do not involve reinforcing materials. These so-called casting applications include buttons, bowling balls, putties, cultured marble, gel coats and decorative products. The marble industry and the more recently developed polymer concrete industry represent outstanding applications for highly filled unsaturated polyesters which offer very economical materials to the building and construction industry. Fiberglass reinforced polyesters (FRP) are used in the manufacturing of boats, automobile and truck parts, building panels, corrosion resistant equipment such as pipes, tanks, ducts, scrubbers, etc., appliances and business equipment, electrical equipment, construction products such as grating and railing, sporting equipment and consumer products that are almost endless. According to the Composites Institute of the Society of Plastics Industry (SPI), automotive, construction, marine and corrosion resistant equipment are the four largest FRP markets, in that order, in the United States which produces 2.5 billion pounds of FRP annually. Mechanical properties are most often the critical factor in the selection of a polyester resin for a specific application. Testing of mechanical properties for both resin castings and fiberglass remforced composites is carried out using standardized ASTM (American Society for Testing and Materials) tests for all plastics. ASTM D-638

Standard Test Method for Tensile Properties of Plastics ASTM D-790 Standard Test Method for Flexural Properties of Plastics ASTM D-695 Standard Test Method for Compressive Properties of Rigid Plastics ASTM D-256 Standard Test Method for Impact Strength (IZOD) of Plastics ASTM D-648 Standard Test Method for Heat Distortion Temperature of Plastics ASTM D-2583 Standard Test Method for Barcol Hardness of Plastics

Polyester resins 39 As mentioned earlier, glycol selection has a produce a rigid polyester which tends to be significant effect on the properties of poly- hard, brittle and lower in tensile elongation. esters. Ether glycols are of great value in Higher unsaturation also leads to higher heat increasing tensile elongation and impact distortion temperature resins. The latter is also strength which is of great importance in auto- achieved by formulating higher molecular motive, casting and liner applications. A weight resins with the chlorendic, bisphenol A principal deficiency of polyester resins is lack and dicyclopentadiene building blocks. As of alkali resistance because the ester linkages expected, all of these resin classes are more are subject to hydrolysis in the presence of brittle and have low tensile elongation. The caustics. Accordingly, increasing the size of the major exception in this scenario are the glycol has the same effect as reducing the con- iso/terepolyesters. Using the multi-stage procentration of attackable ester linkages. Thus, a cessing methods described earlier, these resins resin containing neopentyl glycol, propxylated can be formulated with reasonably high molebisphenol A, or trimethyl pentanediol will cular weights (more linearity) to give very exhibit improved water and chemical resis- tough resins having a good balance of tentance which is highly important in gel coats, sile/flexural properties plus higher tensile corrosion resistant equipment, construction elongation and heat distortion temperatures. Obviously then, when the end use criteria products and many consumer products. The major effect on polyester physical prop- requires the 'something more' than is offered erties is, however, provided by the by general purpose polyesters (orthophthalics unsaturation content in the polyester polymer. and dicyclopentadienes),the formulator turns Higher unsaturation makes for more cross to iso/terepolyesters which have no disadvanlinking and a stiffer cured composite. tages compared to general purpose resins Accordingly, the formulators' selection of other than slightly higher cost. Table 2.1 summarizes the property and unsaturated acid to saturated acid ratio which determines cross linking density can move the application status for the various classes of resin flexural modulus from rigid to resilient unsaturated polyesters. to very flexible. In most cases, a 1/ 1 ratio will Table 2.1 Properties and applications of unsaturated polyesters

Class

Characteristics

Orthophthalics, dicyclopentadiene

Rigid, resistant to crazing, light in color

Isophthalics/terephthalics

Tough, good impact and overall mechanical properties, resistant to environmental elements and moderate chemical attack. Highly resistant to aromatics Rigid, high heat distortion, highly resistant to oxidizing chemical environments Rigid, high heat distortion, highly resistant to most chemical environments particularly caustics

Chlorendic Bisphenol A fumarates

Uses Boats, tub/shower, spas, marble, consumer products, buttons, corrugated sheet, building panels, seating, decorative products Automotive parts, gel coats, electrical, bowling balls, trays, gasoline, tanks, septic tanks, swimming pools, tooling, aerospace products, corrosion, construction products Corrosion resistant tanks, ducting, stacks, industrial vessels Corrosion resistant tanks, piping, stacks, industrial vessels

40 Polyester and uiny ester resins 2.3 VINYL ESTER RESINS

Although vinyl esters have often been classified as polyesters, they should be designated separately because they are typically diesters with a recurring ether linkage provided by the epoxy resin backbone.

Vinyl ester resins are the most recent addition to the family of thermosetting polymers. Although several types of these resins were synthesized in small quantities during the late 1950s, it was not until the mid-1960s that commercialization, principally by Shell and Dow 2.3.1 VINYL ESTER RESIN TYPES Chemical led the push to establish an extremely important segment of today’s com- Aside from the fire retardant versions of vinyl posite industry. Vinyl esters are unsaturated ester resins which are discussed in the next resins made from the reaction of unsaturated section, there are two basic types of vinyl carboxylic acids (principally methacrylic acid) esters having commercial significance. These with an epoxy such as a bisphenol A epoxy are the general purpose lower molecular resin. The typical structure of a vinyl ester weight vinyl esters and the higher heat resistant vinyl esters with greater cross link resin is shown in Fig. 2.8. density. The structure of vinyl ester resins shows several important features which account for the resultant exceptional properties of vinyl General purpose vinyl esters ester resins. There is an epoxy resin backbone with a high molecular weight that provides These are principally methacrylated epoxies excellent mechanical properties combined made by the reaction of methacrylic acid with with toughness and resilience. Secondly, vinyl a bisphenol A epoxy resin. When dissolved in esters display terminal unsaturation which styrene monomer they provide a thermosetting makes them very reactive. They can be dis- resin with good heat resistance, excellent solved in styrene and cured like a mechanical properties (particularly high tenconventional unsaturated polyester to give sile elongation) and outstanding chemical rapid green strength. Obviously, the vinyl resistance to acids, bases, hypochlorites and ester structure also enables convenient many solvents. homopolymerization which could lead to high heat distortion products. Finally, vinyl esters Heat resistant vinyl esters have much fewer ester linkages per molecular weight which combined with the acid resistant These vinyl esters have higher density cross epoxy backbone, give outstanding chemical linking sites available which leads to a more resistance (acids, caustics and solvents) to this heat resistant polymer network. They are proclass of resins. duced from novolac modified epoxy resins

OH H - CI CH2- 0 G

0 7H2 I

c=o

O

CH3

OH O - CH2-CI - H

7% 0 I

c = o

I C-CH3

C-CH3

CH2

CH2

II

Fig. 2.8

T

I

II

Vinyl ester resins 41 and methacrylic acid which provides more cal properties can be 'tailored' to meet the unsaturation sites and higher molecular requirements of specific applications. Another weight due to the epoxy backbone. These unique property of vinyl ester is the bondabilvinyl esters increase the heat resistance by ity of these resins to other surfaces. They are 17-27°C (30-50°F) over the general purpose not as good as epoxy resins in this charactertypes. This often translates to higher useful istic, but obviously the epoxy resin operating temperatures for vinyl ester based component gives them a boost over other reinforced plastics even in corrosive environ- unsaturated polyesters in this area. A case can ments. The higher-density cross linked vinyl also be made for vinyl esters providing better esters are less resilient (lower tensile elonga- fiberglass wet out in FRP composites due to tion) but still have excellent mechanical the backbone hydroxyl groups and their interproperties. Cure of the higher cross linked action with these groups on the fiber surface. vinyl esters may require the use of different Some fabricators have reported that observperoxide catalysts to reduce the peak able resin savings can be achieved with vinyl exotherm and thereby prevent cracking/craz- esters because of this characteristic. However, vinyl esters such as bisphenol A ing in resin rich areas. In other words, resins of this type are more reactive and more caution is polyesters and chlorendic polyesters are made required in the fabrication of FRP laminates. from higher cost materials and often require extended process times which leads to higher finished cost. Accordingly, the specifier/fabri2.3.2 PROPERTIES/APPLICATIONS cator turns to commercial applications where The development of vinyl esters has led to the the improved performance of vinyl esters can fastest growing segment of the thermosetting justify the premium price of the finished comresin industry today. This is not surprising, posite. The foremost application for vinyl esters is since vinyl esters combine inherent toughness with outstanding heat and chemical resis- in glass reinforced laminates for corrosion tance. In all other thermosetting resin types resistant equipment. Because of outstanding one has to sacrifice some heat resistance and chemical resistance combined with excellent often chemical resistance to increase resiliency mechanical properties, vinyl ester based FRP and toughness. Unlike polyesters, vinyl ester tanks, piping, scrubbers, fans and ductwork resins possess low ester content and low are being specified for waste water treatment unsaturation which results in greater resis- plants, mining facilities, chemical processing tance to hydrolysis, lower peak exotherms and storage units, semi-conductor chip operaduring cure and less shrinkage during cure. tions, pulp and paper manufacturing and odor They are easily dissolved in reactive control facilities. Since FRP corrosion resistant monomers such as styrene which provides equipment is the fastest growing segment of easy handling and transportation to the fabri- the US composites industry, the future for cation site. As with polyesters, other reactive vinyl esters looks extremely strong. They are monomers such as vinyl toluene, chlorostyrene comparable to other premium resins for chemand f-butyl styrene can be employed with few ical resistance and secondary bonding combined with a good balance of chemical problems. The toughness of vinyl esters comes from resistance (acids, bases, solvents) at the same the epoxy resin backbone. Since the molecular or lower cost. As a result, chlorendics and weight and structure of the epoxy resin can be bisphenol A polyesters have been reduced to varied like the polyester resin building blocks, 'niche' applications where their specific propphysical properties such as tensile elongation, erty advantages such as heat resistance and heat distortion temperature and key mechani- resistance to oxidizing environments demand

42 Polyester and viny ester resins their use. Since iso/terepolyesters also give an excellent balance of properties in corrosion applications, these unsaturated polyesters and vinyl esters now dominate the corrosion market. The bonus provided by vinyl esters is of course higher heat resistance and extended life at higher operating temperatures, but at significant additional cost compared with the iso/ terepolyesters. The next major market area for vinyl esters utilizes the high tensile elongation characteristics of these resins to produce linings and coating with outstanding adhesion to other types of plastics and conventional materials such as steel and concrete. For example, vinyl esters are an excellent barrier coat for fiberglass boats and acrylic spas. Vinyl ester corrosion coatings are used everywhere today for steel tank linings and industrial flooring. In dual laminate structures, a vinyl ester is often the back up for exotic thermoplastics or the superior corrosion barrier for lower cost polyesters in many FRP tank and pipe applications. The growth of vinyl esters has also been boosted by their excellent handling characteristics and ease of cure. For example, vinyl esters are much preferred by FRP fabricators in filament winding operations because of excellent glass wet out and in fabrication of large structures because the resins are forgiving and provide predictable curing over a wide range of temperatures. The latter situation has resulted in a virtual exclusive use for vinyl esters in field fabrication of large FRP structures.

Table 2.2 summarizes the resin casting properties of the various resins used in corrosion resistant applications today. The outstanding balance of properties provided by vinyl ester resins is obvious and bodes well for continued strong growth in US corrosion markets. Other significant markets for vinyl esters includes pultruded construction and electrical components, automotive structural applications, polymer concrete vessels for mining and chemical operations, grating, high performance marine applications and sporting goods. 2.4 FLAME RETARDANT VERSIONS

The need for flame retardant polymers is essential in many plastics applications today. The combustibility of plastics has drawn so much attention to the safety aspects of these materials in construction applications, that designers and specifiers have been pressured by fire officials to provide fiberglass-reinforced construction materials that exhibit low flame/low smoke characteristics. Since all plastics are based on organic constituents, they are inherently flammable and once ignited will burn until they are completely consumed. There are, however several methods available for making thermosetting resin flame retardant and these provide the capability to supply fire retardant FRP and corrosion resistant/fire retardant FRP for the numerous applications that have a need for some degree of fire retardancy.

Table 2.2 Resins for corrosion resistant applications

General purpose vinyl ester Heat resistant vinyl ester Chlorendic polyesters Bisphenol A polyester Rigid isopolyester Resilient isopolyester

Tensile, psi

Flexural, psi

12 500 13 000 5 500 10 000 8 500 12 500

20 500 20 000 10 000 16 500 19 000 20 000

Elonga f ion break, % 6.7 5.6 1.4 3.2 1.9 4.4

HDT, OF

221 248 284 288 234 201

Flame retardant versions 43 range of chemical environments, both acid and alkali, at operation temperatures similar to the general purpose vinyl esters. Brominated high molecular weight isopolyesters offer economic advantages and are suitable for moderate corrosion applications. These two resin types have become the workhorses for the waste water/odor control FRP market and the chemical and pulp /paper industries because they exhibit excellent impact properties combined with good overall corrosion resistance. 2.4.1 CHEMISTRY AND APPLICATIONS Variations of these resins are used to meet Flame retardancy of unsaturated polyester MIL-R-21607 or MIL-R-7575 requirements. Dibromoneopentyl glycol formulated with and vinyl ester resin is an extension of the nonflame retardant systems (as discussed above). carefully selected chemical building blocks Almost all of these resins can be reformulated provides resins for exposure to severe weatherto include a halogen in the chemical composi- ing conditions. The construction industry uses tion by either blending or by an in situ cook of these resin systems, which are specially formuthe resin. There is an advantage to locking in lated to meet optimum fire retardance for the the halogen in the original resin cook, in order continuous line products of corrugated and flat to chemically tie in the halogen (Cl, or Br) to sheet panels. Such systems are formulated with prevent migration of the halogen when sub- ultraviolet (W) stabilizers and acrylates to jected to thermal degradation. While flame achieve excellent color stability with acceptable retardancy can be achieved with additives low smoke and flame spread (FS)properties. In (Dekabrom or Dechlorine), these additives most cases, these formulations offer good have not been used for high performance chemical resistance for splash and spill on the applications in either the corrosion or con- exposed surfaces. Highly filled halogenated resin systems are designed to accept high filler struction industries (corrugated FRP panels). The chlorendic resins were developed in the loading with aluminum trihydrate (ATH) and 1950s and were based on HET acid (hexa- other synergists to meet DOT requirements for chloroxyclopentadiene). Other formulations low smoke, low flame spread properties followed, based on either tetrabromo bisphe- (ASTM-E-662 and E-162 respectively). Values no1 A (TBBPA) or dibromo neopentyl glycol of 190°C. (Shimp, 1993; 1994a). 5.6.2 GALVANIC CORROSION

5.6 DESIGN CONSIDERATIONS 5.6.1 SELECTION OF ARAMID FIBER AND CORE

Aramid fiber and core reinforcements for CE composites should be selected from second generation materials wluch absorb 95% RH

1o4

Fig. 5.10 Changes in 3 mm thick bar volumes during water immersion for a period of one year indicate swelling rates and limits of thermoset resins. The ratio of volume increase to total volume of water absorbed (numbers on right) indicates the fraction of water associated with dipoles. A: BMI-MDA; X: BMI-DAB; 0: TGMDA-DDS; 0:AroCy B; 0 :XU-366

Suppliers of prepreg and other formulated products Table 5.6 Properties of BMI/IM-7 unidirectional composites

Mechanical st rengtk

Rigidite 5250-4"

0" Tensile, MPa (ksi) 25°C 0" Compression, MPa (ksi) 25°C Dry 105°C Wet 149°C Wet 177°C Dry 177°C Wet

Rigidite 5260b

2618 (380)

2691 (390)

1820 (235)

1746 (253) 1346 (195) 1276 (185)

-

-

1310 (190) 966 (140)

0" Compressive modulus 25"C, GPa (msi)

158 (23)

152 (22)

Open hole compression, MPa (ksi) 25°C Dry 177°C Dry 191°C Wet

420 (61) 351 (51) 303 (44)

352 (51) 269 (39) 221 (32)

Compression after impact At 4.5 kJ m-l, MPa (ksi) At 6.7 kJ m-l, MPa (ksi)

248 (36) 214 (31)

380 (55) 345 (50)

Edge delamination, MPa (ksi) 25°C

241 (35)

358 (52)

~~

a

Data courtesy of Cytec. Post cure 6 h at 227°C; 60% fiber vol. Data courtesy of Cytec. Post cure 6 h at 215°C; 60% fiber vol.

Table 5.7 Properties of CE unidirectional composites

Cytec 5245C Reinforcementhre Carbon fiber Max. cure temp., "C

Fiberite 954-2

Hexcel HX-1562

IM-6 210

IM-7 232

IM-7 177

0" Tensile, MPa (ksi)

2439 (356)

2814 (408)

2610 (378)

0" Compression, MPa (ksi) 25"C, Dry 121"C, Wet 132"C, Wet 149"C, Wet

1690 1350 1310 987

(245) (196) (190) (143)

1573 (228) 1331 (193)

1700 (246)

214 (31)

262 (38)

317 (46)

262 (38)

269 (39)

Mechanical strength

CAI, MPa (ksi) At 6.7 kJ m-I Edge delamination, MPa (ksi)

-

-

1290 (187)

-

-

1140 (165) -

-

111

112 Speciality matrix resins

0

°

1

1 O0

79 74

61

Fig. 5.11 (left) Moisture plasticization of cast matrix systems is inversely related to the percentage of dry room temperature flexural modulus retained at elevated test temperatures. : at 149°C wet; : at 177°C wet.

Fig. 5.12 (below) Hydrolysis of unsubstituted CE (bisphenol A dicyanate) homopolymer begins to reduce mechanical properties after 200 h exposure to 121°C steam autoclave at 15 psig. Ortho-methylation is an effective technique for increasing hydrolytic stability of cured CE resins in aggressive environments. 0: AroCy B; 0:AroCy M.

~

AroCy AroCy BMI/

% WEIGHT GAIN

121

AROCY B

91

200 400 TIME, HOURS

0

600

Table 5.8 Sources of formulated/compounded CE and BMI products

Supplier Bryte Technologies Cytec Hexcel Fiberite, Inc. YLA

Prepreg CE BMI, CE BMI, CE CE CE

Adhesive foam

Syntactic compound

compound

CE BMI, CE BMI

CE BMI, CE

CE BMI

-

-

-

-

-

-

BM1,CE

CE

CE

CE

-

-

RTM

Compression molding

AppIications 113 +1.0 AROCYBIEKXY

w H.5

P3

I

AROCYM

-I

0

10

20

30

40

50

60

DAYS IMMERSION IN 20% NaOH AT 50°C

Fig. 5.13 Cured CE and BMI resins hydrolyze (etch) in strongly alkaline solutions, as indicated by the onset of weight loss. Ortho-methylated CE resin and blends with epoxy resin (50/50 blend shown) increase resistance to alkaline environments generated in galvanic cells. 5.8 APPLICATIONS

Toughened BMI/carbon fiber composites have been specified as the principal composite material for F-22 fighter primary and secondary structures (Fig. 5.15).BMI service temperatures are sufficiently high for cowlings, nacelles and thrust reversers of jet engines. CE composites

t

Y

TRANSMISSION Fig. 5.14 Interactions of microwaves with a radome wall. were used to construct EFA (Eurofighter) prototypes and are used in construction of the Dassault Rafale. Both materials are candidates for HSCT (High speed civil transport) use. Principal applications for CE composites (McConnell, 1992) include radomes for military aircraft, fighter aircraft retrofitted with improved tracking systems, skins over phase array radar, weather tracking aircraft radar and missile nose cones. CE prepreg reinforced with high modulus pitch-based carbon fibers are preferred materials for earth orbit service,

Fig. 5.15 F-22fighter constructed with BMI composites. Photograph courtesy of Lockheed.

114 Speciality matrix resins demonstrating low outgassing, microcrack resistance and resistance to lo9rads of ionizing radiation (Willis, 1991). Applications in space include communication satellites, solar arrays, parabolic antennas, optical benches and precision segmented reflectors. BMI film adhesives are employed in jet engine or high speed aircraft sandwich panels where hot-wet service up to 190°C is required. CE film and paste adhesives are used together with syntactic foams in the construction of radomes. BMI molding compounds reinforced with up to 65 wt.% of chopped reinforcements are used to mold ducts, drive sprockets for heated rolls in copy machines, helicopter gear boxes and missile strongback mounting supports. REFERENCES Bargain, M. et al. 1971. US Patent 3 562 223. Boyd, J.D. and D.A. Shimp. 1987. US Patent 4 644 039. Boyd, J.D. and Hon-Son R. 1990. US Patent 4 923 928. Boyd, J.D. 1991a. US Patent 5 037 689. Boyd, J. et al. 1991b. Galvanic corrosion effects on carbon fiber composites. Int. S A M P E Symp., 36,1217-1231. Boyd, J.D. and L.N. Repecka. 1993a. US Patent 5 189 116. Boyd, J.D. and G.E.C. Chang. 1993b. Bismaleimide composites for advanced high temperature applications. Int. S A M P E Symp., 38,357-369. King, J.J., Chaudhari M. and Zahir. S. 1984. Nat. SAMPE Conf., 29 392. Lee, F.W. and K.S. Baron. 1991. US Patent 5 045 609. McConnell, V.P. 1992. Tough promises from cyanate esters. Adv. Comp., May/June pp. 28-37. Olesen, K. 1991. Degradation of graphite/polymer composites in the presence of a corroding metal. Read at the High Temple Workshop, 11, Reno, Nevada, 4 Feb 1991.

Rottloff, G. et al. 1977. US Patent 4 028 393. Shimp, D.A. 1988. US Patent 4 785 075. Shimp, D.A., S.J. Ising and J.R. Christenson. 1989. Cyanate esters: a new family of high temperature thermosetting resins. SPE/Case Western Conf. on High Temperature Polymers and Their Uses, 1, 127-140. Shimp, D.A. and J.E. Wenhvorth. 1992. Cyanate ester-cured epoxy resin structural composites. Int. SAMPE Symp., 37,293-305. Shimp, D.A and M. Southcott. 1993. Controlling moisture effects during the curing of high T cyanate ester/aramid composites. lnt. S A M P f Symp., 38,370-379. Shimp, D.A. 1994a. Technologically driven applications. In Chemistry and Technology of Cyanate Ester Resins (I. Hamerton Ed.) Chap 10. Blackie, Glasgow, pp. 282-327. Shimp, D. and B. Chin. 1994b. Electrical properties and their significance for applications. In Chemistry and Technology of Cyanate Ester Resins (I. Hamerton Ed.) Chap 8. Blackie, Glasgow, pp. 230-257. Speak, S.C., H. Sitt and R.H. Fuse. 1991. Novel cyanate ester based products for high performance radome applications. Int. S A M P E Symp., 36,336-347. Stenzenberger, H.D. 1990.Chemistry and properties of addition polyimides. In Polyimides (D. Wilson, P.M. Hergenrother and H.D. Stenzenberger, Eds) Chap 4. Blackie, Glasgow. Stenzenberger, H.D. et al. 1991. BMI/bis(allylphenoxy phthalimide)-copolymers: improved thermal oxidative stability. Int. S A M P E Symp., 36 pp. 1232-1243. Stonier, R.A. 1991a. Stealth aircraft and technology from World War I1 to the Gulf, Part I. SAMPE Journal, 27(4), 9-16. Stonier, R.A. 1991b. Stealth aircraft and technology from World War I1 to the Gulf, Part 11. S A M P E Journal, 27(5), 9-18. Willis, P.B. and D.R. Coulter. 1991. Applications of cyanate resins to spacecraft composites. Paper read at 8th Int. Con$ Composite Materials, ECCM/VIII, Honolulu, 15-19 July 1991. Zahir, Sheik A-C. and A. Renner. 1978. US Patent 4 100 140.

THERMOPLASTIC RESINS

4

Lars A.Berglund

6.1 INTRODUCTION

posites offer advantages. They have very low Thermoplastic composites form a fairly new toxicity since they do not contain reactive group of materials. Commercial prepreg tape chemicals (therefore storage life is infinite). such as CF/PEEK (carbon fiber/polyether Because it is possible to remelt and dissolve etherketone) and later CF/PPS (polyphenyle- such thermoplastics, their composites are also nesulfide) was introduced in the early 1980s. easily recycled or combined with other recyHowever, as early as 1966, Menges reported on cled materials in the market for molding improved static strength and fatigue resistance compounds. In the aerospace market, composites based when epoxy was replaced by polyamide 6 as on toughened epoxies dominate. The potena composite matrix (Menges, 1966).In the mid tially cheaper manufacturing of thermoplastic 1970s there was interest in CF/PSU (Po1YSu1- composites has not yet been realized to the fone) due to expectations of better processing methods and improved toughness characteris- extent necessary to motivate large-scale investtics. However, solvent resistance was found to ment in new manufacturing equipment. be a problem. Composites later introduced However, for the next generation of aircraft, based on semi-crystalline thermoplastics, such interest in thermoplastic composites is high. as PEEK and PPS, which have been introduced Higher flying speeds require higher temperatures in the materials than the maximum more recently, have excellent chemical resistemperature available from epoxy-based comtance and are superior to epoxy-based posites in use today. Since the release of gases composites in this respect. Enthusiasm for thermoplastic composites is during processing and inherent brittleness are serious disadvantages of thermoset polyimides, generated for, basically, three different reasons. First, processing can be faster than for thermoplastic composites are of great interest. In the automotive market, thermoplastic thermoset composites since no curing reaction composites are used extensively. Matched-die is required. Thermoplastic composites only compression molding of glass mat thermorequire heating, shaping and cooling. plastics (GMT), primarily based on glass Secondly, the properties are attractive, in particular, high delamination resistance and fiber/polypropylene (GF/PP), is common, because it permits fast processing cycles for damage tolerance, low moisture absorption fairly large components. In the established and the excellent chemical resistance of semifield of injection molded components, matericrystalline polymers. Thirdly, in light of als are used with long fibers (5-10mm) in environmental concerns, thermoplastic commolding pellets. This leads to improved mechanical properties compared with materials based 0; shorter fibeis (Truckenmueller Handbook of Composites.Edited by S.T. Peters. Published and 1991). in 1998 by Chapman &Hall, London. ISBN 0 412 54020 7

116 Thermoplastic resins 6.2 MANUFACTURING METHODS

The principles for thermoplastic composites processing are very different from those for thermoset composites. During the processing of thermosets, the polymer is initially a liquid which then solidifies due to the formation of a three-dimensional molecular network from chemical reactions. Thermoplastics are in the solid state before processing because of their high molecular weight. They are heated above their softening temperature during

processing, for conversion into a high-viscosity melt. Shaping then takes place and the material solidifies on cooling. In Table 6.1, different manufacturing methods used for thermoplastic composites are outlined. Further discussion of thermoplastic composites processing is available in Chapter 24 of this book and in previous reviews (Carlsson 1991; Cogswell, 1992; Kausch, 1993). As with thermoset composites, materials with low fiber volume fractions show ease of

Table 6.1 Manufacturing routes for composites based on thermoplastic resins -

Manufact zi ring ro ii te Open mold processes 1. Autoclave

2. Filament winding 3. Folding

Outline of fabrication and processing methods Unidirectional or woven fibers pre-impregnated by the resin (prepreg) are used. Other forms of prepreg have reinforcing fibers in combination with the resin as fibers or as powder. The prepreg layers are stacked on the mold surface and covered with a flexible bag. Consolidation is obtained by external pressure applied in an autoclave at elevated temperature. Prepreg tape or tape with the resin as fibers or powder are wound onto a mandrel at pre-determined angles. Heat and pressure are applied to the tape in order to continuously weld it onto the underlying material. Preconsolidated sheets are heated. Simple fixtures are then used to shape the sheets into the desired geometry.

Closed mold processes A mixture of molten thermoplastic and short fibers is injected into a colder 4. Injection molding (short fibers, 0.1-10 mm) metal mold at very high pressure. The component is allowed to solidify and is automatically ejected. Semi-finished sheets of glass mat thermoplastics are heated and placed in 5. Compression molding the lower part of the mold in a fast press. The press is quickly closed and (short fibers, 5-50 mm) pressure is applied so that the material can flow to fill the mold. Technology is also available where the hot molding compound reaches the mold from an extruder. The same principle as for short fiber materials. 6. Compression molding Continuous fibers require special clamping fixtures for the sheets and can (continuous fibers) primarily be used for simple geometries. A stack of prepreg is placed inbetween two diaphragms (superplastic 7. Diaphragm forming aluminium or polymer film). The diaphragms are fixed whereas the prepreg can move freely. The material is slowly deformed by external pressure and the mold. Prepreg tape or tape with the resin as fibers or powder is pulled through 8. Pultrusion a heated die to form beams or similar continuous structures with constant cross-section geometry. The material is allowed to cool and solidify. Dry reinforcing fibers are placed in the mold. Monomers and/or low 9. Resin injection molecular weight polymer with low viscosity are injected, the reinforcement is impregnated. Polymerization to a high molecular weight thermoplastic occurs by mixing of reactive components and/or thermal activation.

Material forms processing but low stiffness and strength. On the other hand, materials with high fiber content have high stiffness and strength but require slow processing and are difficult to shape into geometrically complicated structures. For high fiber content materials, the high viscosity of a molten thermoplastic usually requires some kind of prepreg fabrication step before final processing. The prepregs may need to be combined into the consolidated, semi-finished sheets before the final processing step. Regular autoclave processing can be used for thermoplastic composites. For most highperformance thermoplastics, however, temperatures have to be higher than the typical 177°C used for epoxy-based composites. Often, the composite manufacturer must purchase a new autoclave if this is the preferred processing route. Autoclave processing of thermoplastics has been modeled (Lee and Springer, 1987). Consolidation of the prepreg layers is an important issue. At a given temperature, sufficient time must be available for the polymer molecules to diffuse from one prepreg layer into the other and form strong physical entanglements (Howes, Loos and Hinkley, 1989). In addition, the air initially present in the material must be displaced. For thermoplastic composites, filament winding has demonstrated good economic potential (Egerton and Gruber, 1988). The major problem is in the welding of filaments or the tape onto the underlying composite layers. Heat has been applied by means of a gas flame, IR, laser beam or simply from a hot metal surface. Pultrusion of thermoplastic composites offers potential for faster processing than with thermoset composites (Astrom, Larsson and Pipes, 1991), due to the absence of exothermal heat generation from chemical reactions. Profiles may also be produced by roll-forming techniques similar to those used in metalworking. The shape of existing profiles can be changed. The low-cost folding technique (GE Plastics, 1990) has been used commercially by Fokker and TenCate in Holland for quite large components of fairly simple geometry.

117

Compression molding of glass mat thermoplastics (GMT) is a wide-spread process of great sigruficance in the automotive industry (Berglund and Ericson, 1994).Resin injection of polymerizing prepolymer molecules of low viscosity is in principle the same process as for thermosets although the chemical reactions lead to increased molecular weight rather than to cross-linking. Such a process does not provide the advantages of infinite storage life materials with low toxicity. Diaphragm forming is a processing route where the problem of low extensibility of prepreg-based materials is addressed (Mallon, O’Bradaigh and Pipes, 1989). 6.3 MATERIAL FORMS

Thermoplastic composites are usually supplied as semi-finished materials, with the exception of resin injection materials. In Table 6.2, material forms for thermoplastic composites are presented. Prepregs of high fiber volume fractions (V, = 0.6) may be prepared by solvent-, melt-, prepolymer- or powderimpregnation of the reinforcing fibers. Solvent-impregnation is limited to amorphous resins with high solubility. Melt-impregnation is a technique successfully developed by IC1 (Cogswell, Hezzell and Williams, 1981) producing high-quality prepreg. The resulting prepreg is considered too stiff, for some processing situations with little drapability in comparison with CF/EP (epoxy) prepreg. This problem is addressed in prepolymer- and powder-impregnated prepreg. One example is the FIT-technology where small tubes containing reinforcing fibers and polymer powder are used (Thiede-Smet, 1989). In addition, commingled weaves (prepregs) are available. The resin is present in the form of fibers which are melted during processing to form a matrix. Composites produced from commingled material forms may have a fairly inhomogeneous distribution of fibers (Olson, 1990). Film-stacking is a simple method often used for preparation of laboratory samples

118 Thermoplastic resins Table 6.2 Material forms for composites based on thermoplastic resins

Material forms

Outline of preparation procedure

Prepregs 1. Solvent impregnated

Reinforcing fibers are impregnated by a mixture of solvent and thermoplastic. The solvent is removed by evaporation. Reinforcing fibers are impregnated by thin molten films to which pressure 2. Melt impregnated is applied. 3. Prepolymer impregnated Low viscosity prepolymers are used to impregnate the fibers. Polymerization to high molecular weight thermoplastic takes place during processing. Fibers are enclosed by thermoplastic powder, either in small tubes 4. Powder impregnated containing reinforcing fibers and powder or by powder particles adhering to the fibers from partial melting in a fluidized bed. Other material forms 5. Film stacked composites Stacks with alternating layers of dry fibers and polymer film are heated and compressed. 6. Fiber hybridized weaves Roving of commingled reinforcing fibers and the matrix in fiber form. Weaves are produced from the roving. and roving Reinforcing weaves are impregnated by prepolymer liquid and 7. Prepolymer liquid and polymerized to a thermoplastic composite. dry reinforcement 8. Semi-finished glass mats Supplied as sheets prepared in belt press by extrusion and melting of films which impregnate fiber mats. Porous sheets are produced from slurry (low fiber content) of fibers and thermoplastic powder in water, by technology similar to manufacturing of paper. Recycled material can be used. Often prepared by pultrusion of unidirectional fibers and matrix 9. Pellets followed by chopping into pellets. Used in injection molding or in (low fiber content) plasticizing and/or compounding unit combined with compression moulding. Recycled material is easily incorporated.

(Hartness, 1982) although the technique has also been used commercially. In the category of materials with low fiber volume fraction (V,=: 0.2), semi-finished sheets of GMT-materials are available. They usually have random, chopped or continuous fiber mat reinforcements. Unidirectional prepreg may be used in order to selectively provide additional stiffness, strength a n d creep resistance. A n interesting step forward is provided by extrusion compounded GMT (Composite Products Inc, 1994; Hoechst AG, 1994). No semi-finished sheets are used, instead a special extruder is used to produce a hot, soft 'cake' constituted of chopped fibers and the polymer matrix, often PP. The cake is placed in a press and molded. The investment in technology is higher than for conventional GMT molding.

On the other hand, the material cost and energy consumption is reduced, greater freedom in materials selection is obtained and recycling is facilitated. Suppliers of thermoplastic composites are listed in Table 6.3. 6.4 THERMOPLASTIC RESINS

Thermoplastics have either amorphous or semi-crystalline structure (Sperling, 1992). The large, chain-like polymer molecules do not show long-range order in amorphous thermoplastics, which may be viewed as polymer glasses and, in the absence of color pigments, are usually transparent. Thermosets are also amorphous. In contrast, crystalline polymers have regions of molecular order. In meltprocessed crystalline polymers, a spherical

Thermoplastic resins 119 Table 6.3 Suppliers of thermoplastic composites

Supplier

Materials

Ba ycomp Burlington, Ontario, Canada

Unidirectional tapes. Matrices PP, HDPE, PA12 PC, PEI, PBT, PES, PPS, PEEK, ABS, PPO. Fibers: glass, carbon, aramid and stainless steel.

CYTEC, Anaheim CA, USA

Commingled yams. Carbon fiber with PEEK, PEKEKK, PA6,6, TPI (Aurum@).GF/PA6,6.

DuPont de Nemours Bad Homburg, Germany and Newark, DE, USA

Prepreg based on Avimid@K,thermoplastic polyimide, and carbon fiber. Sheets laminated of continuous fiber thermoplastic composites or unidirectional discontinuous fibers. Molding compounds of lower fiber content. Matrices: PA6,6, PEKK, PET and others. Fibers: carbon, glass and aramid.

Electrostatic Technology Branford, C, USA

Prepreg fabrication by deposition of polymers in powder form on tow and fabrics. Wide variety of resins and fibers.

GE Plastics Amsterdam, Netherlands and Pittsfield, MA, USA

GMT-materials based on glass fiber mats and PP, PBT, PC and blends PC/PBT. Unidirectional GF/PP.

Hoechst Frankfurt, Germany

Unidirectional prepreg of GF/PP, GF/PA6, GF/PE, CF/PPS, CF/PA6. Pellets > 12 mm for use in plasticating extruder combined with compression moulding.

Huls Marl, Germany

GF/PA12 fabric prepreg.

ICI/Fiberite Monchengladbach, Germany and Laguna Hills, CA, USA

APC-2 (CF/PEEK) prepreg tape and tow and developmental materials, primarily for high-temperature applications.

Porcher Textile Lyon, France

FIT-weaves (Thiede-Smet, 1989).Matrices PA12, PEI, PEEK. Glass and carbon fibers. Enichem, Milano, Italy reportedly produces GF/PP, PET, PBT with FIT-technology.

Quadrax Corp Portsmouth, RI,USA

Prepreg fabrics and unidirectional tape, consolidated sheets. Matrices PA6,6, PMMA, PEI, PPS and PEEK. Carbon, glass and aramid fibers.

Schappe Techniques Charnoz, France

Spun yarns combining reinforcing and matrix fibers for subsequent weaving. Matrices PP, PA6, PA6,6, PPS, PC, PEI, PEEK. Carbon, glass and aramid fibers.

Symalit AG Lenzburg, Schweiz

Glass mat thermoplastic sheets based on GF/PP.

TenCate Advanced Composites Nijverdal, Netherlands and Fountain Valley, CA, USA

Prepreg fabrics and unidirectional tape, consolidated sheets. Matrices PES, PEI and PA12. Carbon, glass and aramid fibers.

120 Thermoplastic resins morphology, termed spherulitic, is often observed (Bassett, 1981). Crystalline lamellae are present within the spherulites although disordered regions exist between and within the lamellae. This is because the large size of polymer molecules inhibits perfect crystallization. The crystalline thermoplastics are therefore more correctly described as semicrystalline, since the degree of crystallinity never reaches 100%. Semi-crystalline thermoplastics can be viewed as two-phase materials with a crystalline and an amorphous phase. To illustrate the difference in behavior of semi-crystalline and amorphous thermoplastics, polyethylene terephthalate (PET),may be used as an example. PET is a thermoplastic polyester which crystallizes fairly slowly. Therefore, upon rapid cooling from the molten state, crystallization can be suppressed and an amorphous polymer is obtained (similar behaviour is shown by PEEK). Samples of PET with different degrees of crystallinity can be produced by changing the conditions of cooling. The shear modulus G' (obtained from dynamic mechanical thermal analysis, DMTA) is plotted against temperature for such samples in Fig. 6.1. The

I

I

I

50

100

I

amorphous sample shows a dramatic drop in modulus at the Tg (glass transition temperature). The drop in modulus for semi-crystalline samples is less dramatic: the higher the crystallinity, the slower the drop. Above T,, material modulus is maintained by the crystalline phase (although strength usually decreases dramatically). Another effect of reduced degree of crystallinity is reduced chemical resistance. A disadvantage with semi-crystalline polymers is the high processing temperature, see Tables 6.4 and 6.5, compared with the heat deflection temperature (see next section, Table 6.11). The melting temperature, Tm,of the crystalline phase must be exceeded during processing, although the maximum use temperature, as for amorphous polymers, is still below T,. Characteristic temperatures of thermoplastics used in applications where only moderate temperatures are experienced are presented in Table 6.4. These materials are available from many different chemical companies, therefore trade names and suppliers are not listed. In Table 6.5, thermoplastics for applications at higher temperatures are listed. These polymers

I I

150 Temperature ("C)

I

I

200

250

Fig. 6.1 Shear modulus (G') compared with temperature for PET of different degrees of crystallinity.

Thermoplastic resins

121

Table 6.4 Characteristic temperatures for thermoplastic resins with T,< 90°C

Polymer type Polyolefin

Chemical name

___

( "C)

Processing temp. ("C)

Crystalline

-10

165

200-240

Crystalline

55

265

270-320

Crystalline

35

180

220-260

Crystalline

70

265

280-310

Crystalline

20

240

260-290

-

~

Polypropylene

T,,,

T ("C)

Structure

___

(PP) Polyamides

Polyamide 6,6 (pA6,6) Polyamide 12 (PA12)

Polyesters

Polyethylene terephthalate (PET) Polybutylene terephthalate (PBT)

Table 6.5 Characteristic temperatures for thermoplastic resins with T, 290°C

Polymer fYPe Polyester Polyarylene ether or sulfide

Polysulfones

Chemical name ~~

~~

Polycarbonate (PC) Polyphenylene sulfide (PPS) Polyarylene sulfide PEEK PEEKK PEKK PEKEKK Polyketone Polysulphone (PSU) Polyethersulfone (PES)

T,

Trade name and supplier

Structure

Lexan, GE Makrolon, Bayer Ryton, Phillips

Amorphous

150

none

280-330

crystalline

90

280

300-340

PAS-2, Phillips

Amorphous

215

none

330

Victrex PEEK, IC1 Hostatec, Hoechst Declar, DuPont Ultrapek, BASF Victrex HTX, IC1

Crystalline Crystalline Crystalline Crystalline Crystalline

143 165 155 175 205

343 365 340 375 385

380400 390-415 380400 400420 420430

Udel P1700, Amoco

Amorphous

190

none

300-350

Victrex 4100G, IC1 Ultrason E, BASF

Amorphous

220

none

300-320

Torlon C, Amoco Torlon AIX-159 Amoco

Amorphous Amorphous

275 290

none none

350400 350400

Amorphous

217

none

335420

Amorphous Crystalline

250 260

none 390

340-360 400420

Processing temp. ("C)

("Ci

~~

Polyamideimides

Polyamideimide (PAI)

Polyimides

Polyetherimide Ultem, GE (PW Polyimide (TPI) Avimid KIII, Du Pont Polyimide (TPI) Aurum, Mitsui Toatsu

122 Thermoplastic resins of static mechanical properties of the resins is given in Table 6.6. Resins with low Tg,such as PP and PA12, have lower modulus and strength. Their fracture toughness is high and valid data according to linear elastic fracture mechanics are difficult to obtain. Among polymers with T, well above room temperature, the modulus is fairly similar. It is controlled by weak physical forces between the molecules. Viscoelastic effects such as creep and stress relaxation during loading will affect the data. Tensile strength varies more widely than modulus between different resins. As a material property it is unfortunately not very reliable. It is sensitive to loading rate, specimen geometry, specimen preparation and the presence of microscopic flaws on the specimen surface. In addition, uniaxial resin tensile strength is different from resin strength in the composite where the stress state is different. Thermoplastics have higher fracture toughness 6.5 PROPERTIES AND DESIGN than epoxy and other thermosets, although CONSIDERATIONS epoxy fracture toughness can be improved by In contrast to thermoset resins, thermoplastics addition of a thermoplastic (Bucknall and can be dissolved and melted. In general, vis- Gilbert, 1989) or other means. Although not coelastic and plastic effects are more apparent from the table, epoxy modulus is usupronounced in thermoplastics. A presentation ally slightly higher than for thermoplastics.

are more expensive and are often termed ’highperformance’ resins. The higher cost of these materials is due to small material volumes, more expensive monomers and more difficult polymerization procedures. Many resins used in injection molding are so called blends, physical mixtures between two thermoplastics. In the field of commercial composite materials, this technology is primarily used for GMT-materials, where composites based on PC/PBT blends are available (Table 6.3). However, for high-performance resins, blending amorphous with semi-crystalline thermoplastics is an interesting route to improved chemical resistance. Most polymer mixtures form immiscible twophase structures although PEEK and PEI may be mixed to form a miscible blend (Crevecoeur and Groeninckx, 1992).

Table 6.6 Mechanical properties of thermoplastic resins Material -__

PP PA6,6 PET PC Amorph. PA a-2) PPS (Ryton) PAS (PAS-2) PEEK PSU (Udel P1700) PES (Victrex) PA1 (Torlon) PEI (Ultem) TPI (Avimid K-111) TPI (LaRC-TPI) EP(thermoset)

Tensile modulus E (GPu) 1.l-1.6 2.5-3.8 2.74.0 2.3-3.0 3.2 3.5 3.2 3.1-3.8 2.5 2.6 2.84.4 3.0 3.5 3.7 2.8-3.5

Tensile strength 0,( M W 3040 50-80 50-70 60-70 100 80 100 90-100 70 80 90-190 105 100 120 40-120

Fract lire toughness G,? (kJ m-2) -

-

1.6 0.5-0.9 -

4.0 2.5 1.9 3.4

3.3 1.5 1.8 0.1-0.5

Properties and design considerations 123 Many mechanical properties of composites are dominated by the influence of fiber modulus, fiber strength and fiber volume fraction. This is usually true for longitudinal tensile modulus and strength as well as flexural modulus and strength. For thermoplastic composites based on AS-4 carbon fiber, typical tensile data are: longitudinal tensile modulus E,= 130 GPa, longitudinal tensile strength (5, = 1950 MPa. In the present context, we are more interested in properties dominated by the matrix and the fiber/matrix interface. One such property is the transverse tensile strength of unidirectional laminates. When a multidirectional laminate is loaded in in-plane tension, the first major damage mechanism is likely to be matrix cracking in the plies with transverse orientation to the maximum load direction. This reduces laminate stiffness and initiates other damage mechanisms such as delamination. In Table 6.7, transverse strength and modulus are presented for different thermoplastic composites. Fiber volume fractions are high, V ,= 0.5-0.6, For composites based on brittle epoxies, typical transverse strength is 40 MPa. Composites

based on LaRC-TPI, J-2, PAS-2 and K-111 show transverse strengths in the range 3 2 4 1 MPa. Otherwise, typical transverse strengths for thermoplastic composites are in the range 60-90 MPa. Toughened epoxy composites also show fairly high transverse tensile strengths, typically around 75 MPa. The use of transverse tension data in failure criteria will lead to conservative estimates. Data are higher for transverse plies in multidirectional laminates (Berglund, Varna and Yuan, 1991). The modulus data for carbon fiber composites in Table 6.7 appear insensitive to small differences in matrix modulus. Variations in fiber volume fraction and transverse fiber modulus between the materials mask any such effect. The detrimental effect of glass fiber as opposed to carbon fiber is apparent from the GF/PA6,6 and CF/PA6,6 data. GF/PP shows very poor performance, probably due to poor fiber/matrix interfacial adhesion (note the low V, ). Interfacial weakness is also likely to explain the low strength for Kevlar / PEKK. For thermoplastic composites based on AS-4 carbon fiber, Table 6.7 can be used to estimate typical data: transverse

Table 6.7 Transverse tensile properties of thermoplastic composites

Material

Transverse modulus Transverse strengfh ET (GPa) (MPa) ~~~

GF/PP GF/PA6,6 CF/PA6,6 CF/Am. PA CF/PPS CF/PAS CF/PAS CF/PSU K49/PEKK CF/PEKEKK CF/PEEK CF/PAI CF/TPI CF/TPI CF/TPI CF/EP CF/EP

Plytron (V, = 0.35) E-glass/Ultramid G30-500/Ultramid AS-4/J-2 AS-4/Ryton T 650-42/Rade18320 AS-4/PAS-2 AS-4/Udel P1700 Kevlar/PEKK LDFTM G30-500/Ultrapek AS-4/PEEK (ICI) T650-42/PAI-696 AS-4/K-III G30-500/’NewTPI’ AS-4/LaRC-TPI AS-4/3501 (thermoset) HTA/6376C (thermoset)

4 8.6 7.2 9 7.6 8.4 8.3 7 6.2 10.3 8.9 7.6 9 8.3

9 9.9

11 48 72 35 72 61 32 59 21 90 80 67 41 59 33 52 75

124 Thermoplastic resins tensile modulus E, = 8.6 GPa, transverse tensile strength (T, = 75 MPa. High interlaminar toughness is desirable since this suppresses the tendency for delamination crack formation during loading. Interlaminar fracture toughness is determined on double cantilever beam specimens (DCB), usually unidirectional materials are used (Whitney, Browning and Hoogsteden, 1982). In Table 6.8 such data are presented. CF/PEEK shows the highest fracture toughness. All thermoplastic composites show higher toughness than the thermoset composites. There is a difference between crack initiation and crack propagation data (Davies, Benzeggagh and de Charentenay, 1987). The data presented here are crack propagation data; crack initiation data are in general much lower. For tough matrices one may question the applicability of the data to design problems. In DCB experiments, the crack opening displacement (COD) is very high, whereas the COD at small central cracks in stiff laminates is much smaller. Local stress fields and damage mechanisms may therefore be different and affect the measured fracture toughness. At present, delamination fracture toughness from DCB tests are therefore preferentially used to compare material. It has been pointed out that composite data are significantly lower than resin data (Hunston, 1984).In tests

on neat resins, energy is absorbed by yielding and other types of damage when the volume of material is relatively large. In a composite, the presence of fibers tends to limit this material volume. For shear strength, no comparable data for different thermoplastic composites appear to be available in the literature. The interlaminar fracture toughness in mode I1 shear loading, Glrc,is higher for thermoplastic than for comparable thermoset composites (Cantwell and Davies, 1993). This also indicates a higher shear strength for the thermoplastic composites. A typical value for the in-plane shear modulus of thermoplastic composites based on AS-4 carbon fiber is 4.8 GPa which is similar to toughened CF/EP systems but slightly lower than for brittle matrix CF/EP composites. Compressive strength is lower for thermoplastic than for thermoset composites (Table 6.9). Most of the thermoplastic composites are in the range 900-1100MPa whereas typical thermoset composite data are 1700 MPa. Fiber misalignment, shear stiffness and strength have been shown to affect compression strength based on plastic kink band formation (Budiansky, 1983). Compression modulus data in Table 6.9 are similar, in support of similar fiber volume fractions for the materials compared. The composite based on PA6,6 has the lowest strength; PA6,6 also has

Table 6.8 Interlaminar fracture toughness of thermoplastic composites Material

Trade name _

CF/Amorph. PA CF/PPS CF/PEEK CF/PEEK CF/PSU CF/PAI CF/PEI CF/TPI CF/TPI CF/EP CF/EP

_

-~ ~

~

AS-4/ J-2 AS-4/Ryton AS-4/Victrex PEEK IM7/Victrex PEEK AS-4/Udel P1700 T650/Torlon AIX-159 T300/Ultem 1000 AS-4/Avimid K-I11 AS-4/ LaRC-TPI AS-4/3501-6 (thermoset) IM7/8551-7 (thermoset)

Fracture toughness qc(kJ~m-2) ~ 1.3 0.9 2.1 2.5 1.2 1.3 0.9 1.8 0.8 0.2 0.5

Properties and design considerations

125

Table 6.9 Compressive properties of unidirectional thermoplastic composites Material

Trade name

Compression strength (MPa)

__

CFDA6.6 CF/Amorph. PA CF/PPS CF/PAS CF/PEEK CF/PEEK CF /PEKEKK CF/PSU CF/PAI CF/TPI CF/EP CF/EP

G30-500 /Ultramid AS-4/ J-2 AS-4 / Ryton AS-4/ PAS-2 AS4/Victrex PEEK IM7/Victrex PEEK AS-4/Ultrapek AS-4/Udel P1700 C-3000/Torlon C AS-4/Avimid K-111 AS-4/ 3501-6 (thermoset) HTA/6376C (thermoset)

700 1100 940 900 1100 1140 1310 1040 1380 1000 1720 1720

Modulus (MPa)

-~

110 -

130 120 120 -

127 -

110 140 130

the lowest creep modulus and yield stress of superior performance to first generation therthe investigated matrices. It is interesting to moset composites (AS-4/3501-6). This is note that AS-4/PEEK and IM-7/PEEK have because the delaminated area due to the roughly the same strength although the IM-7 impact event is more limited for the thermofiber has higher modulus. The smaller diame- plastic composites. However, toughened ter of the IM-7 fiber appears to have a negative epoxy resin composites combined with tougheffect as expected from Euler-buckling consid- ened interlayers between the plies do in erations. general show as good compression strength Compression strength after impact, a mea- after impact as thermoplastic composites. In sure of laminate and material damage fatigue, delamination resistance is higher for tolerance (Dorey, 1989), is presented in AS-4/PEEK compared with epoxy composites Table 6.10. A quasi-isotropic laminate of given (Gustafsson, 1988). However, in uniaxial tenlay-up and geometry is subjected to impact of sion, brittle CF/EP was found to be superior to a certain energy. Internal damage mechanisms both toughened CF/EP and the thermoplastic such as matrix cracking and delamination composite (Curtis, 1987). Claims have been occur in the laminate. The plate is then sub- made that this observation is due to heating jected to compressive load and the stress and effects in the thermoplastic composite specistrain at failure can be determined. The data mens from testing at high frequency (Moore, show that thermoplastic composites have 1991). Table 6.10 Compression strength after impact of thermoplastic composites Material

CF / PPS CF/PEEK CF / PA1 CF/EP CF/EP

Trade name

AS-4 /Ryton AS-4/Victrex PEEK C-3000/Torlon C AS-4/3501-6 (thermoset) AS-4/8551-7 (thermoset)

Compression strength after impact (impact energy) (28 J) (42 1) (571) o (MPa) o (MPa) o (MPa) 221 179 331 310 290 365 345 317 179 145 131 303 -

126 Thermoplastic resins Table 6.11 Glass and heat deflection temperatures for amorphous thermoplastics

Table 6.12 Glass melting and heat deflection temperatures for semi-crystalline thermoplastics

Material

Material

T, ("C)

PP PA6.6 PET PPS PEEK TPI (Aurum)

-10 55 70 90 143 250

PC Amorph. PA PSU PES PA1 PEI EP (thermoset)

T, ("C)

HDT ("C)

150 160 190 220 290 217 200

132 154 175 203 278 200 180

Increased market need for polymer composites with good performance at elevated temperature has generated interest in thermoplastic composites. Materials with continuous use temperatures above 150°C are of particular interest since they perform better than epoxies. One question is how maximum use temperature relates to Tg. In Table 6.11, Tg and the heat deflection temperature (HDT) for amorphous thermoplastics are presented. HDT is determined by subjecting the material to static load (typically 1.8 MPa) and slowly increasing the temperature. HDT is determined as the temperature at which a critical deflection of the sample is obtained. Table 6.11 shows HDT to be 620°C below the T of amorphous poly6 mers. In comparison with room temperature strength, the strength of the composite is significantly reduced above the HDT. The reason

T,,, ("C) 165 265 265 280 343 388

Structure

PA6,6 PEEK PPS PEI PSU

Crystalline Crystalline Crystalline Amorphous Amorphous Amorphous Amorphous Amorphous Amorphous

PA1

PES Am.PA PC

Hydraulic fluid

0 0 0 A 0 A A -

60 75 41 135 160 238

is that molecular mobility in the polymer is increased dramatically as the temperature approaches T . In Table 8.12, similar data to those in Table 6.11 are presented for semi-crystalline thermoplastics. HDT is usually somewhat higher than Tg. However, for some semicrystalline polymers, HDT is below T (as for amorphous polymers). Creep effects 'will be very strong close to and above Tg.For this reason the maximum temperature for continuous service under significant load is unlikely to exceed a temperature of 20°C below T' for semi-crystalline thermoplastics. The primary advantage of crystallinity is therefore chemical resistance. This is apparent from Table 6.13, where chemical resistance for different thermoplastics is indicated in a qualitative way. Semi-crystalline thermoplastics have much

Table 6.13 Chemical resistance of thermoplastics

Material

HDT ("C)

Chlorinated Ketones kydrocarbons 0 0 0 D D 0 D A D

0 = no effect, A = is absorbed, D = is dissolved

0 0 0 A A 0 A A A

Esters H,O abs. (Yo) 0 0 0 A A 0

A A

8 0.5 0.5 1.2 0.9 2 4 0.3 5 -

Applications better chemical resistance than the amorphous polymers. A notable exception is the high water absorption in PA6,6 caused by hydrophilic groups in its chemical structure. The solvent resistance of a large selection of different thermoplastic composites has been reported (Johnston, Towel1 and Hergenrother, 1991). For most thermoplastics, e.g. PEEK, moisture expansion coefficients of the carbon fiber composite may be taken as 0. Thermal expansion coefficientshave been characterized (Barnes et al., 1990) and are similar to epoxy composites. For AS-4/PEEK, the composite density is 1600 kg m-3. For polymers with high Tg, exposure to elevated temperature may lead to increased density not connected with crystallinity but with the amorphous state. The phenomenon is termed physical aging and leads to a more brittle behavior of the polymer (Kemmish and Hay, 1985). Further work is needed to elucidate the importance of physical aging to composite fracture behavior under practical service conditions. For composites processed at high temperatures, residual stresses will also affect fracture behavior. The magnitude of the residual stresses and, consequently, detrimental effects will increase with increasing cooling rate (Manson and Seferis, 1992). Thermoplastic composites can be joined by the same methods as thermoset composites. Bolted joint performance has been compared for thermoset and thermoplastic composites (Walsh, Vedula and Koczak, 1989) with results in favor of thermoplastic composites. With the semi-crystalline thermoplastics, adhesive bonding requires careful surface preparation (Kinloch and Taig, 1987).This is because of the good chemical resistance and limited solubility of these polymers. However, with careful surface preparation, as good adhesive bonds are obtained as with thermoset composites. The thermoplastic nature of the matrix offers another possibility: fusion bonding. Different methods have been compared (Davies and Cantwell, 1993), including hot gas, IR, laser, ultrasonic, vibration, electrical resistance and

127

induction welding. There are difficulties in controlling the processes and very few methods offer the promise of portable equipment. Various bonding technologies for PEEK composites have been compared (Silverman and Criese, 1989).The study favored a technique for fusion bonding with a polyetherimide film. This approach can also be used in repair of damaged structures, since the temperature needed is below the T, of PEEK. 6.6 APPLICATIONS

Thermoplastic composites can be used in similar applications to thermoset composites. The following examples will demonstrate some of the reasons for choosing a thermoplastic composite material. In Europe, the automotive market for GMT composites is significant. Rapid processing by compression molding, cycle times of typically 30s even for large structures, results in cheaper components. In addition, more functions can be integrated into each component compared with sheet metal structures. Bumper beams dominate the automotive market in the US whereas, in Europe, a wider variety of applications are in commercial production. Typical components are subjected to minor loads or impact and surface appearance is not important. Battery trays, beams supporting the hood, seat supports, oil trays, engine shields and even the complete front end have been produced for Volvo, Volkswagen and others. In the aerospace market, DuPont has supplied thermoplastic polyimide composites to the prototype programs for the F-22 fighter aircraft. In the supersonic civil aircraft program, the same thermoplastic polyimide is considered for wing skins. The main reason for this particular thermoplastic composite is a continuous high maximum temperature. AS-4/PEKK (unidirectional discontinuous fibers) was used by Bell Helicopter Textron in a V-22 tiltrotor thermoplastic wing rib for better open hole compression behavior than thermoset composites for high proportions of

128 Thermoplastic resins Bike helmets are also produced (Fig. 6.2). A semi-finished sheet is heated by IR, transfered to heated dies where a selective clamping system is used to hold the sheet. Forming and consolidation pressure is applied and the mold is cooled before demolding. Hoechst in Germany present applications made by winding, a pressure vessel, tubes, sealings and support rings (Fig. 6.3). Matrices include PA6 and PPS with carbon and glass fibers. In the field of biomedical applications, such as hip prostheses, semi-crystalline thermoplastic matrices offer good potential due to their chemical stability (Williams and Fig. 6.2 Bike helmet, an example of a commercial McNamara, 1987). application of thermoplastic composites (Courtesy: DuPont Europe).

45" ply angles (probably for decreased tendency to delamination). A female steel tool and matched-die press forming was used (Chang, 1992). A stretch forming process was used to fabricate C-section curved fuselage frames for Boeing Helicopter from AS4/PEKK (Chang, 1992). TenCate in Holland supplies materials for Fokker Special Products. Landing flap ribs and impact resistant ice-protection plates are produced for the Dornier 328 aircraft.

REFERENCES Astrom, B.T., P.H. Larsson and R.B. Pipes. 1991. Experimental Investigation of a Thermoplastic Pultrusion Process. 36th Int. SAMPE Symp., pp. 1319-30. Barnes, J.A., Sims I.J., Farrow G. et al. 1990. Thermal Expansion Behaviour of Thermoplastic Composite Materials. J Thermoplastic Composites 3: 66-80. Bassett, D.C. 1981. Principles of Polymer Morphology. London: Cambridge University Press. Berglund, L.A. and Ericson M.L. 1994. Glass Mat Reinforced Polypropylene. In Polypropylene: Structure, Blend and Composites. Vol 3: Composites, ed. J, Karger-Kocsis. London: Chapman and Hall. Berglunh L.A., Varna J. and Yuan J. 1991. Effect of Intralaminar Toughness on Transverse Cracking Strain in Cross-Ply Laminates. Adv. Comp. Mater. 1:225-34. Bucknall, C.B. and A.H. Gilbert. 1989. Toughening of a Tetrafunctional Epoxy Resins Using Polyetherimide. Polymer. 30: 213-18. Budiansky, B. 1983. Micromechanics.Computers and Structures. 16 6-10. Cantwell, W.J. and Davies P. 1993. Short-Term Properties of Carbon Fibre PEEK Composites. In Advanced Thermoplastic Composites, ed. Kausch, H-H. pp. 173-191. Munich: Hanser. Carlsson, L.A., ed. 1991. Thermoplastic Composite Materials. Composite Materials Series, Vol 7. New York: Elsevier. Chang, I.Y. 1992. Thermoplastic Matrix Composites Development Update. 37th Int SAMPE Symp.

I .:'i !

1

Fig. 6.3 Thermoplastic composite applications made by winding: pressure vessel, tubes, sealing and support rings. (Courtesy: Hoechst AG)

References 129 Available from DuPont as report entitled Overview of Thermoplastics Composites Technology. Cogswell, EN. 1992. Thermoplastic Aromatic Polymer Composites. Oxford: Butterworth-Heinemann. Cogswell, EN., Hezzell D.J. and Williams P.J. 1981. US Patent 4 549 920. Composites Products Inc 1994. Brochure of CPI process. Winona, Minnesota, USA. Crevecoeur, G. and Groeninckx G. 1992. MeltSpinning of in situ Composites of a Thermotropic Liquid Crystalline Polyester (TLCP) in a Miscible Matrix of Polyether etherketone (PEEK) and Polyether imide (PEI). Polymer Composites 13: 244-50. Curtis, P.T. 1987. In Investigation of the Tensile Fatigue Behaviour of Improved Carbon Fibre Composite Materials. 6th lCCM and 2nd ECCM Proc., Vol 4. pp. 4.544.64. London: Elsevier Applied Science. Davies, P., Benzeggagh M.L. and de Charentenay F.X. 1987. The Delamination Behavior of Carbon Fiber Reinforced PPS. SAMPE Quarterly 19: 19-24. Davies, I? and Canhvell W.J. 1993. Bonding and Repair of Thermoplastic Composites. In Advanced Thermoplastic Composites, ed. Kausch, H-H, pp. 337-366. Munich: Hanser. Dorey, G. 1989. Damage Tolerance and Damage Assessment in Advanced Composites. In Advanced Composites, ed. 1.K Partridge, pp. 369-98. London: Elsevier Applied Science. Egerton, M, and Gruber M. 1988. Thermoplastic Filament Winding Demonstrating Economics and Properties Via In Situ Consolidation. 33rd lnt. S A M P E Symp., pp. 3546. GE Plastics, 1990. Technopolymer Structures, Design and Processing guide. TPS-400A (6/90)RTB. Gustafsson, C-G. 1988. Initiation and Growth of Fatigue Damage in Graphite/Epoxy and Graphite/PEEK Laminates. PhD thesis 88-9. Royal Inst of Techn, Stockholm, Sweden. ISSN 0280-4646. Hartness, J.T. 1982. Polyether-etherketone Matrix Composites. 14th Natl SAMPE Symp., pp. 2643. Hunston, D.L. 1984. Composite Interlaminar Fracture: Effect of Matrix Fracture Energy. Composites Technology Review 6: 176-80. Hoechst AG. 1994. Compela product brochure. Frankfurt am Main, Germany. Howes, J.C., Loos A.C. and Hinkley J.A. The Effect of Processing on Autohesive Strength Development in Thermoplastic Resins and

Composites. In Advances in Thermoplastic Composite Materials. ed. G.M. Newaz, ASTM STP 1044, pp. 3349. Philadelphia: American Society for Testing and Materials. Johnston, N.J., Towel1 T.W. and Hergenrother P.M. 1991. Physical and Mechanical Properties of High-Performance Thermoplastic Polymers and their Composites. In Thermoplastic Composite Materials. ed. L.A. Carlsson, Composite Materials Series, Vol 7, pp. 27-71. New York: Elsevier. Kausch, H-H, ed. 1993. Advanced Thermoplastic Composites. Munich: Hanser. Kemmish, D.J. and Hay J.N. 1985. The Effect of Physical Ageing on the Properties of Amorphous PEEK. Polymer 26: 905-12. Kinloch, A.J. and Taig C.M. 1987. The Adhesive Bonding of Thermoplastic Composite. J . Adhesion 21: 291-302. Lee, W.I. and Springer G.S. 1987. A Model of the Manufacturing Process of Thermoplastic Matrix Composites. J. Comp Mater. 21: 1017-55. Manson J-A.E. and Seferis J.C. 1992. Process Simulated Laminate (PSL): A Methodology for Internal Stress Characterization in Advanced Composite Materials. J. Comp Mater. 26: 405-31. Mallon, P.J., O'Bradaigh C.M. and Pipes R.B. 1989. Polymeric Diaphragm Forming of Continuous Fibre Reinforced Thermoplastic Matrix Composites. Composites 20: 48-56. Menges, G. 1966. The reinforcement of plastics. Kunststoffe 56: 818-23. (In German). Moore, D.R. 1991. In Thermoplastic Composite Materials. ed. L.A. Carlsson, Composite Materials Series, Vol 7, pp. 331-69. New York: Elsevier. Olson, S.H. 1990. Manufacturing with Commingeled Yarns, Fabrics and Powder Prepreg Thermoplastic Composite Materials. 35th Int SAMPE Symp., pp. 1306-19. Silverman, E.M. and Criese R.A. 1989. Joining Methods of Graphite/PEEK Thermoplastic Composites. SAMPE Journal 25: 34-38. Sperling, L.H. 1992. Introduction to Physical Polymer Science. New York: John Wiley & Sons. Thiede-Smet, M. 1989. Study of Processing Parameters of PEEK/Graphite Composite Fabricated with 'FIT' Prepreg. 33rd Int SAMPE Symp., pp. 2086-2101. Truckenmueller, F, and Fritz H.G. 1991. Injection Moulding of Long Fibre-Reinforced Thermoplastics: A Comparison of Extruded and Pultruded Materials with Direct Addition of

130 Thermoplastic resins Roving Strands. Polymer Eng. Sci. 31: 1316-29. Walsh, R, Vedula M. and Koczak M.J. 1989. Comparative Assessment of Bolted Joints in a Graphite Reinforced Thermoset vs Thermoplastic. S A M P E Quarterly 20: 15-19. Whitney, J.M., Browning C.E. and Hoogsteden W. 1982. A Double Cantilever Beam Test for Characterizing Mode I Delamination of

Composite Materials. J. Reinforced Plas. and

Comp. 1:297-313. Williams, D.F. and McNamara A. 1987. Potential of Polyetheretherketone (PEEK) and CarbonFibre-Reinforced PEEK in Medical Applications. J. Muteu. Sci. Lett. 6 : 188-90.

FIBERGLASS REINFORCEMENT*

7

Dennis J. Vaughan

7.1 INTRODUCTION

(2300°F). The molten glass then flows directly to the fiber-drawing furnace in a direct melt The history of glass is ancient, but its engiflow process (Fig. 7.1) or into a marble making neering scientific development is recent. machine. These marbles can be sorted and can Glass was first produced some 4000 years ago, probably in Egypt in the furnaces used to eventually be remelted and drawn into fibers. Continuous glass fibers are produced when produce pottery. Its first application was as a molten glass from the fiber-drawing furnace is form of adornment in jewelry and as added gravity fed through numerous tiny openings decoration to vases and drinking vessels. in a platinum alloy tank called a bushing The use of glass in fiber form dates back to (Fig. 7.2). The droplets of molten glass that the early seventeenth century when the extrude from each of the bushing’s openings Venetians utilized it to create specialized (Fig. 7.3) are gathered together, mechanically gowns. However, commercial fiberglass did attenuated to the correct dimensions, passed not become a reality until in 1939 the joint through a water spray and over a revolving research efforts of Owen-Illinois and Corning belt that applies a protective and lubricating Glass Works, resulted in the formation of coating known as a size or binder. The fibers Owens-Corning Fiberglas Corporation. are then gathered together in a suitably Textile fiberglass has now grown into a shaped shoe to form a bundle called a strand multi-million dollar industry. Glass fiber can which is wound onto a core at approximately be obtained as a continuous fiber on staple or 190 km/h (120 mile/h). This package of fibers discontinuous fiber. Both forms are made by is then dried or conditioned prior to further the same manufacturing process until the fiber processing and eventually sold as a continudrawing operation. ous filament yarn. Staple fibers are produced by passing a jet of 7.2 FIBERGLASS PRODUCTION air across the openings at the base of the bushing, which pulls individual fibers of The production of glass fibers starts with the approximately2040 cm (8-15 in) long from the dry mixing of silica sand and limestone, boric molten glass that is extruding from each openacid and a number of other products such as ing. These filaments are collected on a rotating clay, coal and fluorspar. These materials are vacuum drum, sprayed with size and gathered melted in a high-refractory furnace, the temperinto a strand. This package of filaments is again ature of the melt being dependent on the glass conditioned or dried prior to processing into a composition, but is generally about 1260°C specific product for further use. Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

* Significant portions of this article appeared in Handbook of Composites, 1982, (G. Lubin ed.) Van Nostrand Reinhold, New York.

132 Fiberglass reinforcement

,

,

RAW MATERIALS LIMESTONE SILICA SAND BORIC ACID ,CLAY ,COAL, F,LMRSPAR

I

U

I

FORYUUTION

I

I

W DECORATIVE AND INDUSTRIAL YARN

r

-1

TWISTING

INSPECTION AND WEIGHING

OVEN HEAT TREATING STRAND CHOPPING

W

7 7 7 -

PACKING

Fig. 7.1 Direct-melt fiberglass manufacturing process.

J

A capsule view of the fiber glass manufacturing process lor yarn. roving and chopped strand takes the fiber from raw rnalerflals batch slage lo 'inished products

Glass composifion 133

Fig. 7.2 Molten glass flows from tiny orifices in platinum bushings (Courtesy of PPG industries).

Each individual fiber is drawn from the bushing opening and must be controlled so that reproducible filaments, strand dimensions and properties are obtained. This control is achieved by the regulation of the melt viscosity, temperature and drawing speed. It is possible, therefore, to obtain a large number of filament diameters by varying the number of openings in the bushing and the drawing conditions. As demand has dictated over the years, the fiberglass industry has established a number of standard filament diameters (Table 7.1). 7.3 GLASS COMPOSITION

Glass is generally defined as an amorphous material, being neither solid or liquid. Chemically glass is made up of elements such as silicon, boron and phosphorus that are converted into glass when combined with oxygen, sulfur, tellurium and selenium. The molecular

Fig. 7.3 Molten glass flows from tiny orifices in platinum bushings (Courtesy of Owens-Corning Fiberglas Industries).

Table 7.1 Fiberglass filament designations

Filament designation B C D DE E G H K

Filament diameter i n ~ l O - ~ 1.5 1.8 2.1 2.5 2.9 3.6 4.2 5.1

Pm 3.8 4.5 5 6 7 9 10 13

arrangement is conducive to formation of an intricate three-dimensional network of oxygen tetrahedra with a silicon atom in the middle, bonded to each oxygen atom. Silicon by itself requires extremely high temperatures for liquefaction. Therefore, other elements are added to the mix to reduce temperatures and to produce

134 Fiberglass reinforcement a viscosity in the molten glass that will allow easy drawing. A number of glass compositions are available depending on the properties desired from the resulting fibers (Table 7.2). A-glass: A high alkali or soda glass is made into fibers for use in applications where good chemical resistance is needed. E-glass: A low alkali glass, based on aluminum borosilicate. This glass possesses excellent electrical insulation properties and is the premium fiber used in the majority of textile fiberglass production. C-glass: A material based on soda borosilicate that produces a fiber that offers excellent chemical resistance. S-2 glass: This glass is made up of magnesium, aluminum silicate and offers higher physical strength. Fibers produced from -~ perthis glass have an approximate cent strength improvement Over those of E-glass composition. D-glass: This fiber made from a low dielectric composition has dielectric loss properties 'Onstant Of 3.8 at mc s-l) superior to that of E-glass (6.O at 1mc s-l). R-glass: A 'Pecial glass that produces fiber that is alkali resistant and is used in reinforcing concrete. Low K: An experimental fiber produced to

0

0

improve dielectric loss properties in electrical applications (similar in performance to D-glass). Hollow fiber: A special glass whose fibers are tube-like or hollow; the material has specific applications in reinforced aircraft parts where weight could be significant. Te glass: A Japanese manufactured S-glass, for higher strength structural application.

7.4 FIBERGLASS PROPERTIES

The composition of the original glass melt probably plays the biggest role in determining the properties of the fiberglass. The continuing widespread use of fiberglass in numerous and diverse applications can be directly related to its inherent unique properties (Table 7.3). The suppliers of these materials in the USA are shown in Table 7.4. 0

High tensile strength: Fiberglass has an exceptionally high tensile strength cornpared with other textile fibers. Its strength to weight ratio exceeds steel wire in applications.

0

Heat and fire resistance: Because fiberglass is inorganic it does not burn or support combustion. Chemical resistance: Fiberglass has excellent resistance to most chemicals and is impervious to fungal, bacterial or insect attack.

0

Table 7.2 Fiberglass compositions (wt.%)

Components

Grade of glass

___

~____._

Silicon oxide Aluminum oxide Ferrous oxide Calcium oxide Magnesium oxide Sodium oxide Potassium oxide Boron oxide Barium oxide Miscellaneous

A (high alkali)

C (chemical)

E (electrical)

72.0 0.6

64.6 4.1

54.3 15.2

-

-

-

10.0 2.5 14.2

13.2 3.3 7.7 1.7 4.7 0.9

17.2 4.7 0.6

0.7

8.0

s (high strength) 64.2 24.8 0.21 0.01 10.27 0.27 0.01 0.2

Fiberglass properties 135 Table 7.3 Properties of fiberglass

Grade of glass -

A

C

E

5

-

Physical properties Specific gravity Mohs hardness Mechanical properties Tensile strength, psi x lo3 (MPa) At 72°F (22°C) At 700°F (371°C) At 1000°F (538°C) Tensile modulus of elasticity at 72°F (22"C),psi X lo6 (GPa) Yield elongation, % Elastic recovery, % Thermal properties Coefficient of thermal linear expansion, OF-' x lo4 ("C-1) Coefficient of thermal conductivity, Btu in h-' f f 2 OF-' (Wm-' K-I 1 Specific heat at 72°F (22°C) Softening point, "F ("C) Electrical properties Dielectric strength, V/mil Dielectric constant at 72°F (22°C) At 69 Hz At lo6 Hz

2.50 -

2.49 6.5

2.54 6.5

2.48 6.5

440 (3033)

440 (3033)

500 (3448) 380 (2620) 250 (1724)

665 (4585) 545 (3758) 350 (2413)

-

-

-

-

10.0 (69.0) 4.8 100

10.5 (72.4) 4.8 100

12.4 (85.5) 5.7 100

4.8 (8.6)

4.0 (7.2)

2.8 (5.0)

3.1 (5.6)

-

-

-

0.176

1340 (727)

0.212 1380 (749)

72 (10.4) 0.197 1545 (841)

-

-

498

-

-

-

-

6.9

7.0

5.9-6.4 6.3

5.0-5.4 5.1

Dissipation (power) factor at 72°F (22°C) At 60 Hz At lo6 Hz

-

-

0.005 0.002

0.003 0.003

Volume resistivity at 72°F (22°C) and 500 V DC, ohm-cm

-

-

1015

10'6

Surface resistivity at 72°F (22°C) and 500 V DC, ohm-cm

-

-

1013

1014

Optical properties Index of refraction

-

-

1.547

1.523

Acoustical properties Velocity of sound, ft/s (m/s)

-

-

17 500 (5330)

19 200 (5850)

0

Moisture resistance: Because fiberglass does not absorb water, it neither swells, stretches nor disintegrates. Fiberglass does not readily rot and continues to maintain its mechanical strength in humid environments.

0

Thermal properties: Due to its low coefficient of thermal linear expansion and high coefficient of thermal conductivity, fiberglass exhibits excellent performance in thermal environments.

136 Fiberglass reinforcement Table 7.4 Summary of available fibers from manufacturers S uppl ier

Type of glass --

0

OCF

E S 5-2 D R A C Low K Te

X

Hollow fiber

X

PPG

Nitto Boeski

Vetrotex

NEG

X

X

X

X

___

X

__

--

X

X

X

X

X

X X

X X

X

Electrical properties: Fiberglass being nonconductive is an ideal choice for electrical insulation, where designers can make use of the high dielectric strength and low dielectric loss properties.

7.5 FIBERGLASS TYPES

7.5.1 FIBERGLASS ROVING

Fiberglass roving is a collection of parallel continuous strands or filaments. Conventional The user may take advantage of one or more rovings are manufactured by winding together of the above properties in manufacturing rein- the number of single strands necessary to proforced composites. By utilizing both the high duce the required yield (the number of yards of strength and the excellent electrical properties roving weighing one pound). Single strand of glass fiber, the aircraft industry has found roving, as the name implies, consist of a single fiberglass an excellent reinforcement for strand of fiberglass filaments. These filaments are drawn from a bushing that has the correct radome applications. The printed circuit board industry has used number of openings, so that the single strand the combination of electrical properties and will have the correct yield. Rovings are generally manufactured from superior dimensional stability of fiberglass to manufacture circuit boards that can be used G or K filaments (see Table 7.1), larger diameunder the various adverse environmental con- ters are also available. Roving yields can vary from 1800 to 225 yd/lb (276 to 222 Tex). The ditions. In numerous applications glass reinforce- use of the word Tex indicates a unit of fiber ment is chosen because it allows composites to fineness that is assessed by the weight in retain maximum properties in high moisture grams of 1000 m of yarn, the lower the numenvironments. E-glass fibers are the material ber, the finer the yarn. of choice because of their excellent water resistance. (E-glass fibers exhibit only a 1-7% 7.5.2 WOVEN ROVING weight loss when exposed to boiling water for 60 min.) This water resistance helps maintain A number of rovings are woven into heavy, the physical and electrical characteristics of coarse weave fabrics for use in materials that the composite over prolonged time exposure require a quick build-up of composite thickness over relatively large areas, such as various to aqueous atmospheres’. marine products and different types of tooling. Woven rovings are manufactured in weights ranging from 400-1400 g/m2 (1240

Fiberglass types 137 oz/yd2) and thicknesses of from 0.7-1.5 mm wet layup techniques or by impregnating and (0.02-0.04 in) (see Table 7.5). applying low pressure (RTM) (see Table 7.5). The fiberglass rovings can be used in conjunction with polyester resins in hand lay-up 7.5.3 FIBERGLASS MAT ~. techniques that typically be used in the The three basic forms of fiberglass mat, are manufacture of boats. Woven roving reinforced chopped strand, continuous and surfacing or laminates can be made by using conventional veil. Table 7.5 Woven roving and polyester laminate data

Woven roving constructions and properties Count, per in (per em)

Weight, oz/yd2 (g / m2)

Thickness, in (mm)

Weave

5x8 (2 X 3.2)

18.0 (610)

0.031 (0.787)

Plain

5x8 (2 X 3.2)

24.0 (814)

0.038 (0.965)

Plain

5x6 (2 X 2.4)

30.0 (1020)

0.049 (1.24)

Plain

5x8 (2 x 3.2)

36.0 (1220)

0.052 (1.32)

Plain

Laminate properties - 24 oz (814 g ) woven roving reinforced Resin type"

FR Epoxy

FR Epoxy

GI' Epoxy

Resin content, wt.%

49.0

45.6

42.5

Thickness, in (mm)

0.299 (7.59)

0.265 (6.73)

0.263 (6.68)

Cureb

RT

P+T

P+T

Flexural strength, psi (MPa) Cond. A' Cond. D2/100d

34 200 (235.8) 38 800 (267.5)

49 800 (343.4) 44 400 (306.1)

73 100 (504.0) 59 600 (410.9)

Flexural modulus, psi (GPa) Cond. A Cond. D2/100

20.6 X lo6 (14.2) 2.12 X lo6 (14.6)

2.41 X lo6 (16.6) 2.31 X lo6 (15.9)

2.92 X 106(20.0) 2.83 X lo6 (19.5)

Compressive strength, psi (MPa) Cond. A Cond. D2/100

29 700 (204.8) 27 300 (188.2)

25 700 (177.2)

25 200 (173.8)

47 300 (326.1) 40 100 (276.5)

Tensile strength, psi (MPa) Cond. A Cond. D2/100

42 000 (289.6) 40 600 (279.9)

48 900 (337.2) 47 400 (326.8)

51 400 (354.4) 50 400 (347.5)

"FR = Fire Retardant; GP = General Purpose "RT = Room Temperature; P + T = Pressure and Elevated Temperature 'Cond. A = as-received condition dCond. D2/100 = immersion for 2 h in water at 100°C.

138 Fiberglass reinforcement Chopped strand mat is a non-woven material which the fiberglass strands are chopped into 2-5 cm (1-2 in) lengths and evenly distributed at random onto a horizontal plane, bound together with some type of chemical size. The mats so produced weigh from 2.6-12g/m2 (0.7-3.0 oz/ft2)and are available in a variety of widths, 60-230 cm (24-90 in) with 9.6 cm (38 in) width being the most typical. Continuous strand mat is made of unchopped continuous strands of fiberglass deposited and interlocked in spiral fashion. This mat is open and flexible, but due to its mechanical interlocking does not require much application of size to achieve adequate handling strength. Surface mat, or veil as it is sometimes called, is a very thin mat of single continuous filaments used frequently as a decorative surface reinforcing layer in hand lay-up or press molding processes in order to minimize surface defects and to prevent ’weep’ in pressurized wound tanks. 7.5.4 TEXTILE FIBERGLASS YARN

A yarn can be described as an assembly of fibers or strands which can be woven into some form of textile material. The continuous individual strand as it emerges from the bushing opening represents the simplest form of textile fiberglass yarn and is designated ’singles’ yarn. However, to ensure the correct and efficient utilization of this yarn in a weaving operation, additional strand integrity has to be added by twisting the yarn slightly (less than one turn per inch). A number of woven fabrics, however, require heavier yarns than can be drawn from the bushing. These heavier yarns can be achieved by combining single strands using a twisting and plying operation. This simply involves twisting one or more single strands together or subsequently plying together two or more of the twisted strands. A yarn can be defined as having an ‘S’ twist if when it is held in a vertical position,

the spirals conform in shape to the central position of the letter ’S’. Alternatively, a yarn is said to have a ’Z’ twist if the spiral conforms in shape to the central position of the letter ‘Z’. Strands that have a simple twist (greater than one turn per inch) will kink, corkscrew and unravel, because their twist is in only one direction. The plying operation will normally eliminate this problem by countering the twist in a twisted ’singles’ yarn with an opposite twist in the plied yarn. For instance, a ’singles’ yarns which have a ’Z’ twist are plied with an ’S’ twist that will result in a balanced yarn. The twisting and plying operations allow the yarn‘s strength, diameter and flexibility to be varied, which in turn allows scope in producing a wide range of suitable woven fabrics. 7.5.5 TEXTURED YARN

’Singles’ or plied yarn can be textured by using a jet of air directed onto the yarn’s surface which results in random but controlled breakage of surface filaments, producing a general increased loft to the fiber surface. The term used to describe such a process is texturizing or bulking and the degree to which it occurs can be controlled by regulating the air pressure and yarn feed rate. The texturing process opens the fiber bundles resulting in some mechanical damage to the surface filaments. The increased surface area allows higher resin absorption during impregnation and produces low glass-to-resin ratios, resulting in more economical laminates. 7.5.6 YARN NOMENCLATURE

Because of the wide variety of fiberglass yarns produced, it is necessary to have a precise system for yarn identification. The standard fiberglass yarn nomenclature is based on both alphabetical and numerical designations. The first alphabetical letter identifies the glass composition (E-glass), the second letter indicates filament type (C = continuous) and

Fiberglass types 139 the third and fourth letters identify filament ECG 150 4/2 3.8s diameter (E 7 micron) [for details see Table 7.11. E = E-glass; C = Continuous filament; G = The first series of numbers in the numerical Filament diameter designation represents 1/1OOth of the basic 150 4/2 = Four basic strands of 150 1/0 are strand yield, the second number series specitwisted together to form 150 4/04 fies the number of single strands twisted 150 4/2 3.8s = Plying two strands of EGG together and the number of twisted yarns 150 4/04 (using 'S' twist to create balance) plied together. The total number of basic The above yarn contains 8 (4 x 2) basic 150 strands in a plied yarn is found by multiplying strands with a glass yield of 1875 (15 000 + 8) these two figures. The yield of the yarns is yards/pound. obtained by dividing the basic strand yield by The nomenclature and typical properties of the total number of strands in the yarn. A third commercially available weaving yarns can be number combined with either the letter 'S' or seen in Table 7.6. 'Z' (designating the type of twist) will sometimes also be included. For example:

Table 7.6 Commercially available fiberglass weaving yarns

Yarn designation

Yield

Glass system

-~

E-glass Yarns ECD 1800 1/0 ECD 1800 1/2

Tex system

ydhb ~~

Minim um breaking strength ___

-

~.

Tex

Ib

N

2.75 5.5

0.25 0.5

1.11 2.22

-

EC5 2.75 1 x 0 EC5 2.75 1 X 2

180 000 90 000

ECD 900 1/0 ECD 900 1/2

EC5 5.5 1 x 0 EC5 5.5 1 x 2

90 000 45 000

5.5 11

0.5 1.1

1.11 4.89

ECD 450 1/0 ECD 450 1/2 ECD 450 1/3 ECD 450 2/2 ECD 450 3/2

EC5 11 1 x 0 EC5 11 1 X2 EC5111x3 EC5112x2 EC5 11 3 X 2

45 000 22 500 15 000 11 250 7500

11 22 33 44 66

1.1 2.2 3.3 4.4 6.6

4.89 9.79 14.7 19.6 29.4

ECD 225 1/0 ECD 225 1/2 ECD 225 1 / 3 ECD 225 2/2 ECD 225 2/3

EC5 22 1 x 0 EC5 22 1 x 2 EC5 22 1 X 3 EC5 22 2 x 2 EC5 22 2 X 3

22 500 11250 7500 5625 3750

22 44 66 88 132

2.2 4.4 6.6 8.8 14.4

9.78 19.6 29.4 39.1 64.1

ECDE 150 1/0 ECDE 150 1/2

EC6 33 1 X 0 EC6 33 1X 2

15 000 7500

33 66

3.5 7.0

15.6 31.1

ECG 150 1/0 ECG 150 1/2 ECG 150 1/3 ECG 150 2/2 ECG 150 2/3 ECG 150 2/4 ECG 150 3/3 ECG 150 4/4

EC9 33 1x 0 EC933 1 X2 EC9 33 1 x 3 EC9 33 2 X 2 EC9 33 2 x 3 EC9 33 2 x 4 EC9 33 3 x 3 EC9 33 4 X 4

15 000 7500 5000 3750 2500 1875 1667 938

33 66 99 132 196 264 297 528

3.5 7.0 9.0 12.0 18.0 24.0 27.0 48.0

15.6 31.1 40.0 53.4 80.1 107 120 214

Continued on next page

140 Fiberglass reinforcement Table 7.6 Continued

Yarn designation Glass system ECDE 75 1 / 0 ECG 75 1/0 ECG 75 1 / 2 ECG 75 1 / 3 ECG 75 2/2 ECG 75 2/3 ECG 75 2/4 ECH 55 1/0 ECDE 37 1/0 ECG 37 1/0 ECG 37 1/2 ECG 37 1 / 3 ECH 25 1/0 ECK 18 1 / 0

Yield

Minimum breaking strength

Tex system 7500 7500 3750 2500 1875 1250 938 5500 3700 3700 1850 1230 2500 1800

66 66 132 198 264 396 528 90 134 134 268 403 198 275

5.7 5.7 11.4 17.1 22.8 34.2 45.6 9.5 11.2 11.2 22.8 34.2 17.0 23.0

25.4 25.4 50.7 76.1 101 152 203 42.3 49.8 49.8 101 152 75.6 102

s c 5 11 1 x 1 s c 9 33 1 x 2 s c 9 33 2 x 2

22 500 7500 3750

22 66 132

-

-

8.2 16.4

36.5 72.9

ET6 33 1 X 0 ET6 66 1 X 0 ET6 134 1 X 0

14 200 7100 3450

35 70 144

EC6 66 1 X 0 EC9 66 1 X 0 EC9 66 1 X 2 EC9 66 1 X 3 EC9 66 2 X 2 EC9 66 2 X 3 EC9 66 2 X 4 EClO 90 1 X 0 EC6 134 1 X 0 EC9 134 1X 0 EC9 134 1 X 2 EC9 134 1 X 3 EClO 198 1 X 0 EC13 275 1 X 0

S-glass Yarns SCD 450 1/2 SCG 150 1/2 SCG 150 2/2 Textured Yarns ETDE 150 1/0 ETDE 75 1/0 ETDE 37 1/0

7.5.7 FIBERGLASS FABRIC

The effect that fiberglass fabrics have on the properties of composite material is largely dependent on the fabric’s construction, which involves fabric count, warp and filling yarn construction, weight and weave pattern. The fabric’s count is made up of the number of warp yarns (’ends’) per inch (cm) in the width and the number of filling yarns (’picks’) per inch (cm) in the lengthwise direction. The warp yarn is the yarn that lies in the lengthwise (machme) direction of the fabric. The filling yarn is the yarn lying in the crosswise direction of the fabric. The physical parameters of the fabric, such as weight, thickness and tensile strength are directly proportional to the types and numbers of yarns used to weave it (see Table 7.7)

1.24 2.2 4.7

5.52 9.79 29.0

There is a variety of weave patterns (Fig. 7.4) that can be used to combine warp and filling yarns to form a fabric. This weave pattern controls the handling characteristics of the fabric and to a large extent the properties of the composite that uses it as a reinforcement. (a) Plain weave: This fabric is constructed so that one warp yarn passes over and under one filling yarn (and vice versa). While the resulting fabric exhibits the greatest degree of stability in respect to yarn slippage and fabric distortion, this stability is also a function of the fabric count and yarn content. (b) Basket weave: A basket weave fabric has two or more warp yarns interlocking over and under two or more filling yarns. While a basket weave is less stable than a plain weave, it is more pliable and will conform or drape more readily to simple contours.

Fiberglass types 141 Table 7.7 Glass-fabricconstruction and properties: physical properties for loomstate fabrics (without finish) Style

104 106

Count, /cm /in

23.6 x 60 x 22.0 x 56 x

20.5 52 22.0 56

236x138 60x35

Warp yarn, Tex system US system

Filling yarn, Tex system US system

Weave

5 5.5 1x 0

5 2.75 1 x 0

Plain

D 900 1/0 D 1800 1/0 D 5.5 1x 0 5 5.5 1x 0 D 900 1 / 0 D 900 1 / 0 D5.51~2 55.51~0 D9001/2 D9001/0 D5.51~2 55.51~2 D9001/2 D9001/2 5111x2 5111x2

Weight,

g/m2 ozf yd?

Plain

Plain

Thickness, Bveaking mm stvengtlz, in N/5 cm lb/in

19.7 0.58 24.7 0.73

0.030 0.0012 0.038 n.nni - _ -,5-

35.6

350 x 130 40x15 394 x 350 4.5 . x 40 .-

D4501/2

D4501/2

60x57

5111x2 DW1/2 D111x2 D 450 1/2

120

23.6 x 22.8

5 111x 2

5531x2 D9001/2 511 1 x 2 D 450 1/2 5 1 11x 2

Crowfoot

107

0.043 613x 175 0.0017 Mx20 0.051 613x350 0.0020 70x40 0.076 718x700 0.003 82x80 0.076 1077 x 525 0.003 123x60 0.102 1095x1050 0.004 125x120 0.102 1095 x 1050

220

23.6 x 22.8 60 x 58

7 22 1x 0

7 22 1x 0

Crowfoot

E 225 1 / 0

E 225 1 / 0

109.2 3.22

0.0889 0.0035

1096 x 1051 125 x 120

118~19.3 30x49.

5111x2

5223x2 D2253/2

Crowfoot

298 8.78

0.229 0.009

438x5250

D4501/2

23.6 x 13.8 60 x 35 23.6 x 18.5 60 x 47

5111x0 D 450 1/0 5111x0

55.51~0 D 900 1/0 5111x0

Plain

35.6 1.05 48.8 1.44

0.043 0.0017 0.051 0 002

613 x 175 70x20 613 x 350 70 x 40

1Q7

108

1 1 116

341 1070

1080

23.6~18.5 60x47 151Fx95.4 6x39 23.6x25.2 60x64

I

236~22.8

1.05 Plain Plain Plain

Plain

Plain

4.8.8 1.44 71.9 2.12 83.7 2.47 107 3.16

1

50x600

700x613 80x70 0.089 788x1 0.0035 9 D x 1 0.109 1310 x 1180

0.060

0.0024

1 1165

23.6 x 20.5 60 x 52 Continued on next page

5111x2

D 450 1/2

9 331 x 0 G 150 1/0

Plain

124 3.66

0.0043

150 x 135

142 Fiberglass reinforcement Table 7.7 Continued

Count, /cm /in



7511x0 E 110 l / O 751 1/0 E 110 1/0 6331x0

Plain

DE 150 1/0

7511x0 E 110 1/0 751 1 x 0 E 110 1/0 6331x0 DE 150 1/0

1522

1 9 . 3 ~16.5 49x42 18.2 x 12.7 46x45 23.6x19.7 60x50 9.1x 8.7

1523

28 x 20 -_ -

9331 x 2 CG 150 1/2 9333x2 G 150 3 / 2

Plain

24 x 22 11.0 x 7.9

9331x2 G 150 1 / 2 9333x2 G 150 ...3 .,/ 2

13.4 x 12.6 34 x 32

9331x2

9331x2 G 150 1/2

Plain

G 150 1/2

9333x3 G 150 3/ 3 9331x2 G 150 1/2 7221x2 D 225 1/2 7221x0 E 225 1/0 9334x2 G 150 4/2 9331x2 G 150 1 / 2 9331x3 G 150 1/3 9332x2 G 150 2/2 9684x2 G 75 2/2 9334x4 G 150 4/4 968 1x 0 G 75 1 / 0 9331x0 G 150 1/0 6681x0 DE 75 1/0 9331x0 G 150 1/0 6331x0 DE 150 1/0

Plain

Style

Weave

g/m2 oz/yd2 -. _. -

1560 1501 1504

1526 1527 1528 1543 1557 1564 1581 1582 1583 1584 1588 1614 1652 1665 1674 1675

1677

6.7 x 6.7 17 x 17 17.3 x 12.6 44 x 32 19.3 x 11.8 49 x 30 22.4 x 11.8 57 x 30 9241x2 G 37 1/2 22.4 x 21.3 57 x 54 23.6 x 22.0 60 x 56 21.3 x 18.9 54 x 48 17.3 x 13.8 44 x 36 16.5 x 14.2 42 x 36 11.8 x 5.51 30 x 14 20.5 x 20.5 52 x 52 15.7 x 9.5 40 x 24 15.7 x 12.6 40 x 32 15.7 x 12.6 40 x 32

15.7 x 15.7 40 x 40 Continued on next page

~

9333x3

G 150 3/ 3 9331x2 G 150 1/2 9662x2 G 75 2/2 9331x2 G 150 1 / 2 9241x2 G 37 1/2 9331x2 G 150 1/2 9331x3 G 150 1/3 9332x2 G 150 2/2 9684x2 G 75 2/2 9334x4 G 150 4/4 933 1x 0 G 150 1 / 0 9331x0 G 150 1/0 9331x0 G 150 1/0 9331x0 G 150 1/0 6331x0 DE 150 1 / 0

6331x0

6331x0

DE 150 1/0

DE 150 1/0

Weight,

Rain

Plain

Plain

Plain Crowfoot Crowfoot Plain Satin Satin Satin Satin Satin Len0 Plain Plain Plain Plain

Plain

-

Thickness, Breaking mm strength, in N/5 em lb/in

_ -

.

-

0.132

2285 x 515

0.0052 0.127

o m

261x253 2145 x 2102 245x240

0.127

2 6 8 8 ~2233

167.8 4.95 166.1 4.90 147.8

4.36

0.01)50

125 3.70 403 11.9

0.140 0.0055 0.356 0.014

1400 x 160 x 4595 x 525 x

1970 x 1705 225x195

307x255 1180 135 3500 400

185

0.165

5.45

0.0065

437 12.9 203 6.00 298 8.34 184 5.42 431 12.7 302 8.92 454 13.4 570 16.8 861 25.4 1756 51.8 79.0 2.33 141.7 4.18 141.7 3.50 95.6 2.82 95.6 2.82

0.381 4375 x 4245 0.015 500 x 485 0.178 2185 x 1750 0.007 250 x 200 0.229 5250 x 525 0.009 600x60 0.140 3240 x 525 0.0055 370 x 60 0.406 4375 x 3940 0.016 500 x 450 0.216 2975 x 2885 0.0085 340 x 330 0.356 4595 x 4375 0.014 525 x 500 0.457 5690 x 5165 0.018 650 x 590 0.686 8315 x 7005 0.035 950 x 800 1.27 16 635 x 11820 0.050 1900 x 1350 0.127 657x744 0.005 75 x 85 0.114 2154 x 1926 0.00454 246 x 220 0.114 2154 x 1926 0.0049 180 x 180 0.107 1225 x 832 0.0042 140 x 95 0.107 1225 x 832 0.0042 140 x 95

108 3.20

0.114 0.0045

1225 x 1135 140 x 130

Fiberglass types 143 Table 7.7 Continued . . .

Style

Count, /cni


Weave

/in

Warp yarn, 7ex system US system

Weight, g/m’ oz/yd2

1678

15.7 x 15.7

9331x0

9331x0

Plain

108.5

16 x 14

2112

15.7 x 15.4 40x39 23.6 x 22.0 60x56

K 18 1 / 0 7221x0

R 18 1 / 0 7221x0

Plain

71.2

E2251/0

E2251/0

7221x0

5111x0

E225L’O

E4501/0

7221x0 F 225 1/11

c 225 I / i )

_ _ _ ~ _ _ ~ _ ~ ~

-~ ~~~

2il3 2116

23.6 x 22.8 60 x 58

7221x0

2117

26~21.6

7221x0

E2251/0

7221x0 R2251/0

23.20

66x55 23.6~228

5221x0

5221

D2251Jf) 0 2 2 5

2123

60x58 15.7x15.4 80x39 23.6x20.5 60 x 52

D225 l / O

2165 2313 2523 2532

3070 3313

13.6 x 25.2 60 x 64 11.0 x 7.9 28 x 20 6.3 x 5.5 16 x 14

276x76 70 x 70 2.3.6 x 24.4 60x62

5221x0 5221x0 D 225 1/0 7221 x 0 E 225 1 / 0 10 198 1x 0 H 25 1 / 0 10 198 1x 0 H251/0 6 16.5 1 x 0 ~

DE 300 1/0 6 16.5 1 x 0

DE 300 1/0 34 1

371 3733

Thickness Breaking mrn strength, in N/5 cm lb/in

7 1 x7.1

18 x 18 3743 19.3 x 11.8 49 x 30 3783 21.2 x 18.9 54 x 48 4522 9.4 x 8.7 (6522) 24x22 4533 7.1 x 7.1 (6533) 1 8 x 18 Continued on next page

91341x0 G 37 1/0 9 134 1x 0 G 37 1/0 9 134 1x 0 G 37 1/0 9331x2 G 150 1/2 9681x2 G 75 1/2

9.66

Plain Plain

Crowfoot

1839 x 1751

0.014

431 x 41_ o.

0.086

788x700

2.10

0.0034

90x80

80.7 2.38

0.081 0.0032

1225X!i2B 14Ox~

0.102 u.001

1095 x 1050 125 x 120

0.085 0.0037

1182xlfM9 135x115

002 0.004

1095XlW W x 120. rnXlP35

107 1.16

Plain

0.109

20.4 3.18 107

3.16

888

9331x0 G 150 1/0

Plain

9331x0

Plain

225

Plain

80.7 2.38 403 11.9 246 7.25 ~.

0.81 3 0.0032 0.330 0.013 0.254

92.1 2.74 81.4 2.40

0.0788 0.0031

2.62

G 150 1/0 511 1 x 0 D 450 1 / 0 10 198 1x 0 H 25 1/0 10 198 1x 0 H251/0

6 165 1 x 0 DE 300 1/0 6 16.5 1x 0 DE 300 11’0 91341

3.70

Plain Plain

Plain

Plain

crowfoot

G 37 1 91341x0 G 37 1/0 5 22 1x 0 D 225 1/0 9 134 1x 0 G 37 1/0 9331x2 G 150 1/2 9681x2 G 75 1/2

431 12.7

Plain

Crowfoot

8Harness Plain Plain

197 5.80 286 8.45 556 16.4 123.4 3.64 200.7 5.92

0.094

0.0051 0.114 0.0045

n.nin 0.0076 0.0030

9ox13Q 1095x1220 125x140 1226 x 525 140 x 60 5075 x 3370 580 x 385 2625 x 2450 3nn x 2x13

1226xk26 lLPoxlg0 1050x2094 120x125

0.390 5900x3700 0.0134 571x423 0.203 0.008 0.203 0.008 0.406 0.0160 0.130 0.0051 0.188 0.0071

2185 x 1750 250 x 200 5250 x 525 600 x 60 4816 x 4290 550 x 490 1226 x 1130 140 x 129 2382 x 2601 272 x 297

144 Fiberglass reinforcement Table 7.7 Continued ~~

Style

Count, /cm /in -

4700

6060

6581 6781 7500 7520 7532 7533 7544

7626 7628 7629

76281

-

5.5 x 5.1 14x 13 23.6x23.6 60x60 22.4x21.3 57x54 22.4 x 21.3 57 x 54 6.3 x5.5 16 x 14 7.1 x 7.1 18 x 18 6.3 x 5.5 16 x 14 7.1 x 7.1 18 x 18 11.0 x 5.5

~


Filling yarn,

9681x0

9681x0

G 37 1/0

G 37 1/0

Weave

Tex system US system -

-

6 8 . 2 7 1 ~ 0 68271x0 l b 0 DE 600 5 / 0 9331x2 9331x2 f;1 9 1/2 G 150 112 9681x0 9681x0 S2C6 75 1 / 0 S2CG75 1 / 0 9682x2 9682x2 G 75 2/2 G 75 2/2 9681x3 9681x3 G 75 1/3 G 75 1/3 9681x3 9681x3 G 75 1 / 3 G 75 1 / 3 9681x2 9681x2 G 75 1/2 G 75 1 / 2 9682x2 9682x2

-

-

-

Rain

147.5

Plain

DH

13.4 x 12.6 34 x 32 17.3 x 12.6 44 x 32 17.3 x 13.3 44 x 34

9681x0 G 75 1/0 968 1 x 0 G 75 1/0 9681x0 G 75 1 / 0

9681x0 G 75 1/0 968 1 x 0 G 75 1/0 9681x0 G 75 1 / 0

44 x 20

G 75 1/0

G 37 1/0 (Tex)

59 x 54 17.3 x 12.6 44 x 32

DE 75 1/0 968 1 x 0 G 75 1/0

DE 75 1 / 0 9 6 8 1x 0 G 75 1/0

Weight, g/m2 oz/ydZ

8ws X HS

Plain Plain Plain Plain 2xlBSK.

Plain Plain Plain

4.30 39.0 1.19

Shaded items may be obsolete or not in all vendor current catalogues

0.1%

2189x1445

0.0077 250x165 0.0019 656.7 x 718 0.048 75x82

301.8 8.90 301.X 8.90 327 9.66 294.3 8.68 254 7.50 193 5.70 610

0.228 0.28 0.0090 0.356 0.014 0.304 0.0120 0.254 0.010 0.203 0.008 0.559

3520x2846 400x325 3320 x 28-36 400 x 325 3940 x 3590 450 x 410 2890 x 2890 338 x 330 2930 x 2765 335 x 316 2010 x 1925 230 x 220 6520 x 7265

183 5.40 203 6.00 211.9 6.25

0.168 0.0066 0.178 0.007 0.178 0.007

1970 x 1750 225 x 200 2190 x 1750 250 x 200 2190 x 1883 250 x 215

0.011

250 x 120

6.85

Crowfoot

Thickness, Breaking mm strength, in a N/5 em 1b/in

9.15 203 6.00

0.90

0.010 0.173 0.0068

340 x 330 2190 x 1750 250 x 200

Fiberglass types 145 Notes to Table 7.7

a

All fabrics listed in Table 7.7 which contain D 225 5 22 yam can also be woven with the original E 2257 22 yarn equivalents. Breaking strength of fabrics in SI units: N was determined by testing a specimen of 5 cm width and is for loomstate fabrics. Finishing reduces fabric strength. This thickness is an estimate for dry ply thickness and is an average of several tests. The manufacturer usually reports a range of thicknesses. To estimate the cured ply thickness, divide the dry estimate by the expected fiber volume, i.e. if the dry thickness is 0.1 mm and the fiber volume is 0.5 then the estimated cured ply thickness is 0.1/0.5 or 0.20 mm. See introduction for methods of converting between weight and volume percent. Styles 4522 (6522) and 4533 (6533) are S-2 glass; Styles 6581 and 6781 are S-glass. Data from Reference 3

Fig. 7.4 (a) Plain weave; (b)basket weave; (c) twill weave; (d) crowfoot satin (four harness weave; (e) eightharness satin weave.

(c) Twill weave: This fabric consists of one or more warp yarns over and under two or more filling yarns in a regular pattern. This produces either a straight or a broken diagonal line in the fabric, which produces greater drapeability and stability. (d) Crowfoot satin weave: This weave pattern has one warp yarn interlocking over three and under one filling yarn in an irregular pattern. The fabric that results is extremely pliable and lends itself to conforming to complex contours. (e) Eight-harness satin weave: A fabric constructed with one warp yarn interlocking

over seven and under one filling yam in an irregular pattern. The resulting fabric is very pliable and readily conforms to compound contours. Because this weave pattern allows comparatively high fabric counts, it contributes maximum strength to composites reinforced by it. (f) Unidirectional fabrics: Fabrics produced with heavy warp yarns and light filling yarns, in either crowfoot or long shaft satin weaves. The filling yarns can also be composed of yarns other than glass. These fabrics offer high strength reinforcements in the heavy yarn direction.

146 Fiberglass reinforcement

(g) Non-woven: Unidirectional fabrics that can be produced by sticking the 'warp' and 'filling' yarns together chemically. Although this chemical bonding contributes to the fabric's stability, these fabrics have a firm hand and do not drape over complex contours. 7.5.8 OTHER WOVEN FORMS

Fiberglass yarns can also be woven into tapes, contoured fabrics, fluted core fabrics and three dimensional fabrics. Tapes are usually narrow fabrics [less than 30 cm (12 in) wide], with a secured selvedge to prevent unravelling. Contoured fabrics are woven into a specific geometrical shape. Fluted core fabrics are two parallel layers of fabric tied together by stringers of woven fabrics so that the cross-section is triangular or rectangular. Three dimensional fabrics are really planar or fabrics woven with yarns in three distinct directions within the fabric plane. That is, yams are interwoven in: (i) the machine direction, (ii) +45" from the machine direction, or (iii) -45" from the machine direction. 7.5.9 MILLED FIBERS

Continuous fiberglass strands can be hammer milled into very short fiber lengths [approx. 2-6 mm ( H 6 - X in) long]. The actual fiber lengths are determined by the diameter of the screen openings through which the fibers pass during milling. These milled fibers are usually used as inert fillers for thermoplastic and thermoset resin systems. 7.6 SURFACE TREATMENTS

7.6.1 ROVINGS

In order for fiberglass rovings to be compatible with processing methods and materials, a chemical size is applied during the basic fiber forming operation. The formulation of this size is designed to hold the individual glass fila-

ments together, lubricate the roving for contact with various processing equipment, and to permit the glass filaments to be thoroughly wetted when combined with other materials. Fiberglass roving sizes consist of polyvinyl acetate, polyvinyl alcohol, or PVA/starch as a film former with the addition of such chemicals as chrome complexes, organosilane antistatic agents and lubricants that impart the desired strand characteristics.The film former aids in adhering the filaments together and giving strand integrity that will reduce filament abrasion during fiber drawing, strand conversion and the end use of the rovings by the fabricator. An organosilane or coupling agent is added, that can react with hydroxyl groups on the glass fiber surface and also possesses one or more reactive groups to react with other materials, specifically those present in thermoplastic and thermosetting resins. The silane produces a form of chemical bridge between the glass surface and the resin matrix. The inclusion of antistatic agents and lubricants improves the softness and choppability of the roving. The selection of a roving size is based on its resin compatibility, processing and the expected end use performance. The largest percentage of rovings are used in processes that require relatively short fiber lengths (1-5 cm, 0.5-2.0 in). This type is used in processes such as spray-up, preform molding, bulk molding compounds and continuous lamination. Rovings suitable for chopping are available in varying degrees of softness; the softer rovings are generally more difficult to chop, but they are recommended for use in applications that require intricate shapes, sharp comers and difficult radii. Conversely, the harder types of rovings are generally recommended for use where the choppability is of prime importance and they are used in simple parts with minimum contours and radii. Single strand rovings that are commonly used in filament winding and pultrusion should have good strand integrity, controlled

Design considerations 147 level of broken filaments, good wettability and uniform processability under controlled tension. S-glass based rovings have been used in applications where the composite needs increased physical properties ( e g filament wound pressure vessels).

The continuous fiberglass strands that will be used in weaving are treated at the bushing with a starch-oil binder; the general formula for such a binder can be a partially or fully dextrinized starch or amylose hydrogenated vegetable oil, a cationic wetting agent, emulsifying agent and water. These sizes or binders are intended to protect the fibers from damage during their formation and subsequent operations of twisting, plying and weaving.

nated Y) and hydrolyzable groups (designated X) in materials with a generic structure X,SiRY. The hydrolyzable groups are intermediaries in the formation of silanol groups that bond to the glass surface, whereas the organofunctional groups are designed for reactivity or compatibility with the polymer to be used by the composite manufacturer. (Table 7.8) The chemical functionability of the coupling agent can determine the resistance to varied environments, chemical, physical and thermal (Table 7.9). There are many functions that may be attributed to coupling agents at the glass-resin interface. They may provide lubrication to protect against abrasion during fabrication. The coupling agent mechanism can protect against stress corrosion as a result of water incursion and can help improve the mechanical and electrical properties of reinforced composites.

7.6.3 FIBERGLASS FABRICS

7.7 DESIGN CONSIDERATIONS

Fabrics directly from the loom still have the original binders that were applied at the yarn bushing; the warp yarns have also been treated with a polyvinyl alcohol solution to help protect them during weaving. These binders and sizes are not compatible with the resins used by the composite manufacturers and must be removed prior to impregnation with other polymers. This is usually accomplished by treating the fabric to carefully controlled time-temperature cycles, which results in complete removal of all organic material. However, after this heat treatment no suitable interface exists between the fiberglass surface and the resin matrices that will eventually be used. An improved interface with resulting good adhesion even in adverse environment conditions can be achieved by the application of a coupling agent to the glass fiber surface. The most commonly used coupling agents in current use are those termed silanes. The coupling agent mechanism of these organofunctional silanes depends on a link between the organofunctional group (desig-

It is only recently that the engineering and design criteria in the combination of such different materials as thermosetting and thermoplastic polymers with reinforcing of glass were properly understood. There are a number of basic considerations that are common to the use of fiberglass as a reinforcing media in the production of composites. The type form, weave pattern, weight, permeability and alignment of the fibers have direct bearing on the mechanical properties of the composite. Also in some instances the chemical, thermal and electrical properties may also be influenced by the judicious choice of reinforcing material (Table 7.10).

7.6.2 FIBERGLASS TEXTILE YARNS

7.7.1 GLASS COMPOSITION

The chemical composition of the glass can have a direct bearing on the fiber properties which can indirectly affect composite performance. It has been noted that 'E' glass composition fibers can contribute 210 000-225 000 psi to the

148 Fiberglass reinforcement Table 7.8 Commercial coupling agents

Organofunctionalgroup

Clzemical structure

Vinyl Chloropropyl

CH,= CHSi(OCH,), ClCH,CH,CH,Si(OCH,),

EPOXY

CH2CHCH20CH2CH2CH2Si(OCH3)3

/”\

CH2 Methacrylate Primary amine Diamine Mercapto Cationic styryl

I

(CH2=C-COOCH2CH2CH2Si(OCH3)3 H,NCH,CH,CH,Si(OC,H,), H,NCH,CH,NHCH,CH,CH,Si(OCH,)3 HSCH,CH,CH,Si(OCH,),

CH,=CHC,H,CH,NHCH,CH~NH(CH~),Si(OCH,),HCl CH3

Cationic methacrylate

c1

I

I

+

CH2=CXOOCH2CH2- N(Me2)CH2CH2CH2Si(OCH& CH3

I

CH,=C

I

Chrome complex

CH2 Titanate Cross-linker Mixed silanes Formulated

I

(CH2=C-C00)3 TiOCH(CH& (CH,O), SiCH,CH,Si (OCH,), C,H,Si(OCH,), + F Melamine resin + C

tensile strength of a composite which they reinforce. Other glass compositions/ ’” and ’2’ glass, have produced even greater strength improvements in reinforced composites.

7.7.2 FILAMENT DIAMETER

A number of different ’E’ glass filament diameters have been evaluated comparing the influence of filament diameter on resulting laminate properties. The fabrics evaluated were woven with the same count and weave pattern as Style 1581, using both plied and singles yarn containing different filament diameters (Tables 7.11 and 7.12).

Design considerations 149 Table 7.9 Effect of coupling agents on the mechanical strength properties of laminates Finish

Loomstate

Volan

A1 72

A174,26030

Polyester laminates reinforced with style 1581 glass fabric Flexural strength, psi X lo3 (MPa) Cond. A 40.4 (278.6) 32.3 (222.7) 73.2 (504.7) Cond. D2/100 25.3 (174.4) 31.8 (219.3) 59.0 (406.8)

68.8 (474.4) 66.1 (455.8)

75.5 (520.6) 77.3 (533.0)

Compressive strength, psi X lo3 (MPa) Cond. A Cond. D2/100

112

23.8 (164.1) 12.5 (86.2)

40.4 (278.6) 12.1 (83.4)

48.0 (331.0) 38.9 (268.2)

47.1 (324.8) 44.9 (309.6)

66.2 (456.4) 58.8 (405.4)

48.1 (331.6)

20.6 (142.0)

49.9 (344.1) 48.8 (336.5)

50.0 (344.8) 51.2 (353.7)

52.2 (359.9) 50.7 (349.6)

Epoxy laminates reinforced with style 1581 glass fabric Flexural strength, psi X lo" (MPa) Cond. A 78.2 (539.2) 76.0 (524.0) 80.9 (557.8) Cond. D2/100 76.7 (528.8) 72.4 (499.2) 83.1 (573.0)

71.3 (491.6) 68.9 (475.1)

85.7 (590.9) 82.4 (568.1)

Compressive strength, psi X lo3 (MPa) Cond. A Cond. D2/100

64.3 (443.3) 58.5 (403.3)

63.4 (437.1) 54.1 (373.0)

70.7 (487.5) 64.7 (446.1)

63.0 (434.4) 62.9 (433.7)

67.5 (465.4) 59.5 (410.3)

Tensile strength, psi x lo3 (MPa) Cond. A Cond. D2/100

56.9 (392.3) 55.8 (384.7)

48.8 (336.5) 47.2 (325.4)

57.8 (398.5) 56.1 (386.8)

50.9 (351.0) 49.4 (340.6)

58.2 (401.3) 52.2 (359.9)

Tensile strength, psi X lo3 (MPa) Cond. A Cond. D2/100

Filament diameter does not appear to play an important role in the determination of the mechanical properties of such plied yarn fabrics as Style 1581 woven with ECG 150 1/2 (EC 9 33 1 x 2 yarn and style) and Style 1281 woven with ECB 150 1/2 (EC 3.8 33 1 x 2 yarn). However, in comparison with commercially available singles yarn Style 7781 woven with ECDE 75 1 / 0 (EC 6.66 1 x 0) yarn, singles yarn Style B 7781 woven with finer filaments [ECB 75 1/0 (EC3 8.66 1 x 0 yarn)] gives laminates with higher flexural strengths, flexural modulus and tensile strengths. Conversely a singles yarn G 7681 woven with coarser filaments [ECG 75 1/0 (EC9 66 1 x 0 yarn)] gives composites with improved compressive strengths.

The effect of variation in the available surface area for stress transfer is affected by the coupling agent-resin interaction. Larger surface areas with poor coupling agents reduces specific shear loading on the interface and so produces high laminate flexural strengths. But with improved interfacial bonding the surface areas effect disappears because other failure modes appear first. Singles yarns produce higher composite strengths due to the smaller angle between fiber axis and load axis in the lower twist yarns. This produces less tensile loading across the interface; the magnitude of this effect will vary with the resin coupling agent efficiency and the function of the twist of the yarn or the angle of the spiral of the filaments.

150 Fiberglass reinforcement Table 7.10 The effect of fabric construction on laminate mechanical properties2

Fabric construction Fabric no.

E1581

51581

YM31A"

Count/in (cm)

57 X 54 (22.4 x 21.3)

57 x 54 (22.4 x 21.3)

57 x 54 (22.4 x 21.3)

Warp yarn Glass system Tex system

ECG 150 1/2 EC933 1 x 2

SCG 150 1/ 2 SC933 1 x 2

MCG 150 1 / 2 MC9 33 1 x 2

Filling yarn Glass system Tex system

ECG 150 1/2 EC9 33 1 x 2

SCG 150 1/2 s c 9 33 1 x 2

MCG 150 1/2 MC9 33 1 x 2

Weave

Long-shaft satin

Long-shaft satin

Long-shaft satin

Laminate mechanical properties

Fabric no. Flexural strength, psi x lo" (MPa)

E1581

S1581

YM31A"

88.6 (610.9)

87.1 (600.6)

89.8 (619.2)

Flexural modulus of elasticity, psi x loh (GPa) Cond. A

3.78 (26.1)

3.79 (26.1)

5.00 (34.5)

Compressive strength, psi x lo3 (MPa) Cond. A

64.8 (446.8)

56.4 (388.9)

61.0 (420.6)

60.0 (413.7)

61.4 (423.4)

64.5 (444.7)

Cond. A

Tensile strength, psi x lo" (MPa) Cond. A a

Data from Reference 4

7.7.3 WEAVE PATTERN

The amount of yarn and the pattern of the weave often determine the handling characteristics of a fabric. If a fabric is too tightly woven, it will not readily drape and conform to various contours during molding; a tight weave can also adversely affect impregnation of the fabric by the resin. If a fabric has a weave pattern that is very open, this can produce a weaker reinforcement due to insufficient fiber, a tendency to distort and in horizontal treater, difficulty in impregnation. A general assumption can be made in the longer the float (Le. the portion of a warp or filling yarn that will extend unbound over two

or more yarns lying 90 degrees to it), the higher the reinforced composite strength. So by extending the length of the float which reduces the interlocking frequency, the resulting composite's strength will be increased. Changing the weave, for example from a plain weave to a crowfoot weave, will also achieve improved composites strength properties (Table 7.13). 7.7.4 GLASS-TO-RESIN RATIO

In general, approximation of physical properties for fiber glass reinforced composites follows the 'rule of mixtures'. The calculation of loads stresses and strains parallel to the fibers

Design considerations 151 Table 7.11 Composite mechanical properties as a function of glass-filament diameter

Plied-yarn fabrics ~-

Fabric style Count/in (cm) Warp yam Glass system Tex system Filling yarn Glass system Tex system Weave Flexural strength, psi x lo3 (MPa) Cond. A Flexural modulus of elasticity, psi x 106(GPa) Cond. A Compressive strength, psi x lo3 (MPa) Cond. A Tensile strength, psi x lo3 (MPa) Cond. A

181 57 x 54 (22.4 x 21.3)

1581 57 x 54 (22.4 x 21.3)

1281 57 x 54 (22.4 x 21.3)

ECE 225 1/3 EC7 22 1 x 3

ECG 150 1/2 EC9 33 1 x 2

ECB 150 l / 2 EC3.8 33 1 x 2

ECE 225 1 / 3 EC7 22 1x 3 Long-shaft satin

ECG 150 1 / 2 EC9 33 1 x 2 Long-shaft satin

ECB 150 1/2 EC3.8 33 1x 2 Long-shaft satin

93.0 (641.2)

88.6 (610.9)

92.4 (637.1)

3.88 (26.8)

3.78 (26.1)

4.06 (28.0)

58.1 (400.6)

64.8 (446.8)

64.9 (447.5)

60.0 (413.7)

64.8 (446.8)

60.0 (413.7)

~.

Singles-yarn fabrics ~ _ _ _ _

Fabric style Count/in (cm) Warp yarn Glass system Tex system Filling yarn Glass system Tex system Weave Flexural strength, psi x 10’ (MPa) Cond. A Flexural modulus of elasticity, psi x lo6 (GPa) Cond. A Compressive strength, psi x 10’ (MPa) Cond. A Tensile strength, psi x lo1 (MPa) Cond. A

G7681 60 x 54 (23.6 x 21.3)

DE7781

B7781

60 x 54 (23.6 x 21.3)

60 x 54 (23.6 x 21.3)

ECG 75 1/0 EC9661x0

ECDE 75 1/ 0 EC6 66 1x 0

ECB 75 1/0 EC3.8 66 1 x 0

ECG 75 1/0 EC9 66 1 x 0 Long-shaft satin

ECDE 75 1/0 EC6 66 1 x 0 Long-shaft satin

ECB 75 1/0 EC3.8 66 1 x 0 Long-shaft satin

92.6 (638.5)

94.9 (654.3)

101.1 (696.4)

3.78 (26.1)

3.90 (26.9)

4.02 (27.7)

76.7 (528.8)

67.7 (466.8)

69.9 (482.0)

64.2 (442.7)

66.3 (457.1)

69.4 (478.5)

152 Fiberglass reinforcement Table 7.12 Mechanical properties of S-glass fabrics-epoxy composites as a function of plied compared with unplied yarn S7682

S1581

Fabric style ~~~

~~

57681

-~ -

-

Finish

Volan A

Volan A

CS-310

Flexural strength, psi x lo3 (MPa) Cond. A

84.8 (584.7)

120.0 (827.4)

116.0 (799.8)

Flexural modulus of elasticity, psi x 10" (GPa) Cond. A

3.10 (21.4)

4.18 (28.8)

5.21 (35.9)

Compressive strength, psi x lo3(MPa) Cond. A

61.3 (422.7)

55.8 (384.7)

73.9 (509.5)

Tensile strength, psi x lo1 (MPa) Cond. A

70.6 (486.8)

87.5 (603.3)

86.4 (595.7)

Table 7.13 Effect of weave pattern on composite mechanical properties Fabric style

7628"

76281h

16-149'

7781d

Finish

Volan

Volan A

Volan A

Volan A

Resin

Polyester

Polyester

Polyester

Polyester

Plies

18

18

12

12

Resin content, wt.%

37.1

36.7

36.5

37.6

Thickness, in (mm)

0.124 (3.15)

0.121 (3.07)

0.120 (3.05)

0.120 (3.05)

Flexural strength, psi X 10' (MPa) Cond. A

53.8 (371.0)

84.7 (584.0)

63.2 (435.8)

87.0 (600.0)

Flexural modulus of elasticity, psi x 1Oh (GPa) Cond. A

3.89 (26.8)

3.41 (23.5)

3.81 (26.3)

3.24 (22.3)

Compressive strength, psi X lo1 (MPa) Cond. A

25.7 (177.2)

57.0 (393.0)

48.0 (331.0)

64.3 (443.3)

Tensile strength, psi x 103(MPa) Cond. A

45.9 (316.5)

59.2 (408.2)

58.7 (404.7)

60.0 (413.7)

Plain weave, data from Reference 3 Crowfoot satin weave, data from Reference 3

li

A 5-shaft satin weave version of Style 7781 An 8-shaft satin weave

Design considerations 153 and perpendicular to the fibers is quite different. Also, the stress transfer across the material boundaries is greatly affected by the degree of adhesion at the reinforcement-resin interface. A more subtle effect is the possibility that wall effects at the surface may alter the matrix orientation and mechanical properties. It must also be assumed that there will be sufficient resin present to fill all of the spaces between the fibers. Internal voids have been shown to produce considerable stress concentrations resulting in premature mechanical failure of the laminate. Woven fiberglass fabrics by virtue of the weave interlocking reduce the effective glass-to-resin ratio below those of filament wound structures. The optimum glass to resin ratio then becomes dependent on weave

pattern, weave density and yarn diameters. The optimum strengths of laminates are obtained with the lowest practical resin content. With coarse plain weave fabrics using hand lay-up, laminates can be produced with resin contents of 36-38%; with other weaves using less twist in the yarn and optimum yarn spacing, excellent physical properties can be achieved in the 25% resin content range. Plain weave fabrics with higher twist yarn will require higher resin contents than low twist singles fabrics of equivalent weave. In filament winding applications parallel strands of fiberglass are wound around a cylindrical mandrel, the resulting cross-section of round filaments provides a close packing, resulting in low resin contents of the final part.

Table 7.14 Laminate mechanical properties as a function of glass-to-resin ratio (Style 181 glass fabric-epoxy resin composites)

Plies

7

9

11

14

13 ~

~

16

18

~

20

___

Resin content, wt.%

55.0

49.2

44.0

35.6

31.2

28.0

22.5

22.2

Thickness, in mm

0.117 2.97

0.126 3.20

0.127 3.23

0.126 3.20

0.125 3.18

0.131 3.33

0.132 3.35

0.147 3.73

Flexural strength, Cond. A psi x103 MPa

45.6 314.4

58.8 405.4

64.8 446.8

83.0 572.3

81.2 559.9

92.1 635.0

87.4 602.6

91.9 633.7

Flexural modulus of elasticity, Cond. A psi x l o h GPa

2.26 15.6

2.62 18.1

2.92 20.1

3.33 23.0

3.82 26.3

4.04 27.9

4.37 30.1

4.64 32.0

Compressive strength, Cond. A psi x lo3 MPa

45.2 311.7

47.9 330.3

44.7 308.2

53.2 366.8

56.9 392.3

52.4 361.3

54.3 374.4

67.8 467.5

Tensile strength, Cond. A psi x 103 MPa

30.8 212.4

35.7 246.2

40.5 279.2

51.3 353.7

53.7 370.3

61.0 420.6

64.7 446.1

65.1 448.9

Tensile load/ply, lb N

514 2286

494 2197

468 2082

498 2215

480 2135

500 2224

476 2117

480 2135

154 Fiberglass reinforcement For example, resin contents of 25-30% with glass fiber content of 70-75% product final composites with high physical strengths. But, if there is any variation from the parallel alignment of the glass fibers, this will reduce the degree of packing and inevitably the optimum glass-to-resin ratio. The weaving of glass fabrics will of course reduce the effective glass-to-resin ratios because of the weave interlocking. The correlation between glass-to-resin ratios and composite strength is well known, the composite's strength increases as resin contents are reduced (see Table 7.14 and Fig. 7.4). 7.7.5 FIBER DISTRIBUTION

Reinforced composites that have the glass strands aligned parallel to each other have their maximum strength and stiffness in the direction of the alignment. This type of parallel alignment is conducive to certain filament winding and pultrusion operations. When the reinforcement is aligned at right angles to itself (half of the strands are laid at

right angles to the other half), the resulting mechanical strengths at either angle is less than that of the parallel alignment. As the distribution of strands varies between 0" to 90" alignment to a +45" to a 45", this further reduces mechanical strength in the primary direction, but shows an increase in the +45"and 4 5 " directions. Of course the yarn distribution can be varied from 0" and 90" directions as part of the fabric design. A so called %balanced' fabric, that is one with equal yarn distribution in the warp and filling directions, will have comparable composite properties in both directions. When fiberglass yarn is woven with yarn primarily in the warp direction with only the minimum amount of filling yarn (enough to give the fabric stability), the result is a unidirectional fabric. Composites reinforced with this type of fabric have the maximum mechanical strength in the direction of the greatest concentration of yarn5. The differences between the mechanical properties, bidirectional and unidirectional reinforced composites, can be seen in Table 7.15.

Table 7.15 Effect of fabric yarn distribution on laminate mechanical properties

Bidirectional (2582)

Fabric type (style) _ _ ~

Unidirectional (7743)

~

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78.2 (539.2) 68.5 (472.3)

95.0 (655.0) 87.4 (602.6)

23.3 (160.6) 24.1 (166.2)

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3.65 (25.2) 3.41 (23.5)

4.95 (34.1) 4.80 (33.1)

2.47 (17.0) 1.46 (10.1)

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63.4 (437.1) 54.5 (375.8)

64.5 (444.7) 51.1 (352.3)

28.9 (199.3) 23.6 (162.7)

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57.9 (399.2) 55.7 (384.0)

54.9 (378.5) 50.2 (346.1)

94.0 (648.1) 91.6 (631.6)

12.1 (83.4) 11.7 (80.7)

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REFERENCES 0.99 x lob (6.83) 0.94 x lo6 (6.48)

Compressive strength, psi (MPa) Cond. A 27 600 (190.3) Cond. D2/100 23 200 (160.0) Tensile strength, psi (MPa) Cond. A Cond. D2/100

Fiberglass remforcement aligned in a random manner within the polymer matrix (e.g. chopped strand mat) produces composites with fairly uniform mechanical strengths in all directions. However, this also tends to produce composites that have relatively low physical properties in all directions (Table 7.16).

14 100 (97.2) 13 800 (95.1)

1. Pitt, C.F. and Harvey J. 20th Anniversary Technical Conference, SPI Reinforced Plastics Division, 1965, Section 9-C. 2. Knox, C.E. Non-Metallic Materials (SAMPE) 4, 127 1972. 3. Horton, R.C. and Adams, R.G. 21st Annual Conference, SPI Reinforced Plastics Division, 1966, Section 10-A. 4. Knox, C.E. New Horizons in Materials and Processing (SAMPE) 18, 527 1973. 5. Peterson, G.P. Properties of High Modulus Reinforced Plastics, S.P.E.J., 57, 1961 January.

BORON, HIGH SILICA, QUARTZ AND CERAMIC FIBERS

8

Anthony Marzullo

tinuous. Continuous fiber or whisker is capable of being manufactured to an indefinite Boron, high silica, fused quartz and ceramic length. Discontinuous fiber can be chopped fibers are used in demanding industrial, autocontinuous fiber, or a short fiber called staple. motive, electronic, aircraft and aerospace Discontinuous whisker is manufactured to a environments. Boards made of ceramic fibers definite length in batch type processing, while such as alumina, alumina-silica or zirconia are continuous whisker is produced by melt used as supplemental linings to insulate high processes such as the laser-heated floatingtemperature furnaces. Ceramic tiles containzone process. ing alumina and silica fibers protect the Summarizing, there is continuous fiber aluminum skin of USA's space shuttle orbiter (indefinite length), continuous whisker (single during re-entry, where the tile's surface may crystal, indefinite length), discontinuous fiber be aerodynamically heated to 1260°C (Banas, (chopped continuous fiber or staple fiber) and McCormick and Creedon, 1991). Other applidiscontinuous whisker (single crystal, definite cations of these fibers involve reinforcing length). polymer-, glass-, ceramic-, or metal-matrix composites. If the composite is well designed, it will be tougher than the matrix material by 8.2 MANUFACTURE itself. As an example, the fracture toughness of Sic is 1.5 MPa m-2,but 8-15 for Sic reinforced Melt, vapor deposition and chemical processes are used to manufacture fibers and whiskers. with Sic fibers (Richerson, 1992). Fibers and whiskers that are made by melt Fibers can be classified according to strucprocessing include continuous quartz fibers ture, diameter (or cross sectional width) and by drawing from a fused quartz rod, alulength. In general, a material is classified as a mina-silica staple by atomizing a molten fiber if its diameter or cross sectional width is ceramic stream and continuous alumina less than 0.0254 m (0.010 in) and 1ength:diamwhiskers by drawing from a melt. Vapor depoeter ratio is greater than 1O:l.Most commercial sition processes include chemical vapor fibers meet these requirements easily. A fiber is deposition of silicon carbide or boron onto a called a whisker if its microstructure is predominantly single crystal. Fibers or whiskers tungsten core to make continuous fibers and can be subclassified as continuous or discon- the vapor-liquid-solid (VLS) process to make discontinuous whiskers. Chemical processing called sol-gel technology is used to create alumina-based continuous fiber. Other chemical Handbook of Composites. Edited by S.T. Peters. Published processes include the creation of polymer in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

8.1 INTRODUCTION

Continuous fibers precursor fiber that is later pyrolyzed into silicon, carbon and nitrogen-based continuous fiber and the reaction between silica and carbon to make discontinuous whiskers. The manufacturing processes are summarized in Table 8.1. 8.3 CONTINUOUS FIBERS

8.3.1 HIGH SILICA FIBERS

Amorphous high silica fibers can be manufactured by leaching sodium oxide from sodium silicate glass fibers. The glass fibers, made by using conventional glass fiber production techniques, are composed primarily of 74.5%)silica and 24.2% sodium oxide. The fibers are held in a perforated basket, which is suspended within a cylindrical tank. A solution of HC1 and water is pumped from the bottom of the tank up to and over the glass fibers for several hours. The HCl/water/fiber mixture is stirred periodically. The HCl and water are drained away, the fibers rinsed and a second leaching and rinsing cycle is performed. The fibers are dried at a low temperature (Price and Kielmeyer, 1980).Other techniques used to make high silica fibers include spinning from solutions and spinning from a gel (Cooke, 1991).

157

fibers with a minimum of surface flaws, the rod’s surface should be flame polished with an oxy-hydrogen flame, as is done in optical fiber manufacturing (Blyler and DiMarcello, 1991). 8.3.3 CONTINUOUS FIBERS BY CHEMICAL VAPOR DEPOSITION

Producing silicon carbide fibers by chemical vapor deposition (CVD) begins by pulling a melt-spun carbon fiber through a mercury seal and into a tubular reactor. The carbon fiber is heated by coupling the mercury seals on the top and bottom of the reactor to a source of electricity. The melt-spun carbon fiber is called a substrate or a core because silicon carbide is vapor deposited onto it. The deposited silicon carbide is produced by the reaction of a gas mixture of silanes and hydrogen that enters the reactor through a port. Details of producing silicon fibers by CVD are found in Wawner, Jr. (1988) and Textron Specialty Materials (1993b). Producing boron fibers by CVD is a similar process (Tsirlin, 1985; Wawner, Jr., 1988; DeBolt, 1982). 8.3.4 PYROLYSIS OF POLYMERIC PRECURSOR FIBER

Several ceramic fibers composed of various combinations of silicon, carbon, nitrogen, 8.3.2 DRAWING CONTINUOUS FUSED boron and titanium have been produced by QUARTZ FIBERS the pyrolysis of polymeric precursor fiber. The Amorphous fused quartz fibers are made from process consists of four major steps: synthesize preforms of fused quartz rods. The end of a a spinnable polymer, spin the polymer into fused quartz rod is heated with an oxy-hydro- precursor fiber, cross-link (cure) the precursor gen flame and the fiber is drawn, sized and fiber and pyrolyze the cured precursor fiber either taken up on a spool or combined with into ceramic fiber. other simultaneously drawn fibers to form a The synthesis of a polymeric precursor is strand. Fused quartz rods are made by fusing known as Preceramic Polymer Chemistry. chemically purified ground quartz crystal or Many spinnable polymers have been synthequartz sand. Fused quartz rods are classified sized, including polycarbosilane as a as Type I or Type I1 vitreous silica depending precursor for silicon carbide fiber (Yajima, on the fusion technique. Type I signifies that 1985),polytitanocarbosilane as a precursor for fusing was done in an electrically heated cru- Si-Ti-C-0 fiber (Yamamura et al., 1988) and cible, while Type I1 implies fusing with an hydridopolysilazane as a precursor for silicon oxy-hydrogen flame. To obtain fused quartz carbonitride fiber (LeGrow et al., 1987). In

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operation. They claimed that the fiber properties were improved due to uniform heating and even fiber tensioning (Koba et al., 1989). Sumitomo Chemical Co invented a fiber with 85% gamma or delta-alumina and 15%silica by calcining dry-spun fibers. The spinnable mixture was formed by mixing an aluminum compound in a solvent and adding water to get polyaluminoxane. A silica containing compound, an organic polymer to improve spinnability and a compound such as lithium to improve fiber properties were added. The solvent was removed by distillation. The mixture is spun into fiber and the fiber is calcined (Kadokura, Harakawa and Matumoto, 1991). Minnesota Mining and Mfg. Co. (3M) makes fibers composed of alumina, boria and 8.3.5 ALUMINA BASED FIBER silica by drying and sintering fiber that is spun E.I. du Pont de Nemours patented a process to from concentrated aqueous solutions. The make continuous alumina fibers that involved aqueous solution can be composed of an aluspinning a viscous slurry. One slurry composi- mina and boria precursor such as a basic tion was prepared by dissolving an alumina aluminum acetate, silica sol and water. The precursor such as aluminum chlorohydroxide 14%boria in 3M’s Nextel 312 fiber mhibits the in water and then adding solid alumina parti- formation of alumina crystals and aluminum cles. An organic polymer might be added to borosilicate is the predominant crystalline the slurry to aid its spinnability by modifying species. The 2% boria in Nextel 440, however, its viscosity. The viscous slurry was extruded is not enough to inhibit alumina, with etathrough a spinneret and the green (unfired) alumina being the predominant crystalline fibers were wound onto a bobbin. The fibers species (Sowman, 1988). were fired at a temperature below the sintering temperature (calcining) to remove water 8.4 STAPLE and volatile material. The calcined fibers were sintered (densified) to remove porosity and The process to make alumina staple is similar convert the alumina precursor to the stable to the processes that are used in making alumina-based continuous fibers. An aqueous alpha-alumina structure. This slurry-spinning process was advanced and viscous solution of an alumina precursor, by Mitsui Mining Co. where the fiber was spun, silica and an organic polymer is extruded dried, calcined and sintered in a continuous through apertures into a high velocity gas stream, creating green (unfired) staple. The green staple is sintered. The silica is added to toughen the resultant staple by inhibiting aluFootnotes for Table 8.1 mina crystal growth while the green staple is a Most of the trade names are registered trademarks. Each being sintered (Bunsell et al., 1988). trademark belongs to the manufacturer listed to its left. Imperial Chemical Industries (ICI) has a All data on composition for the continuous and staple commercial alumina staple called Saffil. Saffil fibers, except for Tyranno Fiber, are from the manufacturers’ product literature. Tyranno Fiber is 3 ym in diameter and its length is 1-5 cm in composition is from Weddell (1990). the mat form. Saffil comes in six grades ranging

160 Boron, high silica, quartz and ceramicfibers from the catalytic grade with an eta-alumina the end of a polycrystalline feed rod is melted structure, to an aerospace grade with a delta + with a CO, laser, a grain free seed crystal is alpha-alumina or an alpha + delta-alumina + brought in contact with the melt and a whisker is drawn. The LHFZ process can be used to mullite structure. produce whiskers with diameters as low as Alumina-silica staple, also called alumi10 pm or as high as 10 pm (Haggerty,Wills and nosilicate staple, is made by atomizing a Sheehan, 1991). molten ceramic stream. The raw material for the melt is a clay mineral or a blend of alumina and silica. The melt is atomized by pouring it 8.5.2 DISCONTINUOUS WHISKERS BY THE VLS into the path of a high velocity air stream or by PROCESS feeding it onto a rotating wheel that throws out the melt by centrifugal force. The former is VLS stands for vapor feed gases-liquid catacalled the blowing process, while the latter is lyst-solid crystalline whisker growth. In a called centrifugal spinning. One commercial particular VLS process to make single crystal alumina-silica staple called Fiberfrax ranges Sic whiskers, uniformly sized catalyst particles in length from 70-1000 pm and is made by The are applied to carbon substrates. The carbon substrates and catalyst are placed within a Carborundum Co. growth chamber. Also within the chambers are Si0 generators, made by impregnating porous brick with carbon and SiO,. The chamber is 8.5 WHISKERS placed within a furnace and heated to a growth 8.5.1 CONTINUOUS WHISKERS temperature of 1300-1450°C. During t h s time, Single crystal continuous alumina whiskers can gases such as CO, CH,, N, and H, are fed into be grown from a pure melt by using a modifi- the chamber. The catalyst particles melt and cation of the Czochralski method. The modified become supersaturated with material supplied method is still classified as a moving crystal from the vapor and carbon from the substrate. techruque, but, unlike Czochralslu’s technique, Solid Sic precipitates from the liquid catalyst uses a die to shape the growing whisker. A cap- onto the growth substrate. As precipitation conillary tube draws the melt up to and over a die. tinues, the whisker grows, lifting the catalyst The melt flows to the die’s vertical edges and ball from the substrate (Petrovic and Hurley stops; so the die’s vertical edges control the 1990; Milewski et al., 1985). The time necessary to produce Sic whiskers shape of the subsequently drawn whisker. A grain free seed crystal is brought in contact with by the VLS process can be reduced by underthe melt surface and a whisker is drawn. A lin- cooling or pre-alloying the catalyst balls. After ear temperature distribution is maintained reaching the growth temperature, the catalyst along the axis of the growing crystal to avoid balls are undercooled by 150”C,initiating predefects by lowering thermal stresses (LaBelle, cipitation. The chamber is brought back up to Jr, 1980; Antonov, 1990). Saphikon Inc. grows the growth temperature and held there until whiskers in many shapes using a similar whisker growth is completed. Another techprocess that they call EFG (edge-defined,film- nique to reduce the growth time is to pre-alloy fed growth). One of the shapes is a hollow tube the catalyst particles with carbon and silicon that may be used as a waveguide (Harrington so that whisker growth occurs immediately upon reaching the growth temperature and Gregory 1990). Another moving crystal technique to make (Shalek, 1987; Shalek, Katz and Hurley, 1988). continuous alumina whiskers is called the Other whiskers produced by the VLS laser heated floating zone directional solidifi- process include A1,0,, boron, MgO and Si cation process (LHFZ). In the LHFZ process,

Forms 161 (Wagner, 1970); Tic by using TiC1, and CH, as the primary feed gases and a nickel catalyst (Pearson and Easley, 1992; Narasimhan and Bhat, 1992);Si,N, by using Si,Cl, and NH, as the primary feed gases and an iron catalyst (Motojima et al., 1989);and TiB, by using TiC1, and BC1, as the primary feed gases and Au or Si as the catalyst (Withers, Loutfy and Lee, 1988). 8.5.3 OTHER DISCONTINUOUS WHISKER GROWTH PROCESSES

An early vapor phase growth process to make alumina whiskers involved hydrogen reduction. Wet hydrogen was passed over heated aluminum powder. Whisker lengths and diameters were from 1-30 mm and 3-50 pm, respectively. Hydrogen reduction of methyltrichlorosilane was used to produce p-Sic whiskers (Milewski, 1991).Other vapor phase growth techniques are evaporation-condensation and vapor phase reaction. These and other growth techniques, are discussed in Bracke, Schurmans and Verhoest (1984). Some commercial processes to make Sic whiskers use a chemical reaction between silica and carbon powders. Tokai Carbon manufactures p-Sic whiskers by heating a mix of pulverized silica gel, carbon black and a metal catalyst. The mix is heated in a nonoxidizing atmosphere at 1300-1700°C and the whiskers grow within the pores of the mix (Yamamoto, 1985).The whiskers are under 1.0 pm in diameter and under 300 pm in length. Other companies have produced Sic whiskers by chemical reactions: the J.M. Huber Corp. passed perforated trays filled with ground rice hulls over heated carbon pellets and through a horizontal furnace. Ground rice hulls were used as the reactant material because they contain both silica and carbon (Tanaka, Kawabe and Kobune, 1986). Silicon nitride whiskers have been produced by heating a silica and carbon mix or rice hulls in a nitrogen atmosphere (nitriding), or by heating silica powder in a gas mixture of ammonia gas and hydrocarbon gas.

8.6 PROPERTIES

Boron, h g h silica, quartz and ceramic fibers have a high thermal and chemical stability relative to organic fibers. Maximum use temperatures of boron, h g h silica, quartz and ceramic fibers, except for boron fiber in an oxidizing atmosphere, all are over 1000°C. These fibers have a good balance of specific strength (tensile strength/density ratio) and specific modulus (elastic modulus/density ratio) compared with other materials. Tsirlin’s (1985) graphs of specific strength against specific modulus for boron or Sic fibers with cores illustrate this point. In Table 8.2 note the low coefficient of thermal expansion for quartz fibers, which allows them to be used in parts that are subjected to high temperatures, such as in high speed printed circuit boards. 8.7 FORMS

The form that a fiber can take depends on its length, diameter and mechanical properties, such as elastic modulus and tensile strength. In general, discontinuous whiskers are too short to be spun into yarn, however, VLS whiskers about 75-100 mm long might be spun into staple yarn (KO, 1989).A fiber’s flexibility is related to its diameter, elastic modulus and tensile strength. A measure of flexibility is l / [ d ( E / o- l)]where d = diameter, E = elastic (Young’s)modulus and cs = tensile strength. A greater value indicates a greater flexibility. Obviously, the greater a fiber’s flexibility, the easier it is to twist, braid and weave into complex shapes without breakage. As an example, l / [ d ( E / o - l)]is about 126 for continuous Sic fiber with a core (d = 80 pm), 533 for continuous alumina fiber, 1298 for continuous Sic fiber (d = 12 pm) and 10 101 for continuous fused quartz fiber. The Sic fiber with a core, having a lower flexibility relative to other fibers, is generally used for unidirectional reinforcement. The fiber might also be used in a material-geometric hybrid to axially reinforce a woven preform of finer

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Footnotes for Table 8.2 All values, except those for silicon carbide whiskers, are from the manufacturers’ product literature. Test methods may varv between manufacturers. All values are averages, are for reference only and are not to be taken as a specification. Most fibers are under continual development. Contact the manufacturers for the most up to date property values. DNA = Data Not Available. The silicon carbide whisker values are from Petrovic and Hurley (1990). This value in metric units was converted from a value that was given in English units. Most of the tradenames are registered trademarks. Each trademark belongs to the manufacturer listed to its left

and two diagonal directions. Figure 8.1 shows the 3D and 5D fabrics. See Ngai (1990) for diagrams of 4D, 7D and 11D fabrics. Continuous boron fibers are supplied in the form of borodepoxy prepreg tape and boron/aluminum preform sheets. The boron/aluminum preform sheets are made by resin bonding the boron fibers to the side of aluminum foil. The fibers can be aligned in any direction. These preform sheets are used in solid state (low temperature, high pressure) diffusion bonding processes (TextronSpecialty Materials, 1993a). Continuous Sic fibers with a core are formed into plasma-sprayed aluminum or titanium preform sheets for hot molding (a low pressure, hot pressing process) and into ’woven fabric’ for investment casting or diffusion bonding. In diffusion bonding, the fabric is placed between two metal foils prior to consolidation. The ’woven fabric’ is similar to nonwoven roving because the fibers are aligned in one direction and held together by a cross-weave of metallic ribbons (Textron Specialty Materials, 199313; Mittnick 1990). Continuous high silica fibers are processed into a variety of forms such as cloth, tape, sleeving, yarn, cordage, rope, nonwoven mat, blanket, felt, rigid tile and sewing thread. Alumina-silica staple fibers are made into paper, rigid board and cylinders, other rigid three-dimensional shapes and into most of the high silica fiber forms. Alumina staple fibers are processed into paper, mat, felt, blanket, rigid board and cylinders and other rigid three-dimensional shapes.

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8.8 APPLICATIONS

Table 8.3 provides a sampling of the many applications for boron, high silica, fused quartz and ceramic fibers. For other applications, see Bracke, Schurmans and Verhoest (1984), Mortensen and Koczak (1993), Schwartz (1992) and the other chapters within this handbook.

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166 Boron, high silica, quartz and ceramic fibers 8.8.1 NONCOMPOSITES

Alumina and alumina-silica staple fibers and the continuous high silica fibers are used in high temperature applications in many industries. Forms of these fibers are used for thermal insulation, personnel protection and gasketing. See Carborundum (1991) for the many applications of high silica fibers and Carborundum (1989) for the many applications of alumina-silica fibers. 8.8.2 COMPOSITES

Fibers and whiskers have been used to reinforce polymers, epoxies, glasses, metals, or ceramics. The reinforced materials are called composites in general and, more specifically, polymer, epoxy, glass, metal, or ceramic matrix composites. The reinforcements are either discontinuous whiskers or continuous fibers for the applications in Table 8.3, but staple fiber and continuous whisker may be used as a reinforcement material as well. Usually the fiber or whisker is coated to cover its surface flaws, provide a diffusional barrier between it and the matrix material and protect the fiber or whisker from oxidation. In addition, the coating provides a weak interface that deflects cracks propagating through the matrix in ceramic matrix composites REFERENCES

Antonov, P.l. 1990. Review of Factors Controlling the Growth of Shaped Crystals. J. Crysf. Growth 104: 3946. Banas, R., McCormick, M. and Creedon, J. 1991. HTP-6.5-22 and HTP-1-22, Reusable Rigid Fibrous Ceramic Materials for Heat Shields on Future Reentry Vehicles. In Ceramics Today Tomorrow’s Ceramics, ed. P. Vincenzini, pp. 2889-98. Amsterdam: Elsevier Science Publishers B.V. Blyler, Jr., L.L. and DiMarcello, F.V. 1991. Optical Fibers, Drawing and Coating. In Encyclopedia of Lasers and Optical Technology, ed. Robert A. Meyers, pp. 457-67. New York: Academic Press Bracke, I?, Schurmans, H. and Verhoest, J. 3984.

Inorganic Fibres and Composite Materials. New York: Pergamon Press. Brown, A.S. 1993. High-temperature materials warm up for debut. Aerospace America 31 (3): 18-27. Buljan, S.T., et al. 1992. ’Development of Ceramic Matrix Composites for Application in Ceramic Technology for Advanced Heat Engine Program’, Contract DE-AC05-840R21400, U.S. Department of Energy, April 1992. Bunsell, A.R. ef al. 1988. Ceramic fibers. In Fibre Reinforcenien ts for Composite Materials, ed. A.R. Bunsell, pp. 427-78. Vol. 2 of Composite Materials Series, Series ed. R.B. Pipes. Amsterdam: Elsevier Science Publishers B.V. Carborundum 1989. ‘Fiberfrax Ceramic Fiber’. The Carborundum Company. Niagara Falls, NU, USA. Carborundum 1991. ’Refrasil Products. Product Specifications’. The Carborundum Company. Niagara Falls, NY, USA. Cooke, T.F. 1991. Inorganic Fibers - A Literature Review. I. A m Cernm. Soc. 74 (12): 2959-78. Covington, M.A. and Sawko, P.M. 1986. Optical Properties of Woven Ceramic Fabrics for Flexible Heat Shields. AIAA/ASME 4th Joint Thermophysics and Heat Transfer Conference, 2 4 June 1986, Boston, Mass. DeBolt, H.E. 1982. Boron and Other High-Strength, High Modulus, Low Density Filamentary Reinforcing Agents. In Handbook of Composites, ed. George Lubin, pp. 171-95. New York: Van Nostrand Reinhold. Fiber Materials, Inc. 1993. ’3-D Orthogonal Woven Blocks’ and ’Multi-Directional Woven Preforms’. Fiber Materials, Inc. Biddeford, Maine, USA. Haggerty, J.S., Wills, K.C. and Sheehan, J.E. 1991. Growth and Properties of Single Crystal Oxide Fibers. Ceram. Eng. Sci. Proc. 12 (9-10): 1785-1801. Harrington, J.A. and Gregory, Christopher C. 1990. Hollow sapphire fibers for delivery of CO, laser energy. Optics Letters 15 (10): 54103. Kadokura, H., Harakawa, M. and Matumoto, T. 1991. Process for Producing Alumina-Based Fiber. US Patent 5002 750. Mar. 26, 1991. Assigned to Sumitomo Chemical Co Ltd. KO, F.K. 1989. Preform Fiber Architecture for Cerami-Matrix Composites. Ceramic Bulletin 68 (2): 401-14. Koba, K., et al. 1989. Continuous Process for Producing Long a-Alumina Fibers. US Patent

Refcremes 4 812 271. Mar. 14, 1989. Assigned to Mitsui Mining Co Ltd. LaBelie, Jr., H.E.1980. EFG, The Invention and Application to Sapphire Growth. J . Crysf. Growth 50 (1980): 8-1 7. L,asday, S.B. 1993. Ceramic/Ceramic Composite Components Advance Furnace Systems and Processes. lndiistriai Heating LX (4): 26-9. LeGrow, G.E., et al. 1987. Ceramics from tlydridopolysilazane. A m Ceram. Soc. B d l . 66 (2): 363-67. Milewski, J.V., et al. 1985. Growth of Beta-silicon Whiskers by the VLS Process. 1. Mater. Sci. 20 (1985): 1160 66. Milewski, J.V. 1991. Whiskers. In Concise Encyciopediu oj Advanced Cmutric hlatwials, ed. R.J. Brook, pp. 516-19. Oxford: Pergamon Press. Mittnick, M A . 1990. Continuous S i c Fiber Reinforced Metals. SME Metal Matrix Clinic, 13-14 R-ovember 1990, at Anaheim, Ca. Mohn, W.R. and Vukobratovich, D. 1988. Engineered Metal Matrix Composites for Precision Optical Systems. SAMPE J . 24 (1):26--35. Mortensen A. and Koczak, M.J. 1993. The Status of Metal-Matrix Composite Research and Development in Japan. ]OM 45 (3): 10-18. Motojimi, S., et nl. 1989. Preparation of Whiskers and Spring-like Fibres of Si,N, By ImpurityActivated Chemical Vapor Deposition. J. Cryst. Growth 96: 383-89. hhik, S.K., et a / . 1991. Ceramic Matrix Composites and its Application in Gas Turbine Engines. Presented at the International Gas Turbine and Aeroengine Congress and Exposition, 3-6 June 1991, in Orlando, F1. Narasimhan, K. and Bhat, D.G. 1992. Process for Producing Single Crystal Titanium Carbide Whiskers. US Patent 5094711. Mar. 10, 1992. Assigned to GTE Valenite Corporation. Ngai, T. 1990. Carbon-Carbon Composites. In lnternational Encyclopedia of Composites, ed. Stuart M. Lee, pp. 158-87. New York: VCH Ptiblishers. Pearson, A. and Easley, M.A. 1992. Process For Production of Small Diameter 'Titanium Carbide Whiskers. US Patent 5 160 574. Nov. 3, 1992. Assigned to Aluminum Company of America. Petrovic, J.J. and Hurley, G.F. 1990. Vapor-Liquid-Solid (VLS) S i c Whiskers: Synthesis and Mechanical Properties. In Fiber Reinforced Ceramic Conzposites, ed. K.S. Mazdiyasni, pp. 93-121. New Jersey: Noyes Publications.

167

Price, G.B. and Kielmeyer, W.H. 1980. Method for Making I-ligh Purity, Devitrification Resistant, Amorphous Silica Fibers. US Patent 4 200 485. Apr. 29, 1980. Assigned to John-Manville Gorp. Richerson, D.W. 1992. Modern Ceruniic Engineering, 2nd edn. New York: Marcel Dekker, Inc. Schwartz, M.M. 1992. Composite Materials I$andbook, 2nd edn. ed. M.M. Schwartz. New York: McGr a w-Hill , Inc . Shalek, P.D. 1987. Process for Growing Silicon Carbide Whiskers by Undercooling. US Patent 4 702 901. Oct. 27, 1987. Assigned to The United States of America. Shalek, P.D., Katz, J.D., Hurley, G.F. 1988. Prealloyed Catalyst for Growing Silicon Carbide Whiskers. US Patent 4 789 537. Dec. 6, 1988. Assigned to The United States of America. Schramm, Dale E. 1988. Process for Producing Silicon Carbide Whiskers. US Patent 4 789 536. Dec. 6, 1988: Assigned to J.M. Huber Corp. Sowman, H.G. 1988. Alumina-Boria-Silica Ceramic Fibers from the Sol-Gel Process. In Soi-Cel Technology for 7 h i n Films. Fibers. Prefornzs. Electronics. und Specialty Shapes, ed. Lisa C . Klein, pp. 162-83. New Jersey: Noyes Publications. 'Takeda, M. rt al. 1992. Thermal Stability of the Low Oxygen Silicon Carbide Fiber Derived From Polycarbosilane. Ceram. Eng. Sci. Proc. 13 (7--8): 209-1 7. Tanaka, M., Kawabe, T. and Kobune, M. 1986. Method of Manufacturing Crystalline Silicon Carbide Employing Acid Pretreated Rice Husks. US Patent 4 591 492. May 27, 1986. Assigned to Tateho Kagaku Kogyo Kabushiki. Tech Trends 1990. Ceramic Matrix Composites: Technology and fndustriul Applications. Paris: lnnovation 128. 'Textron Specialty Materials 1993a. Boron Aluminum Preform Sheets. Textron Specialty Materials. Lowell, MA, USA. Textron Specialty Materials 1993b. Continuous Silicon Carbide Metal Matrix Composites. Textron Specialty Materials. Lowell, MA, USA. Tricoles, G.P. 1988. Radome Electromagnetic Design. In Antenna Handbook, ed. Y.T. Lo and S.W. Lee, pp. 31-1 to 31-31. New York: Van Nostrand Reinhold. Tsirlin, A.M. 1985. Boron Filaments. In Strong Fibres, ed. W. Watt and B.V. Perov, pp.155-99. Vol.1 of Handbook of Composites, Series ed. A. Kelly and Yu. N. Rabotnov. Amsterdam: Elsevier Science Publishers R.V.

168 Boron, high silica, quartz and ceramicfibers Wagner, R.S. 1970. VLS Mechanism of Crystal Growth. In Whisker Technology. ed. Albert P. Levitt, pp. 47-119. New York John Wiley & Sons. Wawner, F.E. 1988. Boron and Silicon Carbide/Carbon Fibers. In Fibre Reinforcements for Composite Materials, ed. A.R. Bunsell, pp. 371425. Vol. 2 of Composite Materials Series, Series ed. R.B. Pipes. Amsterdam: Elsevier Science Publishers B.V. Weddell, J.K. 1990. Continuous Ceramic Fibers. 1. Text. Inst. 81 (4): 333-59. Withers, J.C., Loutfy R.O. and Lee, C.T. 1988. Process to Produce Titanium Diboride Whiskers as Reinforcement for Metal and Ceramic Composites. NSF Grant ISI-8760300. National Science Foundation, Washington D.C. October 1988.

Xiao, H., Ai, X and Yang, H.S. 1993. Effect of whisker orientation on toughening behavior and cutting performance of SiCW-A1203 composite. Mat. Sci. Technol. Vol. 9:21-25. Yajima, S. 1985. Silicon Carbide Fibers. In Strong Fibres, ed. W. Watt and B.V. Perov, pp. 201-37. Vol.1 of Handbook of Composites, Series ed. A. Kelly and Yu. N. Rabotnov. Amsterdam: Elsevier Science Publishers B.V. Yamamoto, A. 1985. Process for Preparing Silicon Carbide Whiskers. US Patent 4 500 504. Feb. 19, 1985. Assigned to Tokai Carbon Co Ltd. Yamamura,T., et al. 1988. Development of a New Continuous Si-Ti-C-0 Fibre Using an Organometallic Polymer Precursor. 1.Mater. Sci. 23: 2589-94.

CARBON FIBERS

9

Khalid Lafdi and Maurice A. Wright

9.1 INTRODUCTION

microstructural features developed in the spinCarbon fibers exhibit truly outstanding prop- ning process that also influence the stiffness, erties. As shown in Table 9.1, their strength, 0, strength and the thermal and electrical propercompetes with the strongest steels; they can ties of pitch can be optimized. Nevertheless, have stiffness, E , greater than any metal, the tensile strength of PAN-based fibers has ceramic or polymer; and they can exhibit ther- always been superior to pitch-based fibers; since PAN fibers were developed before pitch mal and electrical conductivities that greatly fibers, most structural materials and compoexceed those of competing materials. If the nents use PAN-based fibers. strength or stiffness values are divided by the The most important mechanical and physilow density, p, 1800-2200 kg m-3, then their cal properties exhibited by carbon fibers are the huge specific properties (o/p, E/p) make this elastic modulus, tensile strength and the electriclass of materials quite unique. cal and thermal conductivities. These All carbon fibers sold commercially are fabproperties are sensitive to the crystallite size ricated from polyacrylonitrile (PAN) or from a and perfection of the graphene layers develcoal, petroleum or synthetic pitch. PAN-based oped within the carbon fiber and depend for fibers are produced from a solubilized mixture the most part on the degree of molecular alignthat is wet or dry spun to produce a fiber, ment with respect to the fiber axis. Growth and ostensibly for use in the textile industry. This alignment of such layers occur within the prefiber is stabilized and carbonized to produce a cursor and within the solid carbon fiber when it carbon fiber. Aerospace grade material can be is heated to high temperature. However, it can obtained in tows that contain between 3000 be argued that the most extensive structural and 12 000 fibers. Lower performance materirearrangement occurs when the basic structural als are usually formed using larger tows that units (BSUs) present in the original precursor contain up to 320 000 fibers. PAN-based carare large and plate-like so that the shear stresses bon fibers are cheaper when produced from generated during spinning can align these large larger tows. areas more easily. Heating the fiber to high temPitch fibers are melt spun products obtained in small tow sizes varying from 2000 to 4000 perature after spinning relaxes the spinning stresses and allows the oriented regions to fibers. They are usually of larger diameter grow. Compared to PAN, the basic structural (- 10-15 pm) than fibers formed from PAN. units in an original mesophasic pitch are much The spinning process is controlled by the larger in area and length and are not twisted to pitch-based carbon fiber producer; thus, the same degree. The resulting ease of graphitization and alignment of the graphene layers and the development of large crystallites, proHandbook of Composites. Edited by S.T. Peters. Published duces a large elastic modulus and electrical and in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

-

170 Cnrbon fibers thermal conductivity parallel to the fiber axes. Unfortunately, large graphitized regions tend to produce high local stress concentrations, especially if they are misaligned (Reynolds and Sharp, 1974; Reynolds and Moreton, 1980). 'rhus, pitch-based fibers tend to exhbit high modulus and high electrical and thermal conductivities, but low strength. PAN-based fibers tend to be of intermediate modulus and relatively high strength. The mechanical properties of structural metals generally vary in direct proportion to density. Thus, as shown in Table 9.1, the specific

properties (o/p, E/p) tend to remain relatively constant; structures designed to resist a given set of loads tend to weigh the same irrespective of whether they are made from aluminum, titanium or steel. These mechanical property/ density relationships are not followed by carbon fibers; so, compared with metals, similar structures made from carbon fiber reinforced composites will be lighter. The most successful use of carbon fiber reinforced materials has been for military purposes, especially aeronautical structures (Schwartz, 1984; Hadcock, 1982). Composite

Table 9.1 Physical and mechanical properties of typical carbon fibers -

Siipplier/Fiber name

Tensile strnzgth, a (GPd

Elastic rnoduliis, E (CPal

Fracture elongation,

Density, dp. p ( k g / ~ ~ IOp ) rn? s-:

E/P > I O h rn2s-'

%

Pitch Type Amoco /Thornel

PlOO P120

2.2 2.4

690 830

0.3 0.3

-2150 -2150"

1.02 1.11

321 386

Petoca /Carbonic

HM50 HM60 HM80

2.8 3.0 3.5

490 590 790

0.6

-2000"

1.40

245

0.5 0.4

-2000"

1.75

395

T-40

3.45 2.9 5.65

231 390 290

NA NA NA

1760 1810 1810

1.96 1.60 3.12

131 215 160

Akzo / Fortafil

F-~(c) F-3(c)

2.76 3.80

345 227

NA NA

1800 1700

1.53 2.24

192 133

RK Fibers/RK

RK30 RK25

~3.0 >2.5

-230 -230

NA NA

1780 1780

1.68 1.40

129 129

Toray/Torayca

M46J T300

436 230 294

0.5 1.2 2.0

1750

1.71

143

T800

4.2 3.5 5.5

ST-1 ST-2 ST-3

3.6 4.0 4.4

240 240 240

1.5 1.7 1.8

-1800b -1800b -1800b

2.00 2.20 2.38

160 141 133

0.172 0.324 0.414

73 110 199

NA NA NA

2720 4500 7860

0.063 0.072 0.052

27 24 25

PAN Type Amoco/Thornel

T-300

T-50

Toho/Besfight

Metals Aluminum' Titanium' Steel'

"Estimated from Tanabe et al. 1987; hEstimated from Johnson 1987; 'Eshbach and Souders, 1974; NA, Not Available.

Overview 171 use in civilian aviation has been more limited, largely because of cost. However, in 1985 the European Airbus consortia used carbon/epoxy vertical stabilizers on their A310 and, since 1993, delivered their A340 and A330 models equipped with composite tail sections, floor panels, landing gear doors and carbon-carbon brakes. The new Boeing 777 and the projected McDonnell-Douglas MD-12 contains similar composite structures (High-Performance Composites, 1994).

a precursor material is prepared, spun into a fibrous shape, stabilized to change it from a thermoplastic to a thermoset, then heated until all unwanted elements are expelled. Depending on the final heat treatment temperature, a fiber is produced that is primarily carbon. A strikingly similar production technique is used to produce either type of fiber; however, the initial pretreatment and the chemical reactions that occur within either PAN or pitch during stabilization and carbonization are markedly different. These differences are discussed in the following sections since they determine the ultimate properties and the cost of the carbon fibers that are produced.

9.2 OVERVIEW

Ln this chapter we have chosen to combine, rather than separate, the discussion of the preparation of carbon fibers from PAN or from pitch. This style of presentation is attempted since the methods used to fabricate carbon fibers from either of these precursor materials seem to be almost identical, as shown in Fig. 9.1; for example, both methods involve the preparation, spinning and subsequent thermal degradation of organic precursors. Essentially,

9.2.1 POLYACRYLONITRILE

The chemical composition of PAN-based precursors tends to be proprietary; however, in general it consists of small diameter linear molecules that are made up from nitrogen, hydrogen and carbon (Fig. 9.2). Spinning tends to orient these molecules parallel to the

Pitch precursor

Polyacrylonitrile

I

Particulate removal

I

Pretreatment (chemical, thermal, mechanical)

I

Wet or dry spinning

I

Fiber spinning

Solvent extraction

I

Stabilization at 200-260°C

I

Carbonization to 1500°C

I

Graphitization to 2500°C Fig. 9.1 Schematic illustration of process to produce carbon fibers from polyacrylonitrile and pitch.

172 Carbon fibers

\

CH

CH

CH

CH

C

C

C

C

C

n

NN

/CH2\

a

NN

/CH2\

/

,CH2\ CH

a

/CH2\

a

n

NN

NN

NN

Fig. 9.2 Structure of the ideal PAN molecule (Eggs, 1976).

fiber axis, but they continue to be randomly oriented transverse to that direction. Thus, a fiber with a twisted fibrillar structure is produced. This fiber is supplied by the textile industry to the carbon fiber fabricator who stabilizes it under tension before converting it into a carbon fiber using a controlled heat treatment process. Apart from the obvious generation of the fibrous shape, the importance of the spinning process to the fabrication Initiation

of carbon fibers from PAN is relatively minor; more important is the chemical makeup of the PAN and the presence of small amounts of other constituents that influence the complex chemical reactions that occur during stabilization and carbonization. Stabilization involves cyclization of the oriented molecules and results in the release of most of the hydrogen and part of the nitrogen as NH, and other nitrogen compounds (Fig. 9.3). The role of the Cyclization

Abnormal structures end structures impurities N

H Abstraction

n = 0, 1 , 2 ... m = 0,1,2.

2. Transfer 1.Cyclization reactions

scission

I \

N intra- or intermolecular transfer

I

re-initiation

Fig. 9.3 Release of HCN and NH, during cyclization of PAN fiber (Grassia and McGuchan, 1971b).

Pitch precursor treatments 173 retained nitrogen is very important to both the crosslinking process and to the development of optimum properties during carbonization. 9.2.2 PITCH

A pitch precursor taken from a petroleum or coal tar feedstock initially contains individual molecules that exhibit appreciably different molecular weights. These untreated precursors have been used to produce fibers; however, they are isotropic and exhibit relatively poor mechanical and physical properties. Conversely, the carbon fiber producer can pretreat the pitch to develop a continuous anisotropic phase (similar to a mesophasic liquid crystal) or a two phase mixture that becomes highly oriented during the subsequent spinning process. In contrast to the chemical changes occurring in PAN, physical changes are responsible for the final properties of pitch-based carbon fibers. Essentially, the precursor isotropic pitch is pretreated to produce a two phase mixture that is predominantly anisotropic. During spinning and drawdown, this mixture is very strongly oriented both parallel and transverse to the fiber axis. Oxygen added during stabilization tends to crosslink these large molecules in a simple way before being released on carbonization as CO, CO, and H,O. More recent interest has centered on producing fibers from synthetic pitches. These require no extensive pretreatment and they stabilize faster at a given temperature. Fibers made from these pitches should cost less to produce. Precursor materials that have been used to produce carbon fibers include: polyamides, polyesters, polyvinyl alcohol, polyvinylidene chloride, poly-p-phenylene, phenolic, napthalene, naphthalene-phenanthrene, alkylbenzenes, rayon, polyacrylonitrile (PAN) and various petroleum, coal tar and synthetic pitches (Ezekiel, 1969; US Patent 3 533 741, 1970; Shindo, Nakanishi and Soma, 1969; Boncher, Cooper and Everett, 1970; French Patent 1535 800, 1968; Kawamura and Jenkins, 1970;

Mochida et al., 1988; Lewis and Nazem, 1987a). Today, fibers are produced commercially from rayon, PAN and the various pitches. However, the process to produce fibers from rayon is very expensive because it involves stretching at very high temperatures and the yield of carbon after carbonization is small. Rayon-based fibers are therefore fabricated in such small amounts that they are really of no commercial importance. Additional information can be obtained from: Yanagida et al., 1991; Bacon, 1973; US Patent 3107152, 1963; Yoneshoga and Teranishi,l970. 9.3 PITCH PRECURSOR TREATMENTS

A typical pitch precursor material is obtained from either the distillation products involved in the chemical treatment of decanted oil or as a by-product of the production of metallurgical coke from coal. In a conventional as-received pitch, basic structural units (BSUs) are already present to a degree that pitch can be considered a random suspension of highly aromatic molecules similar to coronene (molecular weights between 600 and 900) surrounded by a liquid of smaller molecular weight. Heating such a system initially reduces the viscosity. Eventually, however, the viscosity increases as the BSUs grow and coalesce to form larger entities (molecular weights between 1000 and 4000) with a specific long range anisotropy called 'mesophase' (Brooks and Taylor, 1965). Continued heating eventually causes an inversion in which the mesophase spheres become the continuous anisotropic phase within which are suspended spheres of the isotropic low molecular weight material. The rate of viscosity increase is very slow at low temperatures, but accelerates as the temperature is increased. At temperatures greater than 350°C, the pitch begins to form coke by a process of thermal degradation and gas evolution. The variation in viscosity with temperature for numerous pitch fractions has been documented (Bathia, Fitzer and Kompalik, 1984) and is shown in Fig. 9.4. As

174 Carbon fibers

60

I-

lS

CTP A240

1:1

1:4

I I I

45

I

-

I

I I I

v)

B

>. c .v) 0

I

I I

30-

.-> v)

I

U

I

C

EQ

P

I 1s-

\

50

100

150

200

250

300

350

b 00

1so

500

550

Temperature OC Fig. 9.4 The variations of apparent viscositv w ~ t l temperature i o f T ariotii pitch fractions (Batha, Fitzer and Kompalik, 1984)

can be noted, at a given temperature, the viscosity of any pitch is greater the more anisotropic phase it contains. By cooling after partial reaction, it is possible to produce a two phase pitch with a viscosity appropriate for spinning or infiltration. A number of experiments have been carried out designed to accelerate the process of producing two phase pitch mixtures. These involved heating, stirring, bubbling an inert gas through the liquid (sparging), or the combination of stirring and sparging (US Patent 3 629 379, 1971; US Patent 3 919 383, 1975; US Patent 3 974 264, 1976; US Patent 4017327, 1977). In this way, an appropriate anisotropic concentration can be produced in much shorter time periods (hours instead of days) with improved spinning characteristics. This latter characteristic seeins to be associated with the smaller molecular size existing within the two phase mixtures formed by stirring and sparging and the smaller variation of molecular size.

A 1980 patent (US Patent 4 208 267, 1980) discussed a different method of producing isotropic/anisotropic pitch mixtures. The method essentially consists of dissolving part of the original isotropic pitch in an organic solvent such as benzene, toluene, xylene, etc. The material that is insoluble can then be converted, by heating, into a material that is greater than 75% anisotropic. The efficiency of this process is quite poor however, since only a \Tery small amount of this 'neomesophase' can be produced from a given pitch. For instance, using Ashland A260 pitch, about 75-90% of the initial pitch will dissolve. Using Ashland A240,80-905:1 of the pitch disso1v es. Variations of the gas-sparge process fpossibly associated with a chemical fractionation) can be made to change the characteristics of a resultant pitch. Lafdi, Bonnamy and Oberlin (1991a; 1991b; 1991c; 1992) and Lafdi and Oberlin (1994a;1994h) have indicated that

Spinning conditions 175 some pitches exposed to a nitrogen sparge at atmospheric pressure or a hydrogenation treatment at high pressure produce a continuous strongly anisotropic material that contains small particles of a weakly anisotropic material. During spinning, the second phase becomes completely absorbed (or transformed) to produce a uniformly anisotropic fiber. They believe that sparging disturbs the formation of an anisotropic phase that exhibits large differences in molecular weight. Indeed, they suggest that the spheres of anisotropic material contain BSUs that are only weakly associated. In contrast to the strong molecular orientation exhibited by the Brooks and Taylor type of ’mesophase’ (Brooks and Taylor, 1968), the common orientation of this new anisotropic material results from the statistical orientation of small units. The pitch then behaves as a two component gel which exhibits a long range anisotropy in the bulk. The short range change in orientation of the carbon units produces sharp changes in orientation similar to grain boundaries. Such regions produce a zig-zag nanotexture in the resultant carbon fiber that prevents inter-sheet gliding and provides a crack inhibiting function that contributes to relatively high ultimate strength values.

9.4 SPINNING CONDITIONS

9.4.1 POLYACRYLONITRILE

Dry and wet spinning of polyacrylonitrileprecursors have been used. In the dry process the polymer is solubilized and spun into a current of hot air that removes the solvent. Unfortunately, solvent removal tends to be quite rapid and can cause the outer portion of the fiber to solidify before the solvent can diffuse from the fiber’s center. The large diffusion gradient that develops can seriously affect the final shape of the fiber (Edie and Diefendorf, 1993). The more common wet spinning method involves solubilizing the polymer with a polar solvent such as dimethyl

acetamide before extruding it into a ’coagulation’ bath through a spinneret. In the wet-spinning process, the fiber is solidified by using a coagulant (such as ethylene glycol) which extracts the solvent from the polymeric fiber. In a manner similar to the dry spinning process, the rate at which the solvent is extracted from the polymer as it passes through the coagulation bath governs the final shape of the fiber. The temperature, concentration and circulation rate of the fluid in the coagulation bath are known to affect the structure and hence the physical and mechanical properties of as-spun fiber. Many companies have added a supplemental stage to the spinning process that is designed to reduce the water content. This additional step tends to increase the molecular orientation within the fiber (US Patent 3846833, 1975; US Patent 3 841 079,1974). A typical acrylonitrile-based precursor contains several percent of various co-monomers such as methyl acrylate or vinyl acetate which improve the precursor’s spinnability or fabric properties. Though not added to aid carbonization specifically, they have been found to influence the properties of the resulting carbon fiber. Many modified PAN polymers such as acrylonitrile-hydroxyethylene, acrylonitrile-vinyl chloride-itaconic acid (French Patent 2 328 723), polyacrylomidoxium (US Patent 3 767 773,1973) have been investigated to obtain a suitable as-spun fiber capable of producing a fiber with a large carbon yield after carbonization. 9.4.2 PITCH According to Singer (US Patent 3919383 (1975)), in order to spin a fiber, pitch must be heated to produce a viscosity between 10 and 200 poise (1-20 Pa s). However, temperatures greater than about 35OOC cannot be used to obtain the required viscosity because thermal decomposition of the pitch will occur. In addition, spinning should be carried out above a minimum temperature of about 200°C since

176 Carbon fibers The spinning of two phase mixtures is not an easy commercial operation since the phases exhibit different viscosities and densities. The strongly anisotropic continuous phase contains within it a less anisotropic (or isotropic) phase which exists as spheres. If the diameter of the spheres is large relative to the spinning orifice, localized weak sections of extrudent can be produced that can break and make it difficult to maintain a continuous fiber thread. In addition, since the stabilization rates of each phase differ, one phase may be over stabilized relative to the other and it becomes difficult to

this determines the maximum temperature that can be used in the subsequent stabilization step. These temperature requirements define a processing window into which suitable pitches must fall. Using literature data (Lewis and Nazem, 1987a;Mochida et al., 1988; Yanagida et al., 1991), White (1992) has shown that the smallest window exists for 100% anisotropic pitches processed from coal or petroleum. A larger window exists for material partially transformed from the same precursors; however, as shown in Fig. 9.5, the largest window exists for synthetic pitches.

Temperature ("C) 450

350

400

250

300

200

Window Boundaries

4 c

h

In (d

4

> + .-

11

Sol poi

I

In

8 In

5

h

In lU" (d

4

>

c .-

lo2

In

0 0 In

5 A 7/////

B

\

I

1.4

l

l

1.5

l

\ Lowest viscosity pitch

TC I

l

1.6

l

1'"'

that can be stabilized

-

High temperature limit, to avoid pitch decomposition t

1.7

l

l

1.8

l

l

1.9

l

l

2.0

l

l

2.1

,

10-3 2.2

Inverse Temperature, I/Tx lo3 (K-l) Fig. 9.5 Processing window for injection of mesophase pitch (White, 1992). A A : The lowest viscosity reported for mesophase pitch prepared from petroleum or coal-tar pitch by pyrolysis to 100% mesophase. BB': The lowest viscosity reported for a petroleum- or coal-tar-based mesophase pitch only partially transformed but with the mesophase acting as the continuous phase. CC': The lowest viscosity reported for a chemically-derived fully-transformed mesophase pitch.

Stabilization of polyacrylonitrile generate optimum properties in the resulting carbon fiber. This fact seems to have been recognized early, since Singer (US Patent 3 919 383, 1975) describes methods to remove the isotropic component from the two phase fiber immediately after spinning. Due to the proprietary nature of the various procedures, little is known of the spinning process itself, however, a number of patents have been awarded (US Patent 4 331 620,1982; Matsumoto 1985; US Patent 4 504 454, 1985). These have been discussed in other publications (Wright, 1989; Wright and Palmer, 1994); most deal with attempts to combat the two phase nature and non-Newtonian behavior by stirring the pitch in the spinneret. Apparently, agitation of mesophase within the spinneret is part of the technique used by the Petoca Oil Company to randomize the cross sectional microstructure and produce the folded graphene layers that are thought responsible for the relatively high strengths exhibited by their ’carbonic’ carbon fiber (Guigon, Oberlin and DesarmotJ984a; 1984b). As seen previously in Table 9.1, for similar modulus materials, the strength of Petoca’s carbonic fibers (HM-series), while still not as good as PAN based fibers (T-series),is appreciably better than other commercially available pitch material (P-series). The flow of pitch through the spinneret (Matsumoto, 1985), the geometry of the die hole (Singer, 1978; Yamada et al., 1984; US Patent 4 504 454, 1985), pitch viscosity, die swell and drawdown are all known to affect the microstructure of the resulting pitch-based fiber. Indeed, the above authors illustrate their comments with fiber microstructures that vary from circumferential through random, radial or radial with a crack (Fig. 9.6). Spinning experiments using 100% ’mesophasic’ synthetic pitches produced from napthalene (Ohtsuka, 1988), napthalene-phenanthrene (Lewis and Nazem, 198%) and alkylbenzenes (Yanagida et al., 1991) have been reported. The viscosity of these materials is appropriate for spinning

177

Fig. 9.6 Typical pitch-based carbon fiber microstructures (Edie, 1990): (a) radial; (b) onionskin; (c) random; (d) flat-layer; (e) radial-folded;(f) radial with wedge.

at temperatures lower than those necessary to spin anisotropic-isotropic mixtures obtained from petroleum or coal tar pitch; thermal decomposition during spinning is therefore not a problem. Stabilization of such pitches must be performed at correspondingly low temperatures; however, when they are compared to petroleum and coal tar mesophasic pitches, stabilization at a given temperature occurs at a significantly faster rate. This feature, when combined with the fact that no pitch pretreatment is necessary, should enable fibers to be produced at significantly lower cost. The result of all spinning processes is a PAN or pitch fiber that can be changed into a carbon fiber by heating at a very slow rate to a temperature of about 1500°C. A higher rate of heating will usually melt the fiber unless it is first converted from a thermoplastic to a thermoset. This conversion or stabilization treatment is done by heating to a relatively low temperature for an extensive time period in an atmosphere containing oxygen (usually air). 9.5 STABILIZATION OF POLYACRYLONITRILE

This process converts the thermoplastic, as-spun polymer into a thermoset that is capable of maintaining its shape during

178 Carbon fibers carbonization (Lavin, 1992; Yooh, Korai and Mochida, 1994).The operation is identical for both PAN and pitch based fibers. However, the stabilization of anisotropic pitches involves simple cross linking of plate-like molecules whereas the stabilization of PAN involves many different chemical reactions. Stabilization of both PAN and pitch is an exothermic process, so great care must be taken to control the rate of reaction and to avoid thermal runaway which melts the fiber and is a fire hazard. Commercial stabilization is carried out by heating the PAN fiber in air between 200 and

CYCL’ZATloN

I

CN

260°C for a period of time that varies between thirty minutes and several hours. During stabilization, several interdependent chemical reactions occur. The reaction that dominates is primarily determined by the chemical composition of the initial precursor, the spinning history, the final composition of the as-spun fiber and the stabilization heating schedule. A PAN polymer mainly consists of acrylonitrile entities -CH,CH(CN)- which are able to cyclize (Johnsonet al., 1972)with the help of an initiator into a presumably linear ’ladder polymer’ similar to that shown in Fig. 9.7. In general, the pendant nitrile groups of PAN

CN CN PAN

Fig. 9.7 The process of PAN stabilization and subsequent carbonization (Fitzer and Heine, 1988).

Stabilization of polyacrylonitrile first become crosslinked to form a ladder polymer. Initiation of this process is catalyzed in some cases by the presence of a small amount of reactive copolymer such as itaconic acid. Oxygen is then incorporated into the ladder polymer under a number of possible schemes which have been described by Watt and Johnson (1975) and by Clarke and Bailey (1973) and are shown in Fig. 9.8. Cyclization and stabilization induce tremendous shrinkage into the polymer. Longitudinal shrinkage is resisted mechanically; however, the diameter of the fiber is allowed to decrease.

A balance should be kept during stabilization as to hydrogenation degree. A large hydrogen content can result in a small N/C ratio which increases the temperature at which the local molecular ordering occurs. Conversely, increasing the available oxygen decreases the size of and the temperature at which the units of local molecular order (LMOs)are formed. In addition, since the viscosity increases as crosslinking (stabilization) proceeds, the mobility and growth rate of the LMOs decrease; hence their final size remains small. The smaller the size of the LMOs, the

la

c

c

c

c

N

N

N

N

Ill

H

Ill

Ill

H

H

H

H

I

I

I

I

I

l

l

Ill

H

H

H /:\> % ;\/;\;/:

H

'LL,

1

/c\N/c\N/c\N/c\N

+ + + +

0

0

179

0

/

IV

0

Fig. 9.8 Incorporation of oxygen into cyclized PAN fiber (Clarke and Bailey, 1973).

180 Cavbonfibers less graphitizable is the carbon and the lower the properties of the fiber. In order to ensure appropriate N/C ratios at reasonable temperatures, stabilization should result in only a moderate degree of crosslinking. In addition, slow heating rates during precarbonization should permit hydrogen and delay nitrogen emissions; both of these effects lower the temperature at which extensive formation and growth of LMOs occur. A commercially acceptable rate of stabilization requires the use of as high a temperature as possible. However, since the reactions that occur during stabilization are exothermic, it is most important to limit the oxidation rate and to prevent uncontrollable temperature increases. These conflicting requirements have resulted in the development of alternative methods of stabilization. These include stabilization in: hydroxylamine solution (US Patent 3 767 773, 1973), aminophenoquinones (US Patent 4 004 053, 1976), aminosiloxanes (US Patent 4 009 248,1977), amine salts (US Patent 4 009 248, 1977; US Patent 4 024 227, 1978; US Patent 4 031 288, 1978) or stabilization in gas phases such as mixtures of NO and NO,, Br, and 0,, or HC1 and 0,. Other stabilization processes have been proposed that are designed to reduce the cost and/or decrease the stabilization time. It seems possible to reduce the time by stabilizing the fibers in persulphate (US Patent 3 650 668, 1972), cobalt salt (US Patent 3 656 882,1972), nitric acid (US Patent 3814377, 1974; US Patent 3 656 883, 1972), or to control the final quality of the fiber by stabilizing in carboxylic acid (US Patent 3 814 377, 1974; US Patent 3 656 883,1972), or nitrophenol. Processes designed to produce fibers of larger modulus have also been developed. These involve stretching the precursor during stabilization (US Patent 3 917 776,1975; US Patent 3 677 705,1976). Fiber manufacturers attempt to fit the physico-chemical conditions of the various operations cited above (nature and proportion of the co-monomers, cyclization, stretching, stabilization, carbonization) to obtain an opti-

mal product, i.e. forming the LMOs at the smallest reasonable temperature, retaining the largest nitrogen content beyond the temperature of LMO formation and incorporating an optimum amount of oxygen during stabilization to prevent polymer melting without inducing too small a LMO size. It is known that the overall oxygen content should be between 8 and 12 wt.% in order to completely stabilize PAN fibers (US Patent 4069 297, 1978). Less than 8 wt.% oxygen gives a large weight loss on carbonization due to excessive evolution of gases from the incompletely stabilized central core; more than 12 wt.% oxygen degrades surface layers and the properties of the final fiber (Johnson, Rose and Scott, 1970). Exactly what an average value of 8 wt.% translates into for the specific oxygen content of the surface layers and the core region is unknown, but it would obviously depend on fiber diameter and the kinetics of the stabilization process (diffusion or reaction controlled). A large diameter PAN fiber containing an average of 8 wt.% oxygen exhibiting diffusion controlled stabilization kinetics would probably be composed of highly degraded surface layers with perhaps an under-stabilized central core. Conversely, a very thin fiber exhibiting reaction controlled stabilization kinetics might be completely and homogeneously stabilized with an average oxygen content of less than 8wt.%. Presumably a similar statement can be made for the stabilization of pitch fibers. In the stabilization process, the effect of fiber diameter on the rate of oxygen uptake is important. The three curves shown in Fig. 9.9 illustrate the slower rate of oxygen uptake exhibited by fatter fibers. In addition, larger diameter pitch fibers tend to exhibit a diffusion controlled stabilization that produces an under-stabilized central core and an over-stabilized fiber skin. Smaller diameter pitch fibers appear to exhibit reaction controlled stabilization since no skincore type microstructures are observed.

Chemical changes during carbonization 181

....................

________---------

Solid line - mesophasic particles

-4.881

-5.08 8.68

( O x i d i z e d Oxygen, "

"

I

'

"

'

I

"

"

388 Deg I

'

Time (hours)

Fig. 9.9 Effect of diameter on the stabilization of mesophasic particles and pitch-based fibers (Kowbel, Wapner, Wright, 1989). 9.6 CHEMICAL CHANGES DURING CARBONIZATION

The carbonization of stabilized PAN and pitch involves controlled heating to a temperature of about 1500°C.The majority of gases emitted from either the PAN or the pitch are emitted before a temperature of 1000°C is reached and the emission is primarily from unstabilized regions (Jain and Alhiraman, 1987; Lewis, 1982). Indeed, the quantity of gases emitted from an unstabilized central core of either PAN or pitch can be so large that the fiber can disintegrate. Great care should therefore be taken to determine the optimum heating rate for stabilized or under-stabilized fibers. In some cases, hold times should be incorporated into the heating cycle. Both materials emit a variety of gas molecules containing oxygen, hydrogen and carbon; however, a major difference between PAN and pitch involves nitrogen containing compounds which are only emitted from PAN. The temperature and

rate of emittance are important control parameters since they affect the strength of the resultant carbon fibers. A stabilized polyacrylonitrile fiber which contains about 11wt.% oxygen can be thermally degraded by heating at a slow heating rate (Riggs, Shuford and Lewis, 1982) in an inert atmosphere such as nitrogen or a reactive environment where nitrogen gas is bubbled through acid (US Patent 3 972 984, 1976) or water (US Patent 3 677 705,1976; US 3 656 903, 1972; US Patent 4 039 341, 1976).As the temperature increases, many complex reactions take place resulting in the evolution of volatile products. For example, when the fiber is initially heated, cyclization occurs with the release of large amounts of HCN and NH,. Up to 450°C, HCN, acrylonitrile, propionitrile, NH, and H,O are emitted. Subsequently, at around 500°C and 700°C copious quantities of HCN and water vapor are emitted, respectively. All of these emissions are believed to

182 Carbon fibers come from reactions involving crosslinking of individual molecules. Evolution of nitrogen starts near 700°C; so fibers produced after being heated to 1000°C retain only about 5.8 wt.% nitrogen and have lost about 50 wt.% of the original PAN precursor fiber. Results obtained from experiments involving slow pyrolysis at 4"C/min indicate that optimum mechanical and physical properties are unobtainable unless high nitrogen contents are retained within the precursor until the later stages of carbonization (Deurberque, 1990; Deurberque and Oberlin, 1991). Therefore, a large nitrogen content (large N/C ratio) should be present when local molecular ordering (LMO)begins and the carbon skeleton is being formed. Since the N/C atomic ratio depends inversely upon the H/C ratio, LMO should occur at large N/C and small H/C ratios. Essentially, small amounts of aromatic hydrogen and a relatively large amount of nitrogen present during the LMO stage allow the carbon skeleton to remain flexible enough at high temperatures that molecular rearrangement is easy. Within this overall fibrous texture, the nanotextural features of the carbonized fibers are the consequence of the variations in cyclization, stabilization, carbonization and graphitization conditions. If the original precursor is CH- and NH-rich but oxygen-poor, the corresponding carbonized fibers will have low porosity, high compactibility and stacking order and a relatively high strength. Likewise, if two nitrogen atoms are present in two aromatic rings contained within adjacent sheets, they are able to promote bonding of contacting BSUs together with a N, release (Watt, 1972). The ultimate strength value of the fibers increases as the compactibility and the availability of 'efficient' nitrogen (i.e. the nitrogen remaining at the moment of LMO occurrence) increases (Oberlin and Guigon, 1988; Guigon, 1985). Bright and Singer (1979) agreed with others when they found that the tensile strength of most heat treated fibers tends to decrease with higher temperature exposures and release of nitrogen (Fig. 9.10). The same

40 .1-

3.5 -

100

3.0-

.-

c

0-

too

z

G2

2.5-

W

W

E

a

t;; W

z w

n

I

I

f

= I)

g (3

2.0-

300

W

$ 2 W

t-

k

1.5 -

1.0

zoo

-

Fig. 9.10 Effect of heat treatment temperature on the tensile strength of carbon fibers (Bright and Singer, 1979).

authors argue that if adequate nitrogen exists within the fiber after LMO occurs, then a greater potential for bonding exists and an improved rather than a diminished strength will result. This conclusion is based on the observation that, in contrast to other fibers, the strength of (nitrogen containing) commercially available Toray T-300 fibers increased from 2.2 to 3.2GPa after heat treatment to 2800°C. Tension also increases cyclization and nitrogen elimination which increases the tensile strength of the final carbon fiber (Watt, 1972). Exactly what influence nitrogen has on the development of high tensile strength is still being debated. However, the texture of the graphitic layers, the amount of nitrogen present in the original precursor and the temperature at which misshapen layers touch -

Microstructural changes during carbonization 183 and hence are available to bond and emit nitrogen - are obviously important. Up to lOOO"C, carbonization leads to an effluentloss and increasing aromatization. As a result, the solid residue transforms from a viscoelastic into a brittle solid material. The stabilized PAN normally carbonizes into a statistically isotropic but nanoporous carbon material which, because of the small dimensions of the initial LMOs, is inherently non-graphitizable (Joseph and Oberlin, 1983a). Continuing pyrolysis up to 1500°C eliminates most of the residual nitrogen and completes the conversion of the PAN molecules into sheets of carbon that are appreciably anisotropic (Mair and Mansfield, 1987).Continued heating eliminates the remaining nitrogen and, since the material is then only made of pure carbon, further modifications are only structural. Graphitization is the name of the process that involves heating the carbonized fiber to approximately 2500°C in times as short as a minute (US Patent 4 005 183,1977).Graphitized pitch fibers exhibit a larger, more graphitic and better oriented crystal structure than PAN-based carbon fibers which are inherently non-graphitizable. Parallel to the fiber axes, pitch fibers have higher stiffness and thermal conductivity values and a reduced thermal expansion coefficient. These changes due to graphitization do not produce any significant increase in relative strength values. As a result of the extreme temperatures required to process them, graphitized pitch-based carbon fibers are more expensive and are fabricated for specialized applications.

Fig. 9.11 Model of crumpled sheet-like structure (Guigon,Oberlin and Desarmot, 1984a).

temperatures, for the bonding of adjacent sheets. These sheets contain numerous vacancy imperfections and are folded to enclose pencil-shaped voids oriented in the general direction of the fiber axis. The lengths of each block or sheet are relatively short, with each 9.7 MICROSTRUCTURAL CHANGES DURING succeedingblock misoriented with respect to its CARBONIZATION neighbor. A schematic illustration of the The initial heating of stabilized PAN fibers microstructure as it exists within the actual causes growth of graphite-like ribbons by a fiber is shown in Fig. 9.12. This structure is typdehydrogenation mechanism. Denitrogenation, ically exhibited by high strength PAN-based which occurs as the temperature is increased, is carbon fibers. Further temperature increases tend to responsible for the growth in area and the transformation of these ribbons into thin sheet- decrease the void space by joining sequentially like structures (Fig. 9.11) and, at higher oriented and touching graphite like layers and

184 Carbon fibers

fibre axis

Fig. 9.12 Model of microstructure within a high strength PAN-based carbon fiber (Guigon, Oberlin

and Desarmot, 1984a). aligning them more parallel to the fiber axis. The distorted sheets of BSUs associated with tilt and twist boundaries are bonded to each other wherever the boundaries of adjacent sheets touch. The more compact the fiber, the larger the number of contact areas and the Fig. 9.13 Model of PAN-based high modulus cargreater the chance for adjacent sheets to bond. bon fiber (Guigon,Oberlin and Desarmot, 1984b). The lateral cohesion thus formed causes the strength of the fiber to increase. Specifically, the distortions within the polyaromatic centrations and, hence, weaker fibers graphene layers or sheets tend to be removed (Reynolds and Sharp, 1974; Reynolds and by accumulating any structural defects at their Moreton, 1980). boundaries. This induces a progressive increase of the width and the radius of curva9.8 ELECTRICAL AND THERMAL ture of the aromatic layers which can be PROPERTIES correlated with the stiffness, stacking order and the diameter of the graphitic layers Studies of microstructural features have been (Oberlin, 1984).A schematic of this microstruc- carried out using techniques that include: ture is shown in Fig. 9.13. The longer, better X-ray and electron diffraction, electron spin oriented and more graphitic microstructures resonance, thermoelectric power, magneto exhibit both higher values of moduIus and resistance, lattice fringe imaging, etc. All thermal and electrical conductivities; unfortu- results indicated that carbon fibers are comnately, the misalignment of the larger posed of turbostratic layers of graphite microstructural units causes large stress con- oriented preferentially at some angle to the

Electrical and thermal properties 185 fiber axis. Increasing the heat treatment temperature results in a reduction of the interlayer spacing, a decrease in void space, a growth in thickness and area of the graphitic crystallites and an increase in the preferred orientation of the microstructure. All of these changes increase the elastic modulus and the electrical and thermal conductance. A corresponding reduction of the tensile strength also occurs by mechanisms that depend on local defects as discussed in the previous section. A comparison of the g-value anisotropy of pitch and PAN in Fig. 9.14 indicates that the degree of anisotropy changes for both fibers after heating to about 1700°C. Although pitch based fibers become more anisotropic when the temperature is increased further, the anisotropicity of PAN seems to saturate at a level which is comparable to that of a pitch fiber heated only to about 2000°C. A simple consequence of this inability to fully graphitize

x104

140

-

t

Long Heat Treatment

:120100-

"

\

I

/

1

I

/ I /

0)

dl

PAN-based fibers is the lower maximum values of the elastic modulus and the electrical and thermal conductivities. Typical results which compare the effect of heat treatment on the electrical and thermal conductivities are shown in Figs. 9.15 and 9.16.

/

1800 2200 2600 3000 Heat Treatment Temperature, deg C

Fig. 9.14 Variation in g-value anisotropy of pitch-base and PAN-base carbon fibers as a function of heat-treatment temperature (HTT) (Aggarival, 1977).

Single Crystal Graphite ----------------

1000

1500

2000

2500

3000

35 0

HEAT TREATMENT TEMPERATURE (OC)

Fig. 9.15 Schematic variation of the room temperature electrical resistivity against T,, for: (0) ex-rayon; ( 0 )hot stretched rayon; (A and V) ex-PAN fibers (Robson et al. 1972); (0) Ex-pitch fibers (Bright and Singer, 1979)and (solid curve) benzene-derived fibers (Chieu et al. 1983). The scatter of typical data points about the mean give an indication of the uncertainty. The dashed line indicates the decrease in resistivity produced by hot stretching the ex-rayon fibers.

186 Carbon fibers E. GPa

l

tiI '

0.01 1

'

20

I

I

30

40

I u)

Ex

I

I

I

I

60

70

80

90

1

106 psi

Fig. 9.16 Conductivity of pitch-base, PAN-base and rayon-base carbon fibers as a function of the tensile modulus of elasticity. (Courtesyof R. Gray, NSWC, Dahlgren, Virginia.)

9.9 MECHANICAL PROPERTIES OF FIBERS

The microstructural changes discussed above have been deduced using X-ray diffraction techniques. In addition, the mean length of the graphite sheets oriented in the fiber direction La and their thickness Lc may be computed using such techniques. It has been found that both of these parameters increase with increasing temperature. The orientation of these layers also becomes increasingly aligned with the fiber axis. The net effect is to increase the tensile modulus continuously as shown in Fig. 9.17. Conversely, the tensile strength tends to decrease (Fig. 9.10). More recent studies, which allow direct

observation, have used electron diffraction (Guigon, 1985) and lattice fringe imaging (Oberlin and Guigon, 1988). This latter technique is particularly powerful, because in dark field, the parallel orientations (with respect to the imaging beam) of the convoluted graphite layers can be imaged. Sketches of the possible layer convolutions, the images they can produce and the measurements that can be made are shown in Fig. 9.18. Most important is the realization that it may be possible to classify the properties of commercial samples in terms of decreasing either transverse (rt) or longitudinal (r,) radius of curvature of the sheets. Such a classification, based on these and other numerical values

Mechanical properfies offibers 187

I

1000

2000

3000

shown in Table 9.2, could form the basis of empirical relationships between micro-texture and mechanical properties. At increasingly higher heat treatment temperatures, the scattering domains within high modulus fibers become large and well defined so that the length of the graphitic sheets in the fiber direction, .La,,Cr can be measured directly (see Table 9.2) from the observed Moire fringes. Correspondingly, the radii r, and rI of the sheets are also measurable from 002 lattice fringes. The lateral cohesion of the fiber is also ensured by bonding between adjacent distorted sheets of carbon wherever two grain boundaries are in contact. The chances of such bonding increase as r, and rI decrease, but decrease as La increases. Hence the extent of lateral cohesion can be defined by a variable S = La [(l/rf) (l/rt + l/rl)]. As shown in Fig. 9.19, a linear correlation has been observed between CY, and 1/ S indicating that long, relatively unbonded graphitic layers result in weak fibers. Conversely, Young's modulus and electrical

1 3500

THT ( ' C )

Fig. 9.17 Effect of heat treatment temperature on the elastic modulus of PAN and pitch-based carbon fibers. (Data from Johnson, 1969 and Aggarival, 1977). 1

m 2rr sin 37.

3

2

n

m

n

2r,sin37*

Fig. 9.18 Sketches of the possible dark-field images of a longitudinal section (lamellar model of Fig. 9.15) as r decreases and the fold develops (Guigon, Oberlin and Desarmot, 1984b).

188 Carbon fibers Table 9.2 Quantitative measurements which are suggested for use to classify microtexture and mechanical properties SAD patterns 002DF ~ _ _

[O]

La

__

N L,

Number of fringes in a stack Length of a perfect fringe Length of a distorted fringe

L?

__

11DF

__

Thickness of elementary bright domain [BSU] Length of elementary bright domain

Lc

002LF __

Half width of 004 reflections Half width of 11[0] ring

L'coo4

L'd,

Length of turbostratic Moire fringes Diameter of a domain showing rotational Moire fringes

Lall TF Lall Cr

~-

P

ODP of 002LF

Arc opening Interfringe spacing spreading

ADO,,

,,

lt +

0

I

1

1

I

"

'

I

I

I

1

0.1

1

'

f

i

I

I

'

I

I

I 0.2

11s

AA-

Fig. 9.19 Numerical relations between tensile strength and the microtexture oc = f(l/S). High modulus fibers (full line). High tensile strength fibers heat treated at 2800°C (dashed line) (Oberlin and Guigon, 1988).

conductivity correlate well with La, as seen in Fig. 9.20. 9.9.1 MICROSTRUCTURAL CONSIDERATIONS

According to the pioneering work of Griffith (1920), the following expression describes the strength, oF,of a brittle solid, containing a crack of length 2 4

oF=d(Ey/4a) (9.1) Carbon is a brittle solid; thus, since no plastic deformation can occur, very high local stresses where E is the elastic modulus and y is the surwill develop at stress raisers, such as disconti- face energy. Inspection of equation (9.1) nuities, changes of section size, cracks, etc. indicates that longer cracks are more effective

Mechanical properties offibers

189

(a)

I

I

longer cracks generate higher tip stresses, it can be inferred that once a crack begins to move it will continue to move (accelerate) until it reaches the geometric boundaries of the material. As a consequence of this, the failure of brittle solids is abrupt and depends on the probability that a crack of some critical length is present. If such a theory can be applied to carbon fibers, then it can be argued that the maximum length of cracks (or similar microscopic stress raisers) that can be contained is limited by the fiber diameter; so small diameter fibers will be stronger than large diameter fibers. In addition, since the chance of a crack being present is greater, long fibers will tend to be weaker than short fibers.

I

I

iil

In a composite, many fibers are arranged more or less parallel to one another and function as a load bearing bundle. A number of publications have appeared (Herring, 1966; Wright and Iannuzzi, 1973; Wright and Wills, 1974) that discuss the distribution of strengths exhibited by brittle fibers and how these distributions can be used to compute a mean strength, om and a corresponding bundle strength, ob. For example, the above authors argued that their individual fiber strength data tended to obey a Weibull distribution characterized by the expression, G(o) = 1- exp{- a(o/oo)")

(9.2)

where G(o) is the probability of failure of a fiber subjected to stress o,oois the distribution

190 Carbon fibers scale factor, w is the distribution shape factor and a is a function of the length/diameter (L/d) ratio of the fiber (Corten, 1967). If many fibers are tested of different length I, then a graphical method can be used to deduce w, a and a. and, the mean strength and the strength of a bundle of fibers can be computed from

om=oo(i/d)-lw(i + i/c~),

(9.3)

where r is the gamma function and Gb=oo(acl)e)-'lY

(9.4)

Bundles of twisted fibers would exhibit lower strengths. 9.10 COMPOSITES FABRICATED FROM

CARBON FIBERS

tion of the same fiber by a shear process. The length of matrix required to do this defines a bundle of short fibers (or segment of composite) which must break in order to break the composite. Such bundles can be modeled as analogous to a link within a chain; failure of the weakest link defines the failure load of the composite. Nevertheless, since shorter fiber bundles are stronger than longer, shorter links are stronger than longer links. The link length, 6, has been discussed by Rosen (1964) and, for purposes of this discussion can be approximated by,

6=~ d / 2 ~

(9.5)

where o = obthe stress in the fibers at failure of the composite, d is the diameter of the fiber and z is the shear strength of the matrix or matrix-fiber bond as it exists in the composite. For a given fiber strength distribution, the stronger composites will all exhibit smaller ineffective (link) lengths. This is accomplished by using small diameter fibers, well bonded using high strength glue.

Carbon fibers are very strong, stiff and lightweight materials. In addition, their small diameter (8-12 p)makes them extremely flexible. Unfortunately, they exhibit little compressive strength and they exhibit a poor abrasion resistance. A solution to these problems, and to the problem of brittleness, is to 9.10.1 SURFACE TREATMENT bond large numbers of fibers together to form a composite solid. In this case, the glue or A freshly prepared fiber does not bond well to bonding agent forms a continuous phase that a polymeric glue (or to anything else for that is usually defined as the matrix. The matter); however, the tendency to bond can be matrix-fiber mixture is called a composite significantly increased by subjecting the fiber material. The function of the matrix is to sup- surface to a controlled oxidation. As discussed port and separate the fibers, to protect them by Eggs, Shuford and Lewis (1982) in the earfrom reaction with the environment and to lier edition of this handbook, this treatment transfer load. In a composite, the tensile and essentially etches the surface, cleans it, compression properties parallel to the fibers increases its surface area and produces polar are much better than those measured on bun- hydrophilic oxygen-containing groups which dles. The transverse properties are also bond to it. The process can be carried out in a optimized, since the matrix serves to improve liquid or gaseous environment; for example, the fiber-matrix connectivity. This function is heating in air or oxygen-nitrogen mixtures, important since it affects the mechanical and CO,, C1, NOz-NO, NH, and plasma-ionized thermal properties in the transverse direction. inert gases oohnson 1969; US Patent 3 754 957, Fortunately, although the strength of individ- 1973a; US Patent 3 723 150, 1973b; British ual fibers exhibits a pronounced size effect, no Patent 1341 161, 1973; U.S Patent 4 374 114, size effect is exhibited by composites. In effect, 1983; US Patent 3 627 466, 1971; US Patent load is transferred around fiber breaks into 3767774, 1973; US Patent 3780255, 1973). adjacent fibers and back into the unbroken sec- Direct wet chemical oxidation has been tried

Mechanical properfies of unidirectional composites 191 using aqueous nitric acid, hypochlorite, chlorate and dichromate in sulfuric acid. Treatments have also been investigated using electrolytes of hypochlorite, ammonium hydroxide, sodium hydroxide and ammonium sulfate (US Patent 3 660 140, 1972; US Patent 3 746 506,1973; US Patent 3 894 884,1975a; US Patent 3 859 187, 197513; US Patent 3 746 450, 1973; US Patent 3989802, 1976; US Patent 3832297, 1974; US Patent 3671422, 1972; British Patent 1371621, 1974; British Patent 2 071 702,1981) . During oxidation, the strong carbon-oxygen complexes which are formed bond tenaciously to the fiber surface and will subsequently react with a matrix resin. In order to preserve this reactivity a thin layer of the final matrix resin is applied to the surface of the fibers as a finish or size. This layer does double duty in both protecting the fiber surface against damage during transportation, further processing and handling and in promoting wetting when the sized fibers are bonded together with the matrix resin. Poor bonding is a sensitive function of the surface morphology, anisotropicity, heterogeneity and the nature of the interphase layer between the fiber and the matrix. For example, it has been found that the greater the degree of graphitization and the better the alignment of the microstructure with respect to the fiber axis, the poorer the fiber will bond. Essentially, the higher modulus carbon fibers will not bond easily in the absence of a surface treatment. In some composites where the failure strain of the matrix is smaller than the failure strain of the fibers (as it is for ceramics or carbon), poor bonding is an asset since the largest fiber-matrix bond strength is not required. Conversely, if the failure strain of the matrix is larger than that of the fibers, a strong bond is desired. The reason for this apparent dichotomy involves the fact that fibers are added to brittle matrices primarily to toughen them; only tough matrices (large failure strain) can be strengthened. The mechanism of toughening depends on the blunting of a running

crack by interaction with low strength fiber-matrix bonds; strengthening depends on load being transferred from the matrix to the fibers through a strong fiber-matrix bond.

9.11 MECHANICAL PROPERTIES OF UNIDIRECTIONALCOMPOSITES

The properties of unidirectionally reinforced composites are strongly orthotropic. Specifically, properties measured parallel to the fibers are quite different from those measured at right angles to them. More importantly perhaps is the sensitivity of property determination with respect to the direction of the measurement. For example, Fig. 9.21 illustrates the change in elastic properties exhibited when a unidirectionally reinforced composite is loaded at some angle to the fiber axis. In this case, it can be observed that the stiffnesses of the composite (Qll,etc.) begin to change significantly when the load is misaligned only fifteen degrees to the fiber axis.

9.11.1 MICROMECHANICS

Using micromechanics, various equations can be used to estimate the properties that might be exhibited by well bonded composites. For example, parallel to the fibers, the modulus (E,), strength (0,) and Poisson's ratio y2are given by;

E , =Vf E, + Vm Em

(9.6)

ol=vfof+ vm0,

(9.7)

'12

= Vfv12

+

'mum

o2z om

(9.8) (9.9)

where V is the volume fraction of fibers (0 or matrix (m) respectively, ofis the strength of a bundle of fibers with length equal to the ineffective length (usually, due to lack of statistical data ofis taken as the mean strength supplied by the fiber processor). Also,

192 Carbonfibers

9.11.2 MACROMECHANICS OF LAMINAE

There are a number of excellent books and monographs that have been written on macro-

I

30

01

[

60

1

60

90

-8

mechanics of composites; thus, the reader is directed to these for a more complete discussion of this subject (Ashton, Halpin and Petit, 1969; Jones, 1975; Tsai and Hahn, 1980; Daniel and Ishai, 1994). When only plane stress conditions exist, (e.g. 03=0 and E, is related to E, and E,), then it is possible to relate stress to strain along the principal axes of an orthotropic lamina,i.e. parallel (1)and perpendicular (2) to the fiber axis:

I

01

30

60

'

90

-6

(42

I

01

I

a

30

60

90

-

8

Fig. 9.21 Transformed, off-axis modulus of T300/5208. The angle is the ply orientation and is positive for counterclockwise rotation (Tsai and Hahn, 1980).

Mechanical properties of unidirectional composites 193

(9.12)

the components of the [Q],, has already been illustrated in Fig. 9.21. (9.15)

or

where

and

(9.13) In addition to the variation of elastic properties, the strength of unidirectionally reinforced composites has been found to be sensitively dependent on the angle of loading. A typical failure curve for tension and compression is shown in Fig. 9.22. This type of curve can be obtained by transforming the applied stresses to directions parallel and perpendicular to the fiber axes and then equating those stresses to the failure strengths actually measured (or computed using micromechanics) along those directions. Failure of a composite is then considered to occur when the transformed stress exceeds the failure stress actually measured in that direction, i.e. when

Inspection of equations (9.12) and (9.13) indicates that parallel and perpendicular to the fiber axes, tensile (compressive)stress produces tensile (compressive) strain and shear stress produces shear strain. There is no coupling between shear stress and tensile (compressive) strain, e.g. Q,, = 0 = Q,, as shown previously in Fig. 9.21 and the shear coupling coefficient which is the ratio of tensile (compressive)stress to shear strain, is zero. The above conclusion is not valid when loads are applied at some angle to the principal axes since appreciable shear coupling can occur; Q,, and Q,, are not zero. This means that tensile (compressive) stresses produce shear strains in addition to the more normal tensile (compressive strains). The stress-strain relationships are then written, (9.14)

where x and y are the orthogonal axes of the composite test specimen that are oriented at some angle to the fiber axes and where [Q] or .xY [SI, are the transformed matrixes of equabons (9.12) and (9.13). The variation with angle of

F,,, = Fl,/(cos28);Fx,= F,, / (sin20); F,, = F6/ (sine c o d ) and

Fxtc = FlC/(~os20); Fxc = F J ( s i n 0 ~ 0 s 0 ) and F, are the measured failure where Fl,(c,,F2t(c) tensile (t) or compressive (c) loads measured parallel (1)or perpendicular (2) to the fiber axis and F, is the measured shear strength. These failure criteria are collectively described as the maximum stress failure criterion. There are other failure criteria such as maximum strain, deviatorial strain energy (Tsai-Hill) and Interactive Tensor Polynomial (Tsai-Wu) that can be used, some of which allow for interaction between the stresses. However, for typical composite structures, the maximum stress criterion give reasonably conservative estimates of expected failure stresses. 9.11.3 MACROMECHANICS OF LAMINATES

In order to eliminate coupling and to reduce the very strong change of mechanical properties

194 Carbonfibers 1.2 1.o a 0.8 a (3 0.6 LL'

160 120

.-

(I)

80 x-

0.4

LX

40

g-

F! = o

O

E

5 -0.2

-40

g

0.2

-0.4 -0.6

m

i5

-80

-0.8

0

10

40 50 60 70 Fiber orientation, 8, deg.

20 30

80

90

Fig. 9.22 Uniaxial strength of off-axis E-glass/epoxy unidirectional lamina as a function of fiber orientation (Daniel and Ishai, 1994).

with the direction of applied load, composites are usually fabricated from multiple layers, each arranged at some angle to their neighbor. The angles that are used can be arranged to bring the fiber axes at some optimum angle to the expected loads; however, there is a requirement for each lamina to be oriented very precisely with respect to one another in order to avoid tensile-shear effects and coupling between in-plane loading and out-of-plane deformation (tensile loads can be made to produce bending and twisting deformation, for example). Various rules of angle-ply materials have been worked out in order to avoid the cross-coupling terms. For example, symmetric laminates exhibit no coupling between in-plane loading and out-of-plane deformation. (Symmetric laminates define a composite in which for every lamina oriented at some angle there is another layer of identical thickness and orientation placed at an equal distance from the mid-plane of the composite.) And, composites which exhibit no shear coupling are 'balanced laminates' (pairs of identical layers oriented at a positive and an equal negative angle with respect to the laminate reference axes).

Of particular interest is the quasi-isotropic composite which is often specified for commercial structures. The in-plane engineering elastic constants of these materials are identical in all directions and there is no shear coupling. Examples of such composites are symmetric arrangements of [0/60/-601 or [0/ +45/90] layers. 9.12 TESTING TECHNIQUES

A knowledge of the behavior of the constituent phases allows the mechanical properties of the resulting composites to be computed. These calculations are made with the use of expressions (9.6)to (9.11).The results of the following tests should provide input to and in some cases confirm these calculations. The following was adapted from the excellent (and much more extensive) discussion contained in Chapter 8 of the book Engineering Mechanics of Composite Materials (Daniel and 0.Ishai, 1994). 9.12.1 PROPERTIES OF FIBERS AND MATRICES

Determination of the elastic and failure properties of the fibers is described in ASTM

Testing techniques 195 specification D3379. The difficulty is mostly verse Young’s modulus, tensile strengths and involved in determining the elastic displace- strains and the major and minor Poisson’s ment parallel to the fiber axis, since no ratio can all be determined from this type of measurement device can be attached directly coupon test specimen. Similar properties can to the fragile fiber. Since the fiber tends to shat- also be measured in compression; usually ter and disintegrate when failure occurs, the however, short, thick specimens are used in mean diameter measured before the test has to order to avoid buckling failure. ASTM D-3410 be used to calculate the failure stress. describes this test method. The method involves attaching the fiber across a slot cut into a paper tab. The composFlexure testing ite specimen is aligned coincident with the load axis of the testing machine, the tab ends A far more expensive test is described in ASTM are gripped and the sides of the tab are cut to C393. This requires a rather large allow only the fiber to be loaded by the (22 in/l in/1.5 in) sandwich flexure specimen machine. The measured compliance Cm is the which is tested in four point bending. The sum of the compliance of the loading system dimension of the honeycomb core and the composite face sheets are adjusted to cause failC, and the compliance of the fiber C,. Thus ure in the approximate face sheet. Good results Cm= C,+ C, = C, + (AE,) can be obtained in both tension and compreswhere 1 is the fiber length and E , is the fiber sion; however, since failure might occur in the modulus. A plot of measured compliance core or in either of the outer skins or at the against fiber length allows calculation of the skin/core interface, care should be exercised in loading system compliance and the fiber both determining the exact failure mode and in reporting the appropriate failure stresses. modulus. In the above paragraphs the reader has Polymer matrices are evaluated using coupons cut from thin sheets. Typical geome- been cautioned about making sure that the tries are described in ASTM specifications loading axis of the testing machine and the symmetry axis of the specimen are coincident. D638, D638 and D882. Indeed, consistent measurements will only be achieved by eliminating any tendency to develop complex stresses in any region of the 9.12.2 PROPERTIES OF COMPOSITES specimen. An additional effect not significant Coupon tests in designing or testing ductile materials is the The determination of the tensile longitudinal tendency for composites to exhibit poor shear and transverse properties of unidirectionally properties. Indeed it is common for engineers reinforced composites can be obtained by test- designing metallic components to assume that ing relatively long coupon specimens. These failure in shear will not occur if failure in tenspecimens, as described in ASTM specifica- sion or compression is designed against. tions D3039-76, are 9 in long, 0.5 in wide and Unfortunately, shear resistance of composites from 0.02 to 0.10 in thick. The apparently is not directly related to tensile or compressive excessive length requirement is an attempt to properties; thus, a separate shear failure criteminimize the effect of specimen misalignment rion must be used when designing with with respect to the loading axis. Glass/epoxy composites. Shear failure can occur in-plane or tabs are bonded to the specimen ends to shear interlamina. the load into the specimen and to avoid damage and failure of the specimen within the gripped length. The longitudinal and trans-

,

196 Carbon fibers

In-plane shear testing A reasonably simple test method which utilizes a coupon with an eight layer symmetric +. 45 deg. layup is described in ASTM standard D35 18. If the specimen has strain gages oriented parallel to the x and y axes of the specimen (not the fibers), then the failure stress in shear is given by (0JmaX/2, the shear failure strain is ( E ~- &Jmax and the shear modulus is given by GI,= E J ~ ( E-~ E,) A rail shear test has been described in ASTM D4255-83. The shear stress and shear strain can be obtained at intermediate and maximum loads. Thus, GI, and the appropriate failure parameters can be obtained.

Interlaminar shear strength The interlaminar shear strength is a measure of the strength of the bond that exists between the various layers within the composite. It is important to know such a value since bending of beams can cause appreciable shear stresses, which while not large enough to cause failure of a traditional metallic structural material can fail a composite. This occurs either within the matrix or at the fiber-matrix bond line. The test specimen, as discussed in ASTM D2344, is a rather short, thick beam which is tested in three-point bending. Basically, the test specimen is sized such that the ratio of the shear stress generated at the midplane is maximized with respect to the tensile or compressive stresses generated in the outer fibers. If inspection of the failed specimen indicates that failure by shear has occurred, then beam theory indicates that a reasonably accurate estimate of the interlaminar shear strength F,, can be calculated from F3,=3P/4wh where P is the maximum load and w and h are the width and depth of the beam. Another way to measure the same maximum shear stress is by using the double-notched specimen as described ASTM

D3846. This specimen contains two square notches, each cut into an opposite face of a tensile or compressive specimen. If the slots are cut a reasonable distance, I, apart in order to avoid interactions between the stress fields of the cut slots, then the interlaminar shear strength is given by

F,, = P/wl

9.13 CONCLUSION AND PERSPECTIVES

Low performance isotropic non-continuous pitch-based fibers and anisotropic PAN-based large tow fibers have been available for a number of years in the USA priced at $20/kg. These fibers have been used to produce electrically conducting flexible heating materials to control the temperature of pipelines, to heat human dwellings, hot-houses, etc. (Karpinos and Izmalkov, 1982; Glushchenko and Griffen, 1982; Levit, 1986; Jakubowski and Subramanian, 1979).In addition, low modulus carbon fiber materials and their composites have been used as abrasive, anti-friction, sealing and heat-insulating materials. Other nonstructural uses include activated carbon fibers which can be used directly for the purlfication of nitrogen oxide-containing fumes, chimney smoke, automobile air conditioners, various respiratory, water purification (Chupolov et al. 1983; Richter, Knoblauch and Juntgen, 1984). In recent years carbon fiber adsorbents have been used in medicine to remove toxic substances from body fluids (Ternovoiet al., 1985). The high modulus/high strength continuous carbon fibers have been available at a cost of $60-80/kg. Thus, reinforced epoxies, polyesters and other polymers have demonstrated properties that have enabled them to find use in the aerospace and aeronautical fields. Smaller quantities of fibers have been used in high quality sports gear such as golf clubs, fishing rods, tennis rackets, marine sports, ski equipment, bicycle equipment, etc. Fibers have also been used in many civil engineering

Conclusion structures. These include short fiber reinforcement of cement mortar and continuous reinforcementof concrete. Some attempts have been used to provide earthquake resistant structures and fiber reinforced ropes and some successes have been reported in fabricating structures with carbon fiber reinforced aluminum and copper. Carbon fiber reinforced carbon has been considered for high temperature load bearing structures and heat shields in spacecraft and supersonic aircraft. And, finally, similar materials are considered useful in nuclear applications and disc brake materials for aircraft, high speed trains and racing cars. At the present time, there is a growing interest in the use of very high thermal conduction properties to manage local temperatures in sensitive electronic equipment. Nevertheless, despite all these apparent successes, it is vitally important to realize that the expanded use of high performance carbon fibers depends very sensitively on the lifetime costs involved in substituting carbon and its composites for competing metals, ceramics and polymers. It has been recognized for many years that market penetration of carbon fibers will always be limited to rather sophisticated structures if the cost remains at the present high level. However, the precursor PAN material presently costs about $5/kg; thus, taking into account the weight loss and processing costs involved in converting PAN to carbon, it is unlikely that large amounts of carbon fiber made from such material will ever be less than $10/kg. Conversely, the cost of the pitch precursor material is almost insignificant, since it is the byproduct of a commercial process established to produce other end products: gasoline, metallurgical coke, etc. In addition, since the fiber is fabricated using a melt spinning process, the production rate can be much faster than the wet or dry process used to produce PAN based fibers. The carbon yield from pitch precursors can average up to 85%, whereas the carbon yield produced from PAN averages about 65%. Pitch based carbon fibers

197

can be made with modulus values much larger than can be obtained from PAN precursors. Unfortunately,rather sophisticated and expensive pitch pretreaments must be applied to a petroleum or coal tar pitch in order to produce a high performance fiber. The pitch softening temperature is much higher than PAN; thus stabilization can potentially be carried out quickly at higher temperatures. Nevertheless, stabilization of both PAN and pitch materials is exothermic and, in order to avoid overheating, thermal runaway and decomposition of the precursor, a less than advantageous temperature of oxidation must be used. Future developments in these areas should therefore involve methods to increase the stabilization rate and the development of new precursor materials. Specific topics might include stabilization in thermally stable environments (fluidized beds, liquids, etc.) and the development of alternative synthetic precursor materials (polymers and/or pitches). Cost of the final component, while very sensitive to the cost of raw materials, also involves all of the design, fabrication and testing costs. All of these component costs must be tightly controlled if economically viable performance increases are to be realized. When using composites, it has been found to be vitally important to pay strict attention to detail design if a maximum expected weight saving is to be realized. Designers must optimize the total vehicle weight and not simply substitute a carbon composite for a metal one. It is rumored for instance that, due to overly conservative design (i.e., the use of metallic joining and fabrication techniques etc.), the resulting weight of some composite structures has, in the past, turned out to be as heavy as similar components built from aluminum. 9.14 CONCLUSION

Many people appear to believe that despite nearly thirty years of development, carbon fibers are still an evolving space age material. Until very recently, there was a production

198 Carbon fibers over-capacity in the carbon fiber industry. For example, in 1995 it was estimated that 10 000 000 kg of carbon fibers were sold from an estimated capacity of 16 000 000 kg. Exactly how accurate these estimates were is difficult to assess; however, it can be concluded that the market is small and cannot accommodate many producers. For this reason, many producers have seen fit to evaluate their position in the industry. For example, Table 9.3 is a listing of carbon fiber processors taken from Chapman and Hall's directory, 'Carbon and High Performance Fibers' which was published in 1991. In 1996, only the first eight of these were still producing significant commercial quantities of PAN-based carbon fibers. Table 9.4 is a similar listing for pitch-based fibers. In this case only the first five appear to be active.

At the present time, downsizing of the industry, increasing use of low cost fibers and the resurgence of orders for new commercial aircraft that now use increased quantities of carbon fiber has brought industrial capacity and market requirements closer together. Indeed, some fiber types are now difficult to obtain. Nevertheless, any major growth of the carbon fiber industry depends on the discovery of a method to produce fibers for one-half or one-third of the present projected large volume price and the development of new inexpensive fabrication methods for structures. These developments will initiate major new transportation based markets for the material. At the present price, however, the use of carbon fibers will always be limited to competitive performance driven applications.

Table 9.3 PAN-based tow manufacturers

Company

Country

Trade name

Akzo Carbon Fibers Inc. (Fortafil Fibers Inc.) Amoco Performance Products Inc. RK Carbon Fibres Limited Mitsubishi Rayon Co. Ltd. Soficar SA (Toray Industries Inc.) Toho Rayon Co. Ltd Toray Industries Inc. Zoltek Corporation Akzo NV (Fibres and Polymers Division)(Enka AG) Anglo-Soviet Materials Ltd Asani Kasel Carbon Fiber Co. Ltd BASF Structural Materials Inc.' Formosa Plastics* Hercules Advanced Materials and Systems Company2 Korea Steel Chemical Co. Ltd Nikkiso Co. Ltd* N.W. Chemical Power Co.* Sigri GmbH Textron Specialty Materials

USA USA UK Japan France Japan Japan USA Holland USSR Japan USA Taiwan USA Korea Japan China Germany USA

Fortafil Thomel RK Pyrofil Torayca F Besfight Torayca Panex Tenax Sapem Hi-Carbolon Celion

* Not available. Now Hexcel; * Now Toho.

?

Magnamite Kosca ?

?

Sigrafil Avcarb

References

199

Table 9.4 Pitch-based tow manufacturers

Company

Country

Trade name

Amoco Performance Products Inc. Mitsubishi Kasei Corporation Petoca Ltd (Kashima Oil Co. Ltd) Tonen Corporation Nippon Petrochemicals Co. Ltd Kawazaki Steel Co. Ltd Kobe Steel Ltd" Mitsubishi Oil Co. Ltd* Nippon Carbon Co. Ltd* Nippon Steel Co. Ltd" Osaka Gas Co. Ltd (Donac Ltd) Showa Shell Sekiyu

USA Japan Japan Japan Japan Japan Japan Japan Japan Japan Japan Japan

Thomel Dialead Carbonic Forca Granoc KMFC ? ? ? ?

Donacarbo-F Carbonexel

* Not available.

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Deurberque, A., Ph.D. 1990. Thesis, Universite de Pau et des pays de l'adour. Deurberque, A. and Oberlin, A. 1991. Carbon 29: 621. Donnet, J.B. and Bansal, R. 1984. Carbon Fibers, Vol. 3. New York: Marcel Dekker, Inc. Edie, D.D. 1990. Carbon Fibers, Filaments and Composites. (eds. Figneiredo, J. Bernard, C.A. Baker. R.T.K. and Hiittenger, K.J.) Kluwer Academic, pp. 647-655. Edie, D.D. and Diefendorf, R.J. 1993. Carbon Fiber Manufacturing. Park Ridge, NJ: Noyes Publications. Erlemenko, I.N., Plyublines I. and Gulko, N.V. 1990, Chemical Modified Carbon Fibers. New York: VCH Publishers, Inc. Eshback, O.W. and Souders, M., Handbook of Engineering Fundamentals, New York: John Wiley. Ezekiel, H.M. 1969. Appl. Polym. Symp. 9: 315. Fitzer, E., Muller, K. and Schaeffer, W. 1971. Chemistry and Physics of Carbon, pp. 237-383. New York: Marcel Dekker, Inc. Fitzer, E. and Heine, M. 1988.Fibre Reinforcementsfor Composite Materials Vol. 2; pp. 73-148. Amsterdam: Elsevier. (ed. A. R. Bunsell). French Patent 2 328 723. National Research Development Corp. French Patent 1 535 800, 1968. North American Aviation, Inc. Glushchenko, V.M. and Griffen, L.A. 1982. Gibkie Elektroprovodnye Materialy I Ustroistva na Osnove dlya Obogreva Lyudei I Tekhniki, pp. 23-29. Grassia, N. and McGuchan, R. 1971a. Eur. Polymer J. 71091-1104.

200 Carbon fibers Grassia, N. and McGuchan, R. 1971b. Eur. Polymer 1. 7: 1357-1371. Griffith, A.A. P. 1920. Trans. R. Soc. A221:163. Guigon, M. 1985. Relations entre la microtexture et les proprietes mecaniques des fibres decarbone ex-PAN. D.Sc. Thesis (These d’Etat), Universite de Technologie de Compiegne, France. Guigon, M. and Oberlin, A. 1986a. Composites Sci. Technol. 25: 231. Guigon, M. and Oberlin, A. 198613. Composites Sci. Technol. 27: 1. Guigon, M., Oberlin, A. and Desarmot, G. 1984a. Fibre Sci. Technol. 20: 177. Guigon, M., Oberlin,A. and Desarmot, G. 1984b. Fibre Sci. Technol. 20: 55. Hadcock, R.N. 1982. Design and Analysis of Composite Structures. In Handbook of Composites. New York: Van Nostrand Reinhold co. Hamada, T., Nishida, S., Matsumoto , Y. and Endo, M. 1987.1,Mater. Res. 2: 850. Herring, H. W. 1966. NASA Rep. No. IND-3202. High Performance Composites. 1994. July/August. Jain, M.K. andAlhiraman,A.S. 1987.1.Mater. Sci. 22: 278. Jakubowski, J.J. and Subramanian, R.V. 1979. Chem. Abstr. 1981, 95: 98874. Johnson D.J. 1987. Chemistry and Physics of Carbon 20: 1.New York: Marcel Dekker, Inc. Johnson, J.W. 1969. A p p . Polymer Symp. 9: 229. Johnson, J.W., Potter, W., Rose, P.G. and Scott, G. 1972. Brit. Polymer I., 4: 527-540. Johnson, J.W., Rose, P.G. and Scott, G. 1970. Proc. 3rd Conf. lndustrial Carbon and Graphite, London: Academic Press, p. 443. Jones, R.M. 1975. Mechanics of Composite Materials, Washington, D.C: Scripta Book Co. Joseph, D. and Oberlin, A. 1983a. Carbon 21: 559. Joseph, D. and Oberlin, A. 1983b. Carbon 21: 565. Karpinos, D.M. and Izmalkov, O.M. 1982. Gibkie Elektroprovodnye Materialy I Ustroistva na Osnove dlya Obogreva Lyudei I Tekhniki. Kawamura, K. and Jenkins, G.M. 1970. A New Glassy Carbon Fiber. 1.Mater. Sci. 5: 262. Kowbel, W., Wapner, P. G. and Wright, M. A. 1988. 1.Phys. Chem. Solid 49: 11. Lafdi, K., Bonnamy, S. and Oberlin, A. 1991a. Mechanism of anisotropy occurrence in a pitch precursor of carbon fibres, Part I - Pitch A and 0. Carbon 29: 831. Lafdi, K., Bonnamy, S. and Oberlin, A. 1991b. Mechanism of anisotropy occurrence in a pitch precursor of carbon fibres, Part I1 - Pitch C.

Carbon 29: 849. Lafdi, K., Bonnamy, S. and Oberlin, A. 1991~. Mechanism of anisotropy occurrence in a pitch precursor of carbon fibres, Part I11 - Hot stage microscopy of pitch B and C. Carbon 29: 857. Lafdi, K., Bonnamy, S . and Oberlin, A. 1992. Textures and structures in heterogeneous pitch-based carbon fibres (as-spun, oxidized, carbonized and graphitized); comparison with homogeneous fibres. Carbon 31: 29. Lafdi, K. and Oberlin, A. 1994a.A tentative to characterize and elaborate anisotropic pitches and derived carbon fibres. Part I: preparation by separation. Carbon 32: 11. Lafdi, K. and Oberlin, A. 1994b.A tentative to characterize and elaborate anisotropic pitches and derived carbon fibres. Part 11: preparation by bubbling. Carbon 32: 61. Lavin, J.G. 1992. Carbon 30: 351. Levit, R.M. 1986. Conducting Synthetic Fibers, Khimiya, Moscow. Lewis, I.C. 1982. Carbon 20: 519. Lewis, I.C. and Nazem, F.F. 1987a. 18th Conference Carbon, Extended Abstracts, American Carbon Society, p. 190. Lewis, I.C. and Nazem, F.F. 1987b. 18th Conference Carbon, Extended Abstracts, American Carbon Society, p. 290. Mair, W.N. and Mansfield, E.H. 1987. William Watt 1912-1985. Biographical Memoirs of Fellows of the Royal Society, 33: 643-667. Matsumoto, T. 1985. Pure Appl. Sci. 57: 1553. Mochida, I., Shimizu, K., Korai, Y., Ohtsuka, H. and Fujiyama, 5.1988. Carbon 26: 843. Nazem, F.F. 1982. Carbon 20: 345. Nazem, F.F. and Lewis, I.C. 1986. Mol. Cryst., Liq. Cryst. 139: 195. Oberlin, A. and Guigon, M. 1988. Fibre Reinforcement for Composite Materials, (ed. A.R. Bunsell). Amsterdam: Elsevier, p. 149. Oberlin, A. and Oberlin, M. 1981. Revue Ckim. Miner. 18: 442. Oberlin A. 1984. Carbon 22: 521. Oberlin, A. and Guigon, M. 1984. Science and New Applicafions of Carbon Fibers, Toyohashi University of Technology, Japan. Ohtsuka, H. 1988. Mitsubishi Gas-Chemical Co. Kurashiki, Okayama 712, Japan. Reynolds, W.N. and Moreton, R. 1980. Philos. Trans. Roy. SOC.A294: 451. Reynolds, W.N. and Sharpe, J.V. 1974. Carbon 12: 103. Richter, E., Knoblauch, K., Juntgen, H. Deutsche Offen. Patent 3 412 761, 1984.

References Riggs, D.M., Shuford, R. and Lewis, R. 1982. Handbook of Composites. New York: Van Nostrand Reinhold. Riggs, D.M., 1979. Doctoral Thesis. Rensselaer Polytechnic Inst., Troy, New York. Robson, D., Assabghy , F.Y.I. and Ingram, D.J.E. 1972. J. Phys. D, 5: 169. Rosen, B.W. 1964. AIAA 2: 1985. Schwartz, M.M. 1984. Composite Materials Handbook. New York: McGraw-Hill. Shindo, A., Nakanishi, Y. and Soma, I. 1969. Appl. Polym. Symp. 9: 305. Singer, L.S. 1978. Carbon 16:409. Tanabe, Y., Yasuda, E., Machino, H. and Kimura. S. 1987. Ann. Mtg Jpn Ceramic Society, Nagoya, 77. Temovoi, K. S., Zemskov, V. S., Kolesnikov, E. B. and Mashkov, 0. A. 1985. Sorbitsionnaya v Khirurgicheskoi Klinike Defoksikatsiya (Detoxification Sorption in Surgery) Kishinev (USSR):Shtiintsa. Tsai, S.W. and Hahn, H.T. 1980. Introduction to Composite Materials, Technomic Publishing Co., Inc., Westport, CT. US Patent 3 107 152,1963. Ford and Mitchell. US Patent 3 533 741,1970. Courtaulds Limited. US Patent 3 627 466,1971. Steingiser. US Patent (12) 3 629 379, 1971. Otani. US Patent 3 650 668,1972. Celanese. US Patent 3 656 882,1972. Celanese. US Patent 3 656 883,1972. Celanese. US Patent 3 656 903,1972. Celanese. US Patent 3 660 140, 1972. Scola. US Patent 3 671 411,1972. Ray. US Patent 3 677 705,1976. Celanese. US Patent 3 723 150,1973b. Druin. US Patent 3 746 450,1973. Goan. US Patent 3 746 506,1973. Aitken. Druin. US Patent 3 754 957,1973~~ US Patent 3 767 773,1973. Turner. US Patent 3 767 774, 1973. Hou. US Patent 3 780 255,1973. Boom. US Patent 3 814 377,1974. Monsanto. US Patent 3 832 297,1974. Paul, Jr. US Patent 3 841 079,1974. Celanese US Patent 3 846 833,1975. Celanese. US Patent 3 859 187,1975. Druin.

201

US Patent 3 894 884,1975. Druin. US Patent 3 917 776,1975. Mitsubishi Rayon. US Patent 3 919 383,1975. Singer. US Patent 3 972 984,1976. Nippon Carbon Co. US Patent (8) 3 974 264,1976. McHenry. US Patent (8) 3976729, 1976. Lewis, McHenry, Singer. US Patent 3 989 802,1976. Loo. US Patent 4 004 053,1976. US Patent 4 005 183,1977. Singer US Patent 4 009 248,1977. US Patent (4) 4 017 327, 1977. Lewis, McHenry, Singer. US Patent 4 024 227,1978. US Patent 4 031 288, 1978. Minnesota Mining and Manufacture Co. US Patent 4 039 341, 1976. National Research Development Corp. US Patent 4 069 297,1978. Toho Beslon Co. Ltd. US Patent (6) 4 208 267,1980. Diefendorf and Riggs. US Patent (5) 4 331 620,1982. Diefendorf and Eggs. US Patent 4 374 114,1983. Kim. US Patent (3) 4 376 747,1982. Nazem. US Patent 4 504 454,1985. Riggs. Watt, W. 1972. Carbon 10: 121. Watt, W. and Johnson, W. 1975. Mechanism of oxidation of polyacrylonitrile fibres. Nature 2 5 7 210-212. White, J.L. 1992. ONR Report for Contract No. 88-K-0424 and 89-J-3056. Wright, M.A., 1989. NASA Conference Publication 3054: 17. Wright, M.A. and Iannuzzi, EA. 1973. J. Comp. Mat., 7: 430. Wright, M.A. and Wills, J.L. 1974. J. Mech. Phys. Sol. 22: 161. Wright, M.A. and Palmer, K.R. 1994. Research into Structural Carbons, Materials Technology Center Publication, SIUC, Carbondale, Illinois, 62901. Yanagida, K., Noda, M., Sasaki, T. and Tate, K. 1991. 20th Conf Carbon, Extended Abstracts, American Carbon Society, p. 160. Yoneshoga, I. and Teranishi, H. 1970. Japanese Patent Specification 2774/70. Yooh, S.H., Korai, Y. and Mochida, I. 1994. Carbon 32: 281.

ORGANIC FIBERS

10

Linda L. Clements

A different type of h g h performance organic fiber, extended chain polyethylene fibers, was Before the first aramid fibers were introduced added in the 1970s. While inferior to inorganic in the 1960s and 1970s, organic fibers were relfibers in some properties, organic fibers provide atively low performance materials, primarily combinations of properties not available with used in textile applications. Now several different types of high performance organic inorganic fibers and so have made possible new fibers exist, all competitive with inorganic designs and applications. In this chapter, only high performance fibers in some or even most of their properties. organic fibers which are commercially availThe market demand for these fibers exceeds able will be discussed in detail, although fibers one billion dollars (Adams and Farrow, 1993a). which are nearing commercialization will be The main applications for high perfordiscussed briefly. For a more complete review mance organic fibers today are in asbestos of both commercially available and experireplacement, ballistics, rubber reinforcement, mental high performance organic fibers, see ropes and cables and composites. Most of the Yang (1989, 1992). usage is of aramid fibers, with over 18000 metric tons used each year. Both usage and existing capacity for other organic fibers are 10.2 ARAMID FIBERS only a fraction of this value (Adams and Farrow, 1993a). 10.2.1 OVERVIEW Tlus broad market for organic fibers is a direct outgrowth of applying the basic princi- Aramid fiber is the generic term for a specific ples of polymer science to produce a new and type of ’aromatic polyamide fiber.’ The US exceptional engineering material. In the 1950sit Federal Trade Commission defines an aramid was recognized that if a means could be found fiber as ‘a manufactured fiber in which the to form certain intractable polymers into fiber-forming substance is a long-chain synextended chain fibers, very high stiffnesses, thetic polyamide in which at least 85% of the strengths and use temperatures could be amide linkages are attached directly to two achieved. The difficulty of producing such aromatic rings.’ Thus, in an aramid, most of the amide fibers was solved in the 1960sby spinning from groups are directly connected to two aromatic liquid crystalline solutions. The first fibers prorings, with nothing else intervening. It should duced by t h s process were the aramids, which have since been followed by other such fibers. not be surprising that aramids have quite different properties from nylons and other conventional polyamides since the latter polymers contain few if any aromatic groups in the Handbook of Composites. Edited by S.T. Peters. Published main chain of the polymer. in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7 10.1 INTRODUCTION

Aramid fibers

203

Aramid fibers can be separated into two include DuPont’s KevlarO, Akzo’s TwarorP, types: the para- aramids and the meta-aramids. Teijin’s TechnoraO and Kaiser VIAM‘s Amosa In para-aramids, the chain-extending bonds are and S W @fibers, while meta-aramids include in the para-position on the aromatic ring, as in DuPont’s Nomexs and Teijin’s TeijinconexO poly-p-phenylene terephthalamide (PPTA) (Fig. fibers. Hoechst AG also markets a para-aramid 10.1(a)),co-poly-p-phenylene/3,4’-oxydipheny- fiber in Europe. The para-aramids are the fibers lene terephthalamide (Fig. lO.l(b)) and used in high performance applicationsand thus poly-p-phenylene-benzimidazole-terephthala- will be emphasized in this chapter. mide (Fig. lO.l(c)). In meta-aramids, on the other hand, the chain-extending bonds are in 10.2.2 MANUFACTURE the meta-position on the aromatic ring, as in poly-m-phenyleneisophthalamide (MPIA) (Fig. Historically, meta-aramid fibers were the first 10.1(d)). Commercially available para-aramids to be produced, with DuPont’s Nomex fiber 0

II

-0 .C

H

\

k-(=&IA

\

H

1 -0;-0 H

0

I

I1 C

H

N

0

H

I1

I

C

-0

0-

HI 0

\

1

H I

0

H 0

0

I1

C

\ H NI

d+0

Fig. 10.1 Structural formulae of (a) the para-aramid poly-p-phenylene terephthalamide (PPTA), (b) the para-aramid co-poly-p-phenylene/3,4’-oxydiphenyleneterephthalamide, (c) the para-aramid poly-pphenylene-benzimidazole-terephthalamide(PBIA), and (d) the meta-aramid poly-m-phenylene isophthalamide (MPIA).

204 Organicfibers being introduced in the 1960s. The first paraaramids were synthesized in 1965 by S.L. Kwolek of DuPont (Kwolek, 1971; Kwolek, 1972; Kwolek, 1974).Forming these into usable fibers is very difficult because para-aramids show no melting point and are soluble in a limited number of solvents. The problem of spinning the polymer into fibers was solved for PPTA following the discovery that the polymer would dissolve in strong acids to form a liquid crystalline solution. Undiluted sulfuric acid is the solvent usually used. Blades (1973, 1974) devised a special manufacturing process - known as continuous dry jet wet spinning - for forming the liquid crystalline solution into filaments. The polymer solution is extruded through spinnerets at elevated temperature through an air layer into a coagulating water bath. The cold water bath also contains a base to neutralize and remove the retained acid. Continuous dry jet wet spinning is the manufacturing technique used for most para-aramid fibers. Teijin’s Technora fiber, however, is produced by wet spinning followed by drawing (Hongu and Phillips, 1990).

10.2.3 STRUCTURE

The excellent properties of para-aramids result from both chemistry and physical microstructures. In both meta- and para-aramids, the aromatic rings in the backbone chain produce high thermal resistance. In addition, in paraaramids the orientation of the chain-extending bonds produces a polymer which is an extended-chain rigid rod. Spinning produces a fiber made up of extended-chain crystallites which are almost completely aligned parallel to the draw direction and to each other. The crystallites have a very high length-to-diameter ratio and extensive interconnection of molecules between crystallites. Thus, an unbroken ’infinite’ filament can be formed. Within the crystallite the chains are bonded to one another by hydrogen bonds, as shown in Fig. 10.2. Although these bonds are not nearly as strong as the covalent bonds which occur within the molecules, hundreds or even thousands of such bonds form between adjacent para-aramid molecules. Since the molecules are rigid, the only way to separate them in tension is to break all of the hydrogen bonds at once. This requires a large force and

0 I

H

H

H

A&=(-*

0

I

I

0

-N

II

\ llc - e

-C

\

H

H

\

0

0

-0 H

Fig. 10.2 Schematic showing hydrogen bonding between PPTA molecules in the crystallite.

Aramid fibers 205 is the reason para-aramid fibers are exceptionally strong in axial tension. However, since the bonds can be broken easily one at a time, the fibers are quite susceptible to damage by bending, buckling or transverse loading. In meta-aramids, on the other hand, a crooked chain results. Since even in pure tension the chain-extending bonds can flex and rotate, meta-aramids are much less rigid than para-aramids and not as strong. However, because the chains are more flexible, metaaramids are easier to manufacture than para-aramids and are less expensive.

mechanical, thermal, physical and other properties. This anisotropy may produce design limitations, but can also be used to advantage.

Physical and thermal properties

Table 10.1 compares the physical and thermal properties of some representative aramid fibers. Due to their highly aromatic and ordered structure, aramids have very high thermal resistance for organic materials. They do not melt prior to decomposition,in spite of the fact that they are technically classified as thermoplastics. This is because melting of the 10.2.4 PROPERTIES crystalline phase, like rupturing the fiber in Aramid fibers offer some significant advan- tension, would require that all of the hydrogen tages over other fibers, but also have their bonds between two molecules be severed at drawbacks and limitations. Both advantages once. Nonetheless, because of decomposition, and limitations will be described more fully in their temperature resistance is not equal to the sections on properties and in the sections that of inorganic fibers. Thermogravimetric on design considerations and applications. analysis of Kevlar fibers shows that weight Both DuPont’s Kevlar family of fibers and loss begins at above 350°C (660°F)in air (Penn Akzo‘s Twaron fibers are based upon PPTA and Larsen, 1979; Yang, 1992), with complete (Fig. lO.l(a)). Teijin’s Technora fiber and the decomposition occurring at between 427 and para-aramid marketed by Hoechst AG in 482°C (800 and 900°F) (DuPont, 1992a). Europe, on the other hand, are a para-aramid Exposure to elevated temperature will copolymer, co-poly-p-phenylene/3,4’-oxy- degrade the properties of aramid fibers. Figure diphenylene terephthalamide (Fig. lO.l(b)).It is 10.3 shows the strength retention of Kevlar 29 likely that Kaiser VIAM’s SVM fibers are poly- and Technora fibers as a function of time and p-phenylene-benzimidazole-terephthalamide temperature. This change in properties occurs (PBIA), (Fig. lO.l(c)) rather than PPTA as a result of slow oxidation. For this reason, (Gerzeski, 1989). Kaiser VIAM’s Armos fiber the long-term use temperature of para-aramid may be PBIA or PPTA. Both DuPont’s Nomex fibers is typically limited to about 150-175°C and Teijin’s Teijinconex fibers are based upon (300-350°F). MPIA (Fig. lO.l(d)).These chemical and strucIn the transverse direction para-aramids are tural differences produce different properties like most other materials in that they expand for the fibers. In addition, differences in spin- with increasing temperature. However, in the ning conditions and, most importantly, longitudinal direction the fibers actually conpost-spinning heat treatments are used to alter tract somewhat as temperature increases. The properties further. For example, by changing negative thermal expansion coefficient of processing conditions, Kevlar fibers can be para-aramids can be used to advantage to produced with elastic moduli ranging from 63 design composites with tailored or zero therto 143 GPa (9 to 21 Msi) and elongations at mal expansion coefficient. Aramids are flame resistant but can be break from 1.5 to 4.4%. Because of the anisotropy of their microstruc- ignited. While pulp or dust of Kevlar may conture, para-aramid fibers have very anisotropic tinue to smolder once ignited, fabrics do not

206 Organicfibers Table 10.1 Physical and thermal properties of representative aramid fibers

Fiber

Kevlar 49

Twaron HM

Technora

Nomex

Teijinconex

Type

para-aramid

para-aramid

para-aramid copolymer

meta-aramid

meta-aramid

DuPont 1992a 1.44 (0.0520)

Akzo 1990,1991 1.45 (0.0524)

Teijin 1989,1993 1.39 (0.0502)

DuPont 1981,1993g 1.38 (0.0499)

Teijin 1991 1.38 (0.0499)

-538°C" (1000°F)

>5OO0C (>932"F)

-

>371"C (>700"F)

Decomposition temperature in air

427482°C (800-900°F)

500°C (930°F)

500°C (930°F)

371°C (700°F)

400430°C (750-805°F)

Long-term use temperature in air

149-177°C (300-350°F)

Longitudinal linear thermal expansion coefficientb 10-6/ "C

-4.9 (-2.7)

Reference for data Density g cm-3 (lb in-?) Melting temperature

-3.5 (-1.95)

-6.2 (-3.4)

+15 (+8.3)

+20 (+11)

1.09 (0.26)

1.21 (0.29)

1.05 (0.25)

0.13 (22)

0.13 (22)

OF)

Transverse linear thermal expansion coefficientb: /"C ( / O F )

+66" (+37)

Specific heatb kJ/kg K (BTU/lb OF)

1.42 (0.34)

1.42 (0.339)

Longitudinal thermal conductivityb W/m K BTU/h ft OF)

4.11' (2.38)

4.0 (2.3)

Transverse thermal conductivityb W/m K BTU/h ft OF)

4.82' (2.79)

5.0 (2.9)

3.5%

3.5%

2.0%

4.5%

5.0-5.5%

Typical filament diameter pm in)

12 or 15 (0.48 or 0.59)

12 (0.48)

12 (0.48)

max: 15-17 (0.6-0.7)

-10 to 15 X 45 (-0.4 to 0.6 X 1.1)

Typical filament shape

round

round

round

oval to dogbone

oval to dogbone

Equilibrium moisture contentb

* Data from Yang, 1992.

Varies with temperature; room temperature values are given. Data from Chiao and Chiao, 1982.

Arumidfibers

rL

I

I

I

I

207

100

.

be

80

C

0

+ C

2w

60

L11

fm

e

I=

40

-

iz 20 -

0' 0.1

- Technorag I

I

I

I

1

10

100

1000

L

Time, h

Fig. 10.3 Strength retention of Kevlar 29 and Technora fibers following elevated temperature exposure (DuPont, 1992a; Teijin, 1989).

continue to burn when the flame source is removed (DuPont, 1992a). The lower thermal conductivity of aramids compared to inorganic fibers can improve the fire resistance of their composites, since aramids do not readily conduct heat into the more volatile matrix.

Mechanical properties Composite materials are most commonly used because of their superior strength and/or stiffness at a given weight as compared to conventional structural materials. Figure 10.4 compares the specific strengths and specific stiffnesses of various reinforcing fibers. (The strengths and stiffness in Fig. 10.4 are expressed in units of grams per denier (gpd). This is a textile term often used for organic fibers which measures specific strength and/or stiffness. This term is further explained in the appendix to this chapter.) As can be seen, aramid fibers perfonn very well. In fact, until the emergence of high strength intermediate modulus carbon fibers and the commercialization of polyethylene

fibers in the mid-l980s, aramid fiber composites had the highest specific strengths of all composite materials. Although composites from newer fibers have taken over that position, aramids still offer outstanding combinations of properties, such as high specific strength, toughness, creep resistance and moderate cost, for specific applications. Table 10.2 compares the mechanical properties in axial tension of several commercially available aramid fibers. Aramid fibers have some definite limitations. They are weak in bending and show obvious damage if subjected to kinking or buckling. As a result, they are also weak in compression (where microbuckling is inevitable) and in transverse tension (wherebond-by-bond breakage of hydrogen bonds is likely). In addition, even though the para-aramid chain is quite polar in nature, almost all of the polar groups are fully involved in hydrogen bonding to other aramid molecules. As a result, paraaramid fibers do not form strong bonds with other materials such as composite matrices,

208 Organicfibers Table 10.2 Axial tensile mechanical properties of representative aramid fibers Fiber

Reference

Spec$c gravity

Initial tensile modulus, GPa (Msi) Bare"

Epoxyimpregnatedb

Tensile strength, MPa ( h i ) Bare"

Bare" 3.6

Kevlar Type 956, 1500 denier

DuPont, 1993h

1.44

71.8 (10.4)

Kevlar 29 Type 964, 1500 denier

DuPont, 1992a, DuPont, 19938

1.44

70.5 (10.2)

83.0 (12.0)

2920 (424)


DuPont, 1992a, DuPont, 1993g

1.44

112.4 (16.3)

124.0 (18.0)

3000 (435)


DuPont, 1993g

1.44

99.13 (14.4)

3050 (442)

-

2.9

Kevlar 119 1500 denier

DuPont, 1990

-

54.6 (7.9)

3050 (442)

-

4.4

Kevlar 129 denier unspecified

DuPont, 1993i

1.44

96.0 (13.9)

3380 (490)

-

3.3

Kevlar 149 Type 965A, 1140 denier

DuPont, 19938

1.47

142.7 (20.7)

2340 (339)

-

1.5

Kevlar HT Type 964C, 1000 denier

DuPont, 19938

1.44

99.1 (14.4)

3370 (489)

-

3.3

Kevlar KM2 850 denier

DuPont, 1992d

-

63.4 (9.2)

3280 (476)

-

4.0

Twaron

Akzo, 1991

1.44

70 (10.2)

2800 (406)

Twaron Perkins, 1993 Type 2000, 930 denier 'microfilament'

1.44

88 (12.8)

3230 (468)

Twaron HM

Akzo, 1991

1.45

Armos 58.8 tex

Kaiser VIAM,

103 (14.9) 147 (21.4)

SVM 58.8(300) X17-1000

Kaiser VIAM, 1993g' Gerzeski, 1989

-

1993a'

1.43

2920 (424)

Epoxyimpregnatedb

Elongation at break, %

-

3600 (525)

3500 (508) -

3500 (508) -

3.6

3.6 3.3

2.5 3.2

123 (17.8) Continued on next page

Aramid fibers 209 Table 10.2 Continued

Fiber

Reference

Specific gravity

Initial tensile modulus, GPa (Msij Bare"

Epoxyimpregnatedb ~~

-

Teijin, 1989

1.39

Nomex Type 430, 1200 denier

DuPont, 1993g

1.38

11.6 (1.68)

-

Teijinconex

Teijin, 1991

1.38

7.9-9.7 (1.1-1.4)

-

11.6-12.2 (1.7-1.8)

-

a

Teijin, 1991

1.38

Bare"


Epoxyimpregnatedb

-

3440 (498)

-

4.6

596 (86.6)

-

28.0

610-670 (88-97)

-

3545

730-850 (110-120)

-

20-30

Data for DuPont fibers taken from conditioned yarns tested according to ASTM Standard D885. Modulus data for Akzo fibers from testing according to ASTM Standard D885M. Test technique unspecified for Akzo fiber strengths and elongations and for all data from Kaiser VIAM and Teijin fibers. Data for DuPont fibers taken from epoxy-impregnated strands tested according to ASTM Standard D2343. Data for Akzo fibers from testing according to impregnated strand test method DIN 65356, part 2. Test technique unspecified for Kaiser VIAM fibers. Preliminary data. 50

40

. . . . I

..

I . I . . . I l l l l l l . . . .

-

l l . . I I I I . . I . I . .

-

PED HM Armos 0

Carbon TlOOOG 30 - Technora 0

'pectra 'Oo0

0

Dyneema SKBO

Carbon

Tekmilon 'spectra 900

. Vectran HS 0 Kevlar 49 . Twaron O O S ~0 Carbon T-300 . .waron HM 2o - S-Glass 0 Kevlar 149 0 Carbon T-50

.

PE (H.C.)

Boron

E-Glass Carbon P-100 .

Steel

Bare"

~

73 (10.6)

Technora

Teijinconex HT

Tensile strength, MPa (ksij

1

2 10 Organicfibers and/or prepared using other fabrication processes, the general trend is valid: aramid fiber composites have poor off-axis properties. In axial tension, both aramid fibers and their composites are linear to failure. In spite of this fact, the same microstructural characteristics which lead to the weakness of aramid fibers in buckling also make them very tough. During failure, the widespread bending, buckling and other internal damage to the fibers absorbs a great deal of energy. Similarly, the strength of aramid fibers is not very strain rate sensitive: an increase in strain rate of more than four orders of magnitude only decreases the tensile strength by about 15%. (Abbott et al., 1975) This property alone provides design advantages over all inorganic and many other organic fibers. The mechanical properties of aramid fibers decrease with increasing temperature. Figure 10.5 shows the fiber elastic modulus as a function of temperature for several organic fibers. At 177°C (350°F) the modulus of para-aramid fibers is about 80% of that at room temperature. Figure 10.6 compares the fiber tensile

further aggravating the poor transverse, bending and compressive properties of the fiber itself. The basic chemical structure differences between the aramid fibers produce many of the mechanical property differences seen in Table 10.2. The ether (-0-) linkages in the backbone of the Technora copolymer fiber produce a lower modulus than that of Kevlar and Twaron PPTA-based fibers. On the other hand, the additional cyclic ring in the SVM PBIAbased fibers produces a higher basic modulus. However, heat treatment and other fabrication steps can also alter mechanical properties significantly, as is seen in the property differences between the various Kevlar fibers. The mechanical properties of aramid composites are illustrated in the data of Table 10.3. For this filament-wound composite the longitudinal compressive strength was about one-eighth that in longitudinal tension, the inplane shear strength was one-seventy-fifth and the transverse tensile strength over two hundred times smaller. While the relative values of properties may change for composites made from other aramid fibers and/or other matrices

Table 10.3 Mechanical properties of a filament-wound composite of 60 vol YO aramid fiber in a room-temperature curable epoxy matrix (Clements and Moore, 1977)

Fiber: DuPonf’s Kevlar 49, Type 968, 1420 denier Matrix: 100 parts Dow Chemical DER 332 (diglycidyl ether of bisphenol-A epoxy) and 45 parts Jefferson Chemical reffamine T-403 polyether triamine 1 day at room tnnperuture, postcure 16 h ut 85°C (185°F) Cure: Elastic constants: Longitudinal Young’s modulus E,,, GPa (Msi) Transverse Young’s modulus E,,, GPa (Msi) Shear modulus G,,, GPa (Msi) Major Poisson’s ratio vl, Minor Poisson‘s ration u,,

Ultimates:

81.8 f 1.5” (11.9 k 0.22) 5.10 k 0.10 (0.74 & 0.014) 1.82 k 0.09 (0.26 f 0.013) 0.310 k 0.035 0.0193 f 0.0014

Tension

Longitudinal strength, MPa (ksi) 1850 f 50 (268 f 7.3) 2.23 f 0.06 Longitudinal ultimate strain, Yo Transverse strength, MPa (ksi) 7.9 k 1.1(1.15 f 0.15) 0.161 f 0.023 Transverse ultimate strain, Yo Shear stress at 0.2%, offset, MPa (ksi) Shear strain at 0.2% offset, Yo

Compression

In-plane shear

235 f 3 (34.1 k 0.4) 0.48 5 0.3 53 f 3 (7.7 f 0.4) 1.41 f 0.12

-

-

24.2 f 2.4 (3.51 k 0.35) 1.55 k 0.16

-

a

Limits are 95% confidence limits. Each value is the result of five or more tests.

Aramidfibers

211

s

si - 5 25

-

1992a; Teijin, 1989).

3500

-1

a

-

1

8

-

Techno ra

m

q

: -

2500 1

Kevlar

8 =

;:0003 n g 2000 0,

t?

1500

3i

500

400

300

-

-

.-

Y"

f m c

200 ?!

Polyester

looo;, Nylon

- 100

500 0

, ,: 0

Fig. 10.6 Tensile strength as a function of temperature for two para-aramid fibers and for two polymer fibers and steel (DuPont, 1993h; Teijin, 1989).

212

Organicfibers

strength as a function of temperature for sev- wet transverse tensile and in-plane shear eral organic fibers. For Kevlar fiber the strengths were only about half of the 52% r.h. strength at 177°C (350°F) is about 80% of that values. The data in boiling water illustrate that at room temperature, while for Technora the the drops in strength due to the presence of strength is about 70% of the room temperature moisture alone were almost as severe as those value. On the other hand, at cryogenic tem- due to the combined presence of moisture and peratures modulus increases slightly and elevated temperature. This relative loss in properties is less for the Technora para-aramid strength is not degraded. The presence of moisture also reduces the co-polymer fiber. Care must be exercised when mechanical properties of aramid fibers and using aramid composites in high moisture their composites. The effect upon longitudinal applications. Both para-aramid co-polymers and homotensile properties is relatively small, but the loss is pronounced for off-axis properties. polymers exhibit very little creep. In general, Table 10.4 illustrates this loss for Kevlar 49 creep strain increases with increasing temperafiber in a room-temperature curable epoxy. ture, increasing stress and decreasing fiber The longitudinal tensile strength in water at modulus. Like all high performance fibers, room temperature was 88% of that for com- under long term loading, para- aramids are posites equilibrated at room temperature and subject to stress rupture, i.e. failure of the fiber 52% relative humidity (r.h.). The wet longitu- under sustained loading with little or no dinal compressive strength, on the other hand, accompanying creep. Figure 10.7 compares the was only 75% of the 52% r.h. value, while the stress rupture performance of Kevlar 49 to that Table 10.4 The effect of environments on the mechanical properties of a filament-wound composite of 50 vol Yo of an aramid fiber in a room-temperature curable epoxy matrix (Wu, 1980) Fiber:

DuPont's Kevlar 49, 4560 denier

Matrix: 100 parts Dow Chemical DER 332 (diglycidyl ether of bisphenol-A epoxy) and 45 parts Jefierson Chemical Jeffamine T-403 polyether triamine Cure:

Infrared heating, postcure 2 h at 100°C (212°F) Strength, MPa (ksi)

________

23"C, dry

23°C' 52% r.k.

23"C, water

1OO"C, water

Longitudinal tension

1370 f 6 2 " (199 f 9)

1340 f 112 (194 f 16)

1190 f 62 (173 f 9)

1150 f 1 2 4 (167 f 18)

Longitudinal compression

188 f 1 2 (27.3 f 1.7)

169 f 20 (24.5 f 2.9)

126 f 22 (18.3 f 3.2)

107 f 2 1 (15.5 f 3.0)

Transverse tension

7.6 f 1.6 (1.10 f 0.23)

74 f 1.2 (1.07 k 0.17)

3.9 f 0.7 (0.57 f 0.10)

3.6 f 0.2 (0.52 f 0.03)

Transverse compression

31.3 f 3 . 2 (4.54 kO.46)

29 f 4.0 (4.21 f 0.58)

22.5 f 3 . 2 (3.26 f 0.46)

22.1 f 23.6 (3.20 f 3.42)

In-plane shear

27 f 3 . 0 (3.92 f 0.44)

26.5 f 1.6 (3.84 f 0.23)

13.8 f 2.2 (2.00 f 0.32)

13.6 f 2.5 (1.97 f 0.36)

-

4.1

7.8

8.9

Hygrothermal Properties Equilibrium moisture concentration, Yo ~

Limits are 95% confidence limits. Each strength is the average of five tests.

Aramid fibers 213

0

+, 100

! 2 .I4

Kevlafl

90

49

1 W

o

80

d

.

a a

70

u

,"

60

a

a 0

50

.rl

rl

2a

40

10-2

10-1

1

io

io2

103

104

105

Lifetime, h

Fig. 10.7 Stress-rupture behavior of epoxy-impregnated Kevlar 49 fibers compared to that of epoxyimpregnated S-glass fibers (Chiao,Chiao and Sherry, 1976).

of Sglass. Para-aramids perform well under these conditions, but the phenomenon of stress rupture must be considered in any design where long term loading is anticipated. Strength retention cannot be used to estimate the remaining life of aramid fibers or composites under long term load (Chiao, Sherry and Chiao, 1976),so estimates of long term behavior must be derived from actual data, or accelerated testing methods (Chiao and Chiao, 1982). Para-aramid fibers and their composites perform very well in fatigue. For aramids, tension-tension fatigue generally is not of significant concern in applications where an adequate static safety factor has been used (Yang, 1992). Aramid composites have been found to be superior to glass fiber composites in both tensile-tensile and flexural fatigue loading. For the same lifetime (cycles to failure), Kevlar 49/epoxy composites can operate at a significantly larger percentage of their static strength than can glass-reinforced composites

(DuPont, 1986). Para-aramids also can be expected to perform better than carbon fibers in fatigue (Teijin, 1989; Yang, 1992). Technora para-aramid co-polymer is found to have even better fatigue resistance than the para-aramid homopolymer fibers (Teijin, 1989).

Chemical and environmental properties PPTA fibers are quite stable chemically; their resistance to neutral chemicals is usually very high. They are, however, subject to attack by acids and bases, especially by strong acids. Because the spin process used for Teijin's Technora para-aramid co-polymer produces a very pure polymer, the chemical and environmental resistance of Technora is superior to that of the PPTA fibers. Table 10.5 reports the resistance of Kevlar and Technora fibers to various chemicals. Technora has better acid and alkali resistance than PPTA and its steam resistance is also superior.

214 Organicfibers Table 10.5 Stability of para-aramid fibers in various chemicals

Concentration, Temperature,

None Slight Moderate Appreciable Degraded

40 40 90 90 20 10 10 10 10 10 10 20 40

21 (70) 95-99 (203-210) 21 (70) 95-99 (203-210) 20 (68) 71 (160) 20-21 (68-70) 20-21 (68-70) 21 (70) 99 (210) 99 (210) 95 (203) 95 (203)

1000 100 Tb 100 K 100 100 T 10 100 T 100 K,T 1000 100 10 100 T 100 T

28 10 10 saturated saturated

21 (70) 21 (70) 95-99 (203-210) 95 (203) 180 (356)

1000 1000 100 100 15

Formic Hydrochloric Nitric Phosphoric Sulfuric

Alkalis Ammonium hydroxide Sodium hydroxide

Organic solvents Acetone Benzene

100 100 100 Carbon tetrachloride 100 Ethylene chloride 100 Ethylene glycol/water 50/50 Ethylene glycol 100 Gasoline 100 Gasoline-leaded 100 Methyl alcohol 100 N-Methyl pyrrolidone 100

3 10 10 Sea water 100 Sea water (New Jersey) 100 Steam 100 100 100 100 Water, tap 100

a

hr

"C ( O F )

Acids Acetic

Other Sodium chloride

Effect on breaking strength"

Yo

Chemical

Portland cement

Time,

K

T

K K K K K

K T

K

T T

100 K 784 T 1000 K 100 1000 T 1000 T 300 784 T 1000 K 1000 K,T 100

21 (70) 99 (210) 121 (250) 95 (203)

1000 100 100 1000 1 yr 400 48

K K

100

T

100 100

K

-

K

K

boil 20 (68) 21 (70) boil 20 (68) 99 (210) 95 (203) 20 (68) 21 (70) 21 (70) 95 (203)

120 (248) 150 (302) 150 (302) 200 (392) 99 (210)

Kb

K K

T

K T K T K T

None, 0-10% strength loss; slight, 11-20% strength loss; moderate, 2140% strength loss; appreciable, 41430% strength loss; degraded, 81-100% strength loss. K is for Kevlar aramid fiber (DuPont 1989,1993h);T is for Technora aramid fiber (Teijin, 1989).

Aramid fibers Para-aramids are strong ultraviolet (W) absorbers. Upon exposure, the yellow or gold fibers turn first orange and then brown, due to degradation. The degradation occurs only in the presence of oxygen and is not enhanced by either moisture or atmospheric contaminants (DuPont, 1992a). Extended exposure may cause a loss of mechanical properties. Bare 1667 dtex (1500 denier) Kevlar 29 was found to have 71% strength retention after 1 month of outdoor exposure in Wilmington, DE and 43% after 4 months (Yang, 1992).In both processing and applications, para-aramids must be protected from W exposure, such as by painting or coating. However, since para-aramids are self-screening, UV protection may also be effected simply by dense packing of the fiber itself, with or without a matrix. Thus, bare 12.7mm (0.5 in) 3-strand Kevlar 49 rope was found to have 90% strength retention after 6 months outdoors in Florida and 69% strength retention after 24 months (DuPont, 1986). Unlike inorganic fibers, aramid fibers absorb water. For some aramid fibers the equilibrium moisture content (see appendix on page 241 for defirution) is quite high (5% for SVM, 7% for Kevlar, Kevlar 29 and Twaron), moderate for others (3.5% for Kevlar 49 and Twaron HM) and reasonably low for some (2% for Armos and Technora and about 1% for Kevlar 149) (Akzo, 1991; Kaiser VIAM, 1993a; Teijin, 1989; Yang, 1992). The equilibrium moisture content is directly proportional to the relative humidity, rising for Kevlar 49 to 6.2% at 96% r.h. (DuPont, 1992a). Absorbed moisture has only a small effect upon the tensile properties of the fibers, but a significant effect upon the transverse tensile, compressive, shear and flexural properties of the composite. The gain of moisture is completely reversible and once removed produces no permanent property changes. Electrical and optical properties Aramid fibers are electrical insulators. The process used to make the Technora fiber, however, leaves it with fewer ionic impurities than

215

in the para-aramid homopolymers and thus improved electrical properties. Technora fiber has a resistivity of 5 x lo’* Q/cm (Teijin, 1989). The dielectric constant of PPTA is 3.85 (Allied, 1989). The refractive index of Kevlar 49 fiber is 2.0 parallel to the fiber axis and 1.6 perpendicular (DuPont, 1986).Aramid fibers are opaque and are yellow to gold in color. 10.2.5 TREATMENTS

Unlike inorganic fibers, few surface treatments are used on aramid fibers to promote matrix adhesion. One reason is the futility of increasing the matrix bonding to the surface of a fiber which readily fails by defibrillation. Most dramatic improvements in fiber/matrix bonding give only modest improvements in off-axis strengths since they simply move the locus of failure from the surface to the interior of the filament. In other cases, longitudinal tensile strengths are adversely affected by otherwise successful surface treatments. Not all attempts at designing surface treatments have been unsuccessful, but for the most part the surface treatment used on commercial fibers is minimal compared to that used for inorganic fibers. Finishes - lubricants which aid in subsequent processing steps - are applied to aramid fibers for some applications. Available finishes are designed for such purposes as lubrication during weaving operations, improving abrasion resistance for cable applications or better performance in rubber goods. If the fiber is to be used in a high performance composite, however, the user will usually wish to avoid or remove any finish before impregnating the fiber with a matrix. Commercial aramid fibers may also be twisted. Twist may be quite useful in some applications and a small amount of twist will increase the strength of bare yarn or cord. [This optimum twist for Kevlar fibers occurs at a twist multiplier of 1.1.At about this value, the strength of bare yarn is the highest and the modulus is only slightly decreased from the

2 16 Organic fibers

untwisted level (DuPont, 1992b).] Twist will make the fiber easier to handle, make subsequent weaving or braiding operations easier and will improve the abrasion resistance of the fiber. It is also required for rope and cable applications. However, once the fiber is used in composite matrix fiber, twist is not desirable. This is because twist interferes with full impregnation of the fiber with resin and with stress transfer between adjacent fiber bundles. It also increases stress concentrations, particularly at higher twist levels. For this reason, most of the aramid fiber manufacturers supply most or all of their fibers untwisted or with minimal twist. 10.2.6 FORMS AND AVAILABILITY

Table 10.6 lists most of the commercial types of para-aramid fibers. Some of these fibers are readily available in a variety of fiber deniers, package sizes, finishes and so forth, while others are available only in limited quantities for specific applications. Due to constant changes in market conditions and other factors, the user is advised to check with the fiber manufacturer concerning current availabilities. The mechanical properties of fibers with different deniers and/or finishes and other treatments will vary somewhat from each other and from the nominal values given in Table 10.2. In addition to the yarns, tows and rovings listed in Table 10.6, DuPont’s Kevlar fibers are also available as staple (short fibers), floc (precision cut fibers of very short lengths) pulp (very short and high fibrillated fibers) and in specialty compounded forms (DuPont, 1992a, 1993h).A variety of fabrics are also produced. In addition, DuPont produces a colored fiber, Kevlar 100, in sage green, yellow, black and royal blue (Yang, 1992).Both Teijin’s Technora and Akzo’s Twaron fibers are available as staple and chopped fiber (very short lengths) and in a variety of fabrics (Teijin, 1989;Akzo, 1991). Technora is also marketed in black as well as natural color. Kaiser VIAM‘s Armos and SVM fibers are also expected to be offered as tape,

staple, pulp and in various fabrics. While the meta-aramid fibers are not usually used as fiber reinforcements in composites, they are used extensively as reinforcements for honeycomb sandwich core materials. The use of such materials along with composite face sheet panels has greatly extended the overall usage of composite materials, particularly in the aerospace industry. Information about the availability and package sizes of the fibers shown in Table 10.6, about other products and about special formulations can be obtained from Table 10.7. At the time of this publication, Kaiser VIAM’s Armos and SVM fibers are just being imported from Russia. For this reason, information on the fibers and their availability is limited in this chapter, but should be readily available later from the contact given in Table 10.7. Pricing Para-aramid fibers are currently priced from about $20 per pound for the larger denier fibers to about $60 per pound for most of the small denier, higher modulus fibers. (However, some of the very fine denier specialty fibers from some manufacturers cost hundreds of dollars per pound.) Prices can vary significantly for similar fibers of different deniers or from different manufacturers and thus price quotes should always be obtained before any decision is made upon use of a specific fiber. 10.2.7 DESIGN CONSIDERATIONS

In the 1970s and early 1980s aramids began to replace carbon and glass fibers in many applications. However, the development of high strength intermediate modulus carbon fibers in the mid-1980s and the commercialization of tough, high strength polyethylene fibers reversed this trend. Today aramid fibers are used mainly in applications where they offer a unique combination of properties, such as high specific strength combined with toughness and creep resistance.

Aramidfibers 217 Table 10.6 Availability of commercial para-aramid fibers"

Product (reference)

Count dtex (den)

Filament

~ _ _ _ _ _ _ _ _

number/ diameter yarn pm ( 1 C 3 i n )

Kevlarb (DuPont, 1993g; Yang, 1992) 666 Type 950 1110 (1000) 1670 (1500) 1000 1000 2500 (2250) 3330 (3000) 1333 Type956

800 (720) 1110 (1000) 1670 (1500) 2500 (2250) 3330 (3000)

Comments/typical applications

490 666 1000 1000 1333

12 12 15 15

(0.48) (0.48) (0.59) (0.59)

Finish: tire reinforcement

12 12 12 15 15

(0.48) (0.48) (0.48) (0.59) (0.59)

Mechanical rubber goods: hoses, belts, etc.

Kevlar 29 (DuPont, 1992b,1993a, 1993b, 1993f, 19938) Type 960 1670 (1500) 1000 12 (0.48) Cordage finish: high lubricity for improved 3330 (3000) 1333 15 (0.59) abrasion resistance; ropes and cables 100OOR' 12 (0.48) 17 OOO(15 000) 1110 (1000) 1670 (1500) 3330 (3000) 5000 (4500) 17 OOO(15 000)

666 1000 1333 3000 lOOOOR

Type 962

1670 (1500) 3330 (3000)

1000 1333

12 (0.48) 15 (0.59)

No finish; ropes and cables

Type 963

3330 (3000)

1333

15 (0.59)

Textile finish; non-apparel ballistic armor

Type964

215 (200) 430 (400) 1110 (1000) 1670 (1500)

134 267 666 1000

12 12 12 12

(0.48) (0.48) (0.48) (0.48)

Textile finish; ballistics and apparel, ignition cables

Kevlar 49 (DuPont, 1992b, 1993c-g) Type 965 61 (55) 25 215 (195) 134 420 (380) 267 1270 (1140) 768 1580 (1420) 1000 2400 (2160) 1000

15 12 12 12 12 15

(0.59) (0.48) (0.48) (0.48) (0.48) (0.59)

Textile finish; woven reinforcement in aerospace composites, ballistic armor, and printed circuit boards

12 12 12 12 15 15 15 12 12

(0.48) (0.48) (0.48) (0.48) (0.59) (0.59) (0.59) (0.48) (0.48)

No finish; marine composites, fiber optic cable reinforcement, ropes, filament-wound composites

Type968

215 (195) 420 (380) 1270 (1140) 1580 (1420) 2400 (2160) 3160 (2840) 4800 (4320) 5070 (4560) 7900 (7100)

134 267 768 1000 1000 1333 2000R 3200R 5000R

12 12 15 12 12

(0.48) (0.48) (0.59) (0.48) (0.48)

Textile finish; ropes and cables

Type 961


218 Organicfibers Table 10.6 Continued

Product (reference)

Count dtex (den)

Filament

Commenfs/fypicalapplications

number/ diameter yarn pm ( 1 P i n ) ~

_

_

Type 978

1580 (1420) 2400 (2160) 5070 (4560)

1000 1000 3200R

12 (0.48) 15 (0.59) 12 (0.48)

Cordage finish: high lubricity for improved abrasion resistance; ropes and cables

Type 989

1580 (1420) 2400 (2160) 3160 (2840) 4800 (4320) 6300 (5680) 7900 (7100) 9500 (8520)

1000 1000 1333 2000R 2666R 5000R 4000R

12 15 15 15 15 12 15

Textile finish; fiber optic cable reinforcement

Kevlar 68 (DuPont, 1992b, 1992c, 19938) 215 (195) 90 Type9568 1000 1580 (1420) Type9898

420 (380) 1580 (1420) 2400 (2160) 3160 (2840) 4800 (4320) 7900 (7100)

267 1000 1000 1333 2000 5000R

(0.48) (0.59) (0.59) (0.59) (0.59) (0.48) (0.59)

12 (0.48) 12 (0.48)

High performance mechanical rubber goods

12 12 15 15 15 12

Textile finish; fiber optic cable reinforcement

(0.48) (0.48) (0.59) (0.59) (0.59) (0.48)

Kevlar 129 (DuPont, 1990, 1993h) Type 956E 1670 (1500) 1000

12 (0.48)

Power transmission belts, high-performance tires, high fatigue applications

Kevlar 129 (DuPont, 1992c, 1993h, 1993i) 666 Type 956C 1110 (1000)

12 (0.48)

Mechanical rubber goods

12 12 12 12

Personal body armor

Type964C

830 (750) 930 (840) 1110 (1000) 1580 (1420)

500 6OOL 666 1000

Kevlar 249 (DuPont, 1992c, 19938) Type965A 420 (380) 267 768 1270 (1140) 1000 1580 (1420) Type 968A

1270 (1140) 1580 (1420) 4730 (4260) 7890 (7100)

768 1000 3000R 5000R

(0.48) (0.48) (0.48) (0.48)

12 (0.48) 12 (0.48) 12 (0.48)

Woven reinforcement in aerospace composites, hard ballistic armor, printed circuit boards

12 12 12 12

No finish; marine composites, fiber optic cable reinforcements, ropes, filament-wound composites

(0.48) (0.48) (0.48) (0.48)

Kevlar HT (DuPont, 19938) Type 964C 1110 (1000)

666

12 (0.48)

Advanced ballistic protection

Kevlar K M 2 (DuPont, 1992d) 945 (850)

560

12 (0.48)

Ballistic protection: helmets, composite armor Continued on next page

Ararnidfibers 219 Table 10.6 Continued

Product (reference)

Count dtex (den)

Filameizf


number/ diameter Fnz ( l C 3 i n ) yarn

Twarond (Akzo, 1990,1991; DeCos, 1993) Type 1000

420 (380) 840 (760) 1100 (990) 1260 (1130) 1680 (1510) 2520 (2270) 3360 (3020)

250 12 (0.48) 500 12 (0.48) 750 10.5 (0.41) 750 12 (0.48) 1000 12 (0.48) 1500R 12 (0.48) 2000R 12 (0.48)

Standard finish; multipurpose

Type1001

420' (380) 840 (760) 1100' (990) 1260 (1130) 1680 (1510) 3360 (3020)

250 12 (0.48) 500 12 (0.48) 750 10.5 (0.41) 750 12 (0.48) 1000 12 (0.48) 2000R 12 (0.48)

Adhesive-activated finish; tires, mechanical rubber goods, composites

Type 1010

1680 (1510) 3360 (3020)

1000 2000R

12 (0.48) 12 (0.48)

Very low finish level; composites

Type 1020

1680 (1510)

1000

12 (0.48)

Special finish for increased abrasion resistance; cables, ropes, nets

Type 1030 17 OOO(15 300)

5000R

12 (0.48)

PTFE + silicone oil impregnated; braided packings

Type 1031 14 OOO'(12 600)

5000R

Type1040

420 (380) 840 (760) 1100' (990) 1260 (1130) 1680 (1510)

250 500 750 750 1000

Type 1041

1260' (1130) 1680' (1510)

750 1000

12 (0.48) 12 (0.48)

Adhesive-activated finish; fabrics

Type2000

930 (840) 1110 (1000) 1680 (1510)

1000 1000 1000

6.6 (0.26) 8 (0.31) 12 (0.48)

Standard finish; high tenacity for ballistic applications. 930 dtex fiber is 'microfilament'.

Twaron HM (Akzo, 1991) Type 1055 1210 (1090) 1610 (1450) 2420 (2180) 3220 (2900) 4830 (4350) 6440 (5800) 8050 (7245) Type 1056

1210 (1090) 1610 (1450) 2420 (2180) 6440 (5800) 8050 (7245)

12 (0.48) 12 12 10.5 12 12

(0.48) (0.48) (0.41) (0.48) (0.48)

PTFE + silicone oil impregnated; braided packings Tangled yarn; multipurpose

750 1000 1500R 2000R 3000R 4000R 5000R

11.5 11.5 11.5 11.5 11.5 11.5 11.5

(0.45) (0.45) (0.45) (0.45) (0.45) (0.45) (0.45)

Standard finish; multipurpose

750 1000 1500R 4000R 5000R

11.5 11.5 11.5 11.5 11.5

(0.45) (0.45) (0.45) (0.45) (0.45)

Very low finish level; composites


220 Organicfibers Table 10.6 Continued

Product (reference)

Count

Filament


-

dtex (den)

number/ diameter yarn pm (1Win)

Twaron IM (Akzo, 1991) Type 1111

420' (380) 1260 (1130) 1680 (1510) 2520 (2270)

250 750 1000 1500R

12 12 12 12

(0.48) (0.48) (0.48) (0.48)

Easily removed finish; fiber optic cable reinforcements, ballistics, composites

-

-

-

Twisted 48 t/m; multipurpose

-

-

-

-

-

-

-

Type A1 lubricating finish Lubricating finish on 'acidic' fiber Lubricating finish on 'neutral' fiber Lubricating finish on 'acidic' fiber Type A1 lubricating finish on 'acidic' fiber Two different heat treatments available Type A1 lubricating finish on 'acidic' fiber

Armos (Kaiser VIAM, 1993a) 588 (530) SVM(Kaiser VIAM, 1993b-j) 63 (57) 143 (130) 294 (265) 294 (265) 294 (265) 588 (530) 588 (530)

-

-

-

-

Technoru' (Teijin, 1989; Mahn 1993) T-200

1110 (1000) 1670 (1500)

666 1000

12 (0.48) 12 (0.48)

Rubber reinforcement

T-202

440 (400) 1670 (1500)

1667 1000

12 (0.48) 12 (0.48)

Rubber reinforcement, pre-activated type

T-220

1110 (1000) 1670 (1500)

666 1000

12 (0.48) 12 (0.48)

Rope, cable, and cord

T-221

1110 (1000) 1670 (1500)

666 1000

12 (0.48) 12 (0.48)

Rope, cable, and cord

T-230

1670 (1500)

1000

12 (0.48)

Fiber-reinforced plastics, rope

T-240

60 (55) 110 (100) 220 (200) 440 (400) 1110 (1000)

36 67 133 267 666

12 12 12 12 12

(0.48) (0.48) (0.48) (0.48) (0.48)

Woven and knitted fabrics, fiber-reinforced plastics

T-241

1670 (1500) 8330 (7500)

1000 5000R

12 (0.48) 12 (0.48)

Woven and knitted fabrics, fiber-reinforced plastics

T-360

608 (55) 220 (200) 440 (400) 1110 (1000) 1670 (1500)

36 67 133 267 1000

12 12 12 12 12

'Spunnized' yarn (made up of long but not continuous filaments) for protective clothing and other fabric applications

T-370

220 (200) 440 (400)

133 267

For footnotes see next page

(0.48) (0.48) (0.48) (0.48) (0.48)

12 (0.48) 12 (0.48)

High tenacity 'spunnized' yarn for reinforcement of rubber, etc.

Aramid fibers 221 Table 10.7 Sources of information o n commercial aramid fibers

Information source

Product Armos and SVM fibers

Kaiser VIAM; 880 Doolittle Drive, San Leandro, CA 94577, USA

Kevlar fibers

DuPont Fibers; P.O. Box 80705, Wilmington, DE 19880-0705, USA, (800)

4-KEVLAR Technora fibers

Teijin Limited, 11, 1-chome, Minamihonmachi, Chuo-ku, Osaka 541, Japan Teijin America Inc; 10 East 50th Street, New York, NY 10022, USA

Twaron fibers

Akzo, Aramide Maatschappij v.o.f., P.O. Box 9300,6800 SB A r h e m , Westervoortsedijk 73, The Netherlands Akzo Fibers Inc., 801-F Blacklawn Rd., Conyers, GA 30207, USA

The outstanding toughness of aramids is often the reason they are used over cheaper, stiffer or even stronger fibers. Unlike glass and carbon composites, aramid composites loaded in compression, flexure or shear fail in a non-brittle manner, with significant work being required to fail the composite. Their fatigue resistance is also excellent. If other concerns such as cost or stiffness preclude the use of aramid composites, aramids are often used as a hybrid with another fiber to improve the toughness of the composite. The poor off-axis and compressive properties of aramid fibers must be considered in any design. However, because of their high strength in axial tension and their toughness, aramid fibers are often outstanding in applications such as pressure vessels where the loading is almost totally in longitudinal tension. Aramid fibers absorb moisture. Where either the physical swelling of the fiber or the amount of moisture absorbed is of significant concern, one of the lower absorption aramids, such as Kevlar 149, Armos, or Technora should be considered.

Aramids are strong UV absorbers and deteriorate when exposed to ultraviolet light. Protective coatings or the self-screening ability of the fiber should be used to avoid deterioration. Aramid fibers are opaque and thus the penetration of resin into the fiber bundles cannot be determined visually for a aramid composite as it can for those made with glass fibers. In fabric applications the weave used is important to the resulting properties. The same is true for sandwich construction. In these cases, the fiber, fabric, or honeycomb supplier can provide design assistance. The choice of resin system for use with aramid fibers is an important one. Epoxy resins give better translation of fiber properties than do polyesters, producing better shear strength and flexural properties, but lower impact resistance. Vinyl ester resins give both good shear strength and impact resistance. Thermoplastic matrices are also used, particularly in chopped fiber composites, because of their improved impact resistance over thermosets. However,

Footnotes for Table 10.6 All availabilities are subject to market conditions and should be verified with the manufacturer. All Kevlar fibers are supplied untwisted. E The ' R indicates that this fiber is a 'roving,' meaning in this case that it is composed of more than one 'end' of yam. Twaron fibers are normally supplied untwisted. In some circumstances twist may be supplied on special request. e Under development. Technora fibers are supplied with twist as requested. Finishes are supplied as requested or as is appropriate to the application. For special applications,Technora fibers can be supplied in larger than 12 pm filament diameters. These fibers, and others 'spun' yarns (composed of discontinuous filaments) are normally measured by '(English) cotton count' (ECC) rather than dtex or denier, where ECC = 5315/denier. a

'

222 Organicfibers Many of these are not as structural composites. For example, aramids are used in many rope and cable applications. In mooring ropes to secure oil tankers and to anchor off-shore oil platforms, the lighter weight compared to steel makes the aramid ropes much easier to handle. In addition, they do not corrode, are easier to maintain and have an extension under load which is far superior to both steel and other organic fibers. Aramids are widely used to reinforce mechanical rubber goods. The largest volume of such usage is in pneumatic tires, where aramids are lighter than steel and offer higher strength and modulus than other organic fibers. Significant usage is also seen in belts and hoses. The excellent fatigue and creep resistance of aramids are important factors in their usage in these applications. Corrosion resistance and electrical resistivity may also be important. Aramids are also used in athletic shoes and in rubberized sheet materials as used in aircraft evacuation slides and life rafts. In some cases, non-composite applications have led to composite uses. For example, aramids have long been used in soft body armor, where the fibers absorb and disperse bullet impact energy to other fibers in the fabric weave. This application has now seen a derivative usage in rigid composite ballistic armor, composite helmets and composite spa11 liners. In these applications the toughness, RF transparency and fire and corrosion resistance of aramid fibers were significant factors in their selection. In spite of significantly higher fiber costs than glass, aramids are used in canoes, kayaks, racing shells and small boats where maximizing strength and minimizing weight are important. Aramids offer weight savings for superior speed and better handling and/or improved range and fuel economy. Toughness and overall durability and vibrational damping are also superior with aramids. The 10.2.8 APPLICATIONS superior properties of aramids allow boats to Aramid fibers are used in numerous applica- be built at an overall cost only 10-15% higher tions, some of which are listed in Table 10.6. than with glass fibers and with superior perfor-

for thermoplastics the penetration of the resin into the fiber bundle and the quality of the fiber-matrix bond is almost always of concern. Because aramids are very tough fibers, they are somewhat difficult to cut and their composites can be difficult to machine. Special shears and other tools are available for cutting aramids and many successful machining techniques have been developed. The fiber manufacturers are an excellent source of information in this area. As with all high performance fibers, aramids should be handled with care before and during processing. Rough handling will damage any high performance fiber. In addition, because of their sensitivity to ultraviolet light, aramids should be protected from such exposure. The fibers also should not be exposed to excessive moisture prior to processing. If the fiber is to be twisted, braided, or woven, it is preferable to condition the fiber for one to two days at room temperature and intermediate moisture content prior to processing (DuPont, 1993h). However, if the fiber is to be resin-impregnated and processed directly into a composite, so long as fiber handling is careful, superior properties may be attained by drying the fiber prior to processing. Tlus is because of improved bonding of resin to the filament surfaces. Aramid fibers present minimal safety or environmental concerns. In lifelong animal inhalation studies with Kevlar fibers, no health effects were observed at any workplace levels. Nonetheless, as with any textile fiber, inhalation of fibrous particles should be avoided. Extensive animal and human skin patch tests with Kevlar fibers have shown no sensitivity and little irritation, and rat feeding studies have shown oral toxicity to be very low. Combustion by-products are similar to wool. Aramid yarns are also essentially inert in the environment (DuPont, 1993h).

Extended

h i i i

polyefhylcviefibers 223

mance (DuPont, 1983). These same properties properties but they also have limitations that have led to the use of aramids in skis. must be considered in design. Their high strength-to-weight ratio comCommercially available high strength, h g h bined with outstanding toughness has led to modulus polyethylene fibers include Spectrakh' numerous applications of aramids in aero- fibers from Allied-Signal Corporation, space. In both civilian and military aircraft, the DyneemaO SK60 from Dyneema Vof, Tekmilon" toughness of aramids - and resulting resis- from Mitsui Petrochemicals and a new, as yet tance to damage from impacts ranging from unnamed, fiber from Hoechst Celanese. bird strikes to shrapnel - insures their continued usage. Engine nacelles and the tail cone on 10.3.2 MANUFACTURE the McDonnell Douglas DC-9-80 are made from Kevlar composites and approximately The traditional method of producing fibers 10% of the empty airframe weight of De from polyethylene is to spin them from a polyHavilland Aircraft's DASH-8 turboprop com- mer melt. This technique yields fibers muter aircraft is Kevlar composite. Aramid composed of folded-chain crystalline regions composites are also widely used in rotorcraft with non-crystalline regions interspersed. With and other vertical lift aircraft. extraordinary means, the modulus of the absolute best of such fibers can be brought to about 80 GPa (11.5Msi). It was long recognized, 10.2.9 CONCLUSIONS however, that if polyethylene could somehow Although composites of other fibers have now be produced with extended chain crystallinity, supplanted aramid composites as having the a very high modulus fiber would result. [The highest specific strengths, aramids still offer theoretical modulus for polyethylene is combinations of properties not available with 320 GPa (46 Msi) (Adams and Eby, 1987).] any other fiber. For example, aramids offer Following earlier work by Pennings, in the high specific strength, toughness and creep late 1970s Smith and Lemstra of DSM (The resistance, combined with moderate cost. Netherlands) developed a process with comHowever, the applications of aramid compos- mercial potential which yielded a highly ites continue to be limited by their poor oriented extended-chain polyethylene fiber compressive and off-axis properties and in (Hongu and Phillips, 1990).At the same time, some applications, their tendency to absorb both Toyobo Inc. of Japan and Allied Chemical water. Nonetheless, aramids will continue to Company in the USA were working on a simbe the fiber of choice where properties such as ilar approach. DSM, however, was the first to outstanding impact resistance combined with patent the process and both Toyobo and Allied judged it impossible to circumvent the basic creep resistance are critical. patent filed by DSM. Thus, both companies entered into technical association with DSM to 10.3 EXTENDED CHAIN POLYETHYLENE produce polyethylene fibers. Toyobo Inc. FIBERS linked with DSM to form the joint venture Dyneema Vof - to produce and market the 10.3.1 OVERVIEW new fiber. In the USA, Allied-Signal is licensed High performance polyethylene fibers, with from DSM/Stamicarbon to produce and maroutstanding strength-to-weight and stiffness- ket a similar fiber. The process which is used to produce most to-weight performance, show promise in various specialized applications. While such commercial high strength, high modulus polyfibers are not as widely known as aramid and ethylene fibers is called gel spinning, the name carbon fibers, they possess many superior derived from the gel-like appearance of the

224 Organicfibers as-spun and quenched fibers. Ultra-high molecular weight linear polyethylene is dissolved in a volatile solvent to form a dilute isotropic solution that is then spun through a spinneret and quenched in cold water to form a gel precursor fiber. Following solvent extraction, this fiber is then hot-drawn to a very high draw ratio (= 30), yielding a very highly oriented, highly crystalline, lightweight fiber (Dyneema, 1987; Jaffe, 1989; Ward and McIntyre, 1986; Yang, 1992). Another approach to producing a high strength polyethylene fiber is melt extrusion followed by multiple stage drawing of a much lower molecular weight polyethylene. The modulus of an experimental fiber of this type, 220 GPa (32 Msi), is the highest ever achieved for polyethylene (Adams and Eby, 1987). The new polyethylene fiber from Hoechst Celanese is the only commercial version of such a fiber. This fiber has only about 50% of the strength and 75% of the modulus of gel spun fibers. In this case, the expense of dealing with a volatile and potentially toxic solvent is avoided, lowering the overall price of the fiber significantly. 10.3.3 STRUCTURE

Fig. 10.8 Schematic illustrating the difference between (a) conventional polyethylene fibers and (b) gel-spun extended chain fibers.

Figure 10.8 illustrates the difference between conventional polyethylene fibers and gel-spun hydrogen bonds nor strong covalent bonds or melt-extruded and drawn extended-chain between them. They are, in fact, held together fibers. Figure 10.8(a)is a schematic of a con- by weak dispersion-type van der Waals bonds ventional melt-drawn polyethylene fiber. The which have a distinct effect upon properties. fiber consists of folded chain crystallites, mostly oriented in the draw direction, which 10.3.4 PROPERTIES are joined to one another by tie molecules and have between them interspersed non-crys- Polyethylene fibers offer a unique combinatalline material. Figure 10.8(b)is a schematic of tion of properties: low specific gravity, high a gel-spun and hot-drawn extended-chain specific modulus, high specific strength, high polyethylene fiber. Such fibers show minimal energy to break, high abrasion resistance, chain folding, high crystallinity and a very excellent chemical resistance, good ultraviolet resistance and low moisture absorption. They high degree of axial orientation (>95%). Since these fibers are based on polyethyl- have outstanding anti-ballistic and vibrational ene, they have a density of only two-thirds damping characteristics, as well as a low that of aramid fibers and about one-half that of dielectric constant. However there are tradecarbon fibers. However, the polyethylene crys- offs involved in the use of polyethylene fibers. tallites have neither relatively strong They are limited to fairly low use temperatures,

Extended chain polyethylene fibers 225 they produce composites with poor off-axis treated Spectra 900 fiber in an epoxy matrix and compressive properties and have poor was found to be -9 x lO"/OC (-5 x 104/OF) in creep resistance. the axial direction and 100 x lO"/"C (56 x As with aramid fibers, the anisotropy of 104/"F) in the transverse direction. The axial their microstructure gives polyethylene fibers thermal expansion coefficient of a similar comanisotropic mechanical, thermal and physical posite of Spectra 1000 fiber was -10 x lO"/"C properties which can be used to advantage in (-5.6 x 10"/OF) and the transverse coefficient some applications. was 105 x lO"/OC (58 x 10"/"F) (Allied, 1989). Polyethylene fibers are the only high performance fibers with a specific gravity of less than Physical and thermal properties 1 and thus are the only fibers that float. Their Polyethylene fibers have a relatively low melt- density is about two-thirds that of aramid ing point [147"C (297"F)I and thus a low use fibers and about half that of carbon fibers. temperature. In general, polyethylene fibers Polyethylene fibers will burn slowly if ignited, are limited to use below 100°C (212°F). They decomposing into carbon dioxide and water. will, however, tolerate brief exposure (30 min The filament diameters of commercial polyor less) at temperatures near the melting point ethylene fibers are relatively large, typically without major property loss (Dyneema, 1987; 23-38 pm (0.91-1.50 x in), although the Weedon and Tam, 1986). diameter of Mitsui's Tekmilon monofilament As would be expected from the lower melt- fibers can be as large as 121 pm (4.76 x in). ing temperature, the properties of polyethylene The filament cross-section is typically irregufibers are much more sensitive to temperature lar and somewhat elliptical. than are aramids. Like aramid fibers, polyethylene fibers contract with temperature Mechanical properties in the axial direction, while expanding in the transverse direction. The thermal expansion Gel-spun polyethylene fibers offer some coefficient of a composite of 60 vol% plasma- tremendous advantages over other fibers. As Table 10.8 Axial tensile mechanical properties of representative high performance polyethylene fibers

Fiber

Reference

Fiber type

Spec@ gravity

Tensile modulus, GPa (Msi)

Tensile strength, MPa (ksi)


-

Dyneema SK60

Dyneema, 1987 gel-spun

0.97

87 (12.7)

2620 (380)

-

Hoechst Celanese fiber

Hoechst meltCelanese, 1993 extruded

0.96

1300 (189)

4

Spectra 900

Allied, 1993

gel-spun

0.97

55 (8.0) 86-103 (12.5-14.9)

2080-2400 (300-350)

3.6-3.7

Spectra 1000

Allied, 1993

gel-spun

0.97

128-171 (18.6-24.8)

2740-3000 (397435)

2.8-3.1

Tekmilon monofilament

Mitsui, 1989

gel-spun

0.96

59-98 (8.6-14.2)

1470-3430 (213498)

4-6

Tekmilon multifilament

Mitsui, 1989

gel-spun

0.96

88.3 (12.8)

2450 (356)

3

226 Organicfibers

can be seen in Fig. 10.4, these fibers offer very high specific stiffnesses and specific strengths, equivalent or superior to all of the aramid fibers and to most of the carbon fibers. This superior performance is offered at a lower price than that of competitive fibers. Table 10.8 compares the mechanical properties of representative commercially available polyethylene fibers. Like aramid fibers and for similar reasons, polyethylene fibers have poor compressive and off-axis properties. Since the fiber is held together internally by only very weak van der Waals bonds, the transverse strength of the fiber is even worse than that for the aramids. In addition, the inertness of the polyethylene fiber means that the untreated fiber bonds

very poorly to a matrix. Although gas plasma surface treatment can improve the interfacial bond strength significantly, polyethylene fiber composites will still have poor off-axis properties. Table 10.9 gives mechanical properties for Spectra fiber composites, including those made from plasma-treated fibers. In spite their weak transverse strength, but because of the non-stick nature of polyethylene and thus its low coefficient of friction, polyethylene fibers perform much better than aramids in abrasion resistance and polyethylene fabrics are much less easily damaged than are those of aramid fibers. The abrasion resistance of polyethylene fibers can be up to ten times that of aramids (Dyneema, 1987)and can be improved

Table 10.9 Mechanical properties of Spectra polyethylene fiber composites” (Allied, 1989)

Matrix: Bisphenol A based epoxy __

-

Spectra 900 Axial Tensile Properties: Volume percent fiber Modulus, GPa (Msi) Strength, MPa (ksi) Elongation, % Axial Compressive Properties: Volume percent fiber Modulus, GPa (Msi) Strength, MPa (ksi) Elongation, YO Flexural Properties: Volume percent fiber Modulus, GPa (Msi) Strength, MPa (ksi) Short Beam Shear Properties: Volume percent fiber Strength, MPa (ksi) a

58 27 f 1 (4.0 f 0.1) 552 zk 90 (80 f 13) -

70 32 f 5 (4.7 k 0.7) 52 k 2 (7.5 f 0.3) -

~

Spectra 900 P T ~ Spectra 1000

Spectra 100 PT

-

54 50 f 3 (7.2 f 0.5) 889 f 55 (129 f 8) 2.1 i-0.4

50 24 f 1 (3.5 f 0.1) 676 f 103 (98 15) 3.6 f 0.2

53 50 (7.3) 1034 c 228 (150 f 33)

70 40 k 5 (5.8 + 0.7) 59 + 1 (8.6 c 0.2)

55 19c6 (2.7 f 0.9) 72 f 3 (10.5 f 0.4) 3.8 k 0.5

65 54 f 3 (7.8 f 0.4) 69 f 1 (10.0 f 0.2) 3.8 f 0.2

*

-

-

58 22 k 1 (3.2 f 0.2) 145 f 7 (21 k 1)

54 30f1 (4.3 f 0.2) 200 f 7 (29 f 1)

54 23 zk 1 (3.3 k 0.2) 159 +. 7 (23 f 1)

53 38 f 3 (5.5 f 0.5) 214 f 7 (31 k 1)

58 8.3 f 0.7 (1.2 f 0.1)

54 28.3 f 0.7 (4.1 f 0.1)

54 9.0 c 0.7 (1.3 f 0.1)

53 21 e 3 (3.1 f 0.4)

Numbers of specimens tested and criteria for limits not specified. PT indicates a fiber with gas plasma surface treatment.

Extended chain polyethylene fibers 227 even further by the use of lubricants. Because of their high strength, polyethylene fibers exhibit very high energy to break. On a per-weight basis, the impact energy absorption of polyethylene composites is superior to that of all other fiber composites. Polyethylene fibers are more affected by temperature than are higher melting point fibers. The loss in modulus as function of temperature is shown in Fig. 10.9 for Tekmilon multifilament fiber and Spectra fibers. Fig. 10.10 shows the loss in strength as a function of temperature for Tekmilon multifilament, Spectra 900 and Spectra 1000 fibers. Because of their very high specific strength at room temperature, however, polyethylene fibers still outperform most other fibers to about 100°C (212°F). Room temperature strength retention of polyethylene fibers following annealing at temperatures of up to 125°C (260°F) is excellent, while modulus loss following such

r L 1 . 8 8

100

0 a

I

,

I

I

I

I

,

exposure is 20-30%. The loss in both modulus and strength are reduced if annealing is performed under tensile loading. Unlike aramid fibers and their composites, polyethylene fibers and composites show very little or no loss of properties, axial or off-axis, when exposed to moisture. Creep resistance of extended-chain polyethylene is of concern. Because of its low melting temperature, the resistance of the fiber to creep, even at room temperature, is less than ideal. This is significant, since the creep of carbon, glass and aramid fibers is minimal. Spectra 1000 is a 'stabilized' version of the fiber, which shows better creep resistance than the Spectra 900 fiber. Figure 10.11 shows the creep response of the two Spectra fibers at room, elevated and low temperatures. At low load levels at room temperature and/or at low temperatures the creep encountered is not severe, especially for the Spectra 1000 fiber, but at higher loads or temperatures the creep is much

,

I

S

*

- 15

-

.-ul

75 -

- 10

c)

i

ul

-33

= -3

50 -

U

0

z

I

- 5 25

0

-

~

~

.

l

l

~

l

.

~

'

~

~

~

0

~

l

~

~

~

~

I

Fig. 10.9 Modulus as a function of temperature for Spectra 900, Spectra 1000, and Tekmilon multifilament polyethylene fibers (Prevorsek, 1989; Mitsui, 1989).

*

l

l

l

l

228 Organicfibers Temperature,

200

150

100

50

OF

250

300 500

-

3000

- 400

-

2500 0

%

-

2000 1

.-

300 -$

f0,

f

VI

5

1500

1

1000

-

500

-

C

i7l

50

25

0

100

75

Temperature,

125

e

-

200 3;

-

100

0 150

OC

Fig. 10.10 Strength as a function of temperature for Spectra 900, Spectra 1000, and Tekmilon multifilament polyethylene fibers (Allied, 1991e; Mitsui, 1989).

-

Spectra 900

.___-0

------RT, 10% Load

~

0

~

10

20

30

40

50

60

70

80

90

100

time, h

1300 1200 1200 1000 950 800

1400 1100 1350 1120 1000 1010

>1400 1200 1400 1200 1250 1190

of a silicon carbide based textile fiber for composite reinforcement. All commercially available fibers in this category contain oxygen but can also contain nitrogen and titanium. Nicalon fiber manufactured by Nippon Carbon Company and marketed in

Constituent materials 311 the USA by Dow Corning Corporation is by far the most commercially developed. X-ray diffraction analysis indicates that Nicalon consists of ultra fine p-Sic particles dispersed in a matrix of amorphous SiO, and free carbon6. Nicalon has excellent resistance to thermal degradation in argon and air exposure at temperatures to 1000°C for as long as 100 h7.The loss of tensile strength for Nicalon by exposure to temperatures to 1400°Cin both air and a r m n are nrewnted in FiP 14 7

- 3 -

0"

t

9. b

d

g 2 g!

r

n

-

Monofilament reinforcements Monofilament Sic and boron fibers are produced by chemical vapor deposition onto a fine substrate filament. For the case of Sic fibers the core is 37 pm amorphous carbon filament, while for boron a 13 pm tungsten wire has been commonly used. The principal advantage of monofilament reinforcements are their ability to tolerate some degree of surface reaction with the matrix during fabrication or high temperature service. These fibers can be infiltrated by a number of processing methods including powder sintering, powder hot pressing, plasma spraying and melt infiltration. These fibers are limited to structures with relatively simple shapes such as sheet, plates and large diameter cylinders because of their large critical bend radius. Table 14.4 lists the properties of these fibers.

Q,

Table 14.4 Properties of monofilament reinforcements

e l -

f

'Original fiber 0

I

I

I

I

I

I

I

I

.

Manufacturer Composition Tensile strength (MPa) Tensile modulus (GPa) Density (g ~ r n - ~ ) Diameter (pm) Critical bend radius (mm)

(4

0

0

Boron

SCS-6

Textron B 2.5 400 2.5 140 11

Textron Sic 4.3 427 3 140 7

14.2.2 MATRIX MATERIALS

0'O\14000C

'0

0 1

10

1d

1o3

Heat treatment time, t (h)

(b)

Fig. 14.3 Loss of strength of NicalonTMafter exposure to (a) argon and @) air at temperature to 14OO0C7.

The selection of matrix materials for ceramic composites is strongly influenced by thermal stability and processing considerations. The properties of matrix materials commonly used in ceramic composites are shown in Table 14.5. These include oxides, carbides, nitrides, borides and silicides. The first indication of the ability of a material to resist high temperature service is melting temperature. With the exception of glass ail these materials have melting temperatures above 1600°C. As the melting temperature increases the ease of processing decreases.

312 Ceramic composites Table 14.5 Properties of typical ceramic matrix materials ~~

Materials

Young’s modulus (GPn) -.

LAS Pyrex Alp, Mullite ZrO, PS ZrO, FS

TiO, Si,N, SN Si,N, RB Si,N, HP SiOz Sic Sn Sic HP B4C TiB, Tic TaC Be0

wc

Cr $4 Cr& BNL BNII

NbC

117 48 345 145 207 207 283 310 165 310 76 331 414 290 552 427 283 359 669 103 386 34 76 448

Poisson’s Modulus of rupture ratio (MPa) __

0.24 0.20 0.26 0.25 0.23 0.23 0.28 0.24 0.24 0.24 0.16 0.19 0.19 -

0.20 0.19 0.24 0.24 0.20 -

0.20 -

0.21

138 55 483 186 648 248 83 496 303 827 -

386 462 310 896 248 200 234 -

262

Thermal expansion

Fracture toughness (MPa m”?)

Density

2.42 0.08 3.52 2.20 8.46 2.75 2.53 5.60 3.41 5.60 0.77 4.94 4.94

2.61 2.23 3.97 3.30 5.75 5.56 4.25 3.18

5.76 3.24 8.64 5.76 7.92 13.5 9.36 3.06

-

-

-

6.92 -

-

3.85

-

-

76 110

-

-

-

-

(g

Me1t ing point (“C)

(10-6PC)

3.19 2.20 3.21 3.21 2.41 4.62 4.92 14.50 3.00 15.80 5.21 6.70 1.94 1.94 7.82

3.06 0.54 4.32 4.32 3.06 8.10 8.46 6.66 5.76 4.50 7.56 9.67 6.66 0.36 6.66

-

1252 2050 1850 2760 -

1849 1870 -

1870 1610 1980 1980 2350 2900 3140 3880 2530 2870 2435 1890 2982 2982 3499

Mechanical and chemical compatibility of the 14.3 PROCESSING METHODS matrix with the particular reinforcement ultiProcessing of ceramic composites can be permately determines whether a useful formed by solid, liquid, or gas phase composite can be made. For the case of processing of the matrix material to achieve whisker reinforced composites the chemical infiltration of the matrix around the reinforcreactions with matrix are particularly critical ing phase. The goals in processing ceramic since even minor reactivity can consume the composites are to achieve minimum porosity entire reinforcement. Large differences in the with a uniform dispersion of the constituents coefficients of thermal expansion between and controlled bonding between the reinforcreinforcement and matrix can lead to large ing phase and the matrix. residual stresses during the fabrication and ultimately result in serious degradation of mechanical strength. Small or optimum differ- 14.3.1 POWDER PROCESSING ences can be beneficial to mechanical Fundamental steps in processing ceramics properties by placing the weaker constituent composites from powdered constituents are: in compression or by inducing crack deflection 0 powder selection; between reinforcements. 0 powder characterization; 0 agglomerate reduction;

Processing methods 313 constituent mixing; green body fabrication; green dressing (machining and gate removal); binder removal; consolidation and densification; final dressing (burr removal); inspection.

To minimize voids and interfacial weakness and maximize the toughening effect of the reinforcing phase, a uniform finely dispersed mixture must be produced. Arranging the constituents to minimize free space between them is referred to as ’packing’. When the constituents are not effectively packed, subsequent densification becomes difficult, requiring higher pressing temperatures, presThe selection of constituent powders is the sures and duration. Both constituent shape first step in composite design and consideraand particle size difference can affect packing. tion must be given to chemical, mechanical Optimum packing occurs when the particle and thermoelastic compatibility between the size distribution contains 30 vol. % of very constituents as well as the desired final small particle and 70 vol. % of large particles’. mechanical and physical properties of the If uniform round fibers (or whiskers) are percomposite. In addition to the obvious problem fectly aligned in a closed packed array then of reaction between constituents, other incommatrix particles approximately 0.15 times the patibilities such as large differences in melting fiber diameter would theoretically fit in the temperature of matrix and reinforcement can interstices. preclude successful processing. Thermal Most ceramic powders can be comprised of expansion mismatch between constituents can a mixture of primary particles and agglomercause premature failure in them or at their ates. Agglomerates are primary particles interface. bonded by surface chemical forces, electrostaThe rational selection of constituents usutic forces or solid bridging. In order to produce ally requires knowledge of certain physical a finely dispersed, homogeneous mixture of characteristics such as particle size distribumatrix and reinforcement successfully, the tion, shape, specific surface area, bulk density, agglomerates must be reduced. A typical electrical charge, impurities, etc. The ultimate agglomerate (mullite in this case) is shown in aim of such constituent characterization is to Fig. 14.5. The agglomerate is 8-9 ym in diamepredict the final characteristics of the ceramic ter while the constituent primary particles composite, as shown schematically in range from 0.1 to 1.5 pm. To uniformly incorFig. 14.48. porate 0.5-1 ym diameter whisker or particles,

MILL TIME

LUBRICANTS

Fig. 14.4 Use of powder characterization in process control8.

314 Ceramic composites

Fig. 14.5 Typical agglomerate found in mullite powders.

the large agglomerates must be broken down by mechanical action or chemical treatments if the agglomeration is due to surface forces. If the whiskers are robust or some degree of whisker breakage can be tolerated, both agglomerate reduction and constituent mixing can be accomplished simultaneously by ball milling. Organic binders are usually mixed with the particle-whisker mixture for near-net-shape processing by a variety of cold forming operations including uniaxial pressing, cold isostatic pressing, tape casting, extrusion, compression molding and injection molding. The ceramic preform after cold consolidation is referred to as the 'green' form. The part in the green form can usually be machined without damage. In this state additional near-net-shape processing can be applied such as gate removal and machining. Final consolidation and densification is performed at high temperatures. Three of the most common methods are sintering, hot (unidirectional) pressing and hot isostatic pressing. For low fiber or whisker contents

(5 % or lower) sintering may produce satisfactory results. For complete densification of even low fiber volume fraction composites, sintering may impractical due to excessive temperatures and durations. For high fiber or whisker volume fractions, hot pressing and hot isostatic pressing are the only effective methods for densification. Table 14.6 shows the effect of hot pressing time and temperature at 31 MPa pressing pressure on the theoretical density of Sic whisker-Al,O, composites for various vol.% whiskers. Theoretical densities of over 99 vol.9'0 can be achieved in unreinforced A1,0, at 15OO0C, in lO%SiC whisker composites at 1650°C and in 2O%SiC whisker composite at 1800°C10. 14.3.2 LIQUID PROCESSING

When high temperatures and mechanical forces are used to consolidate composites from the powder, the optimum strength properties can be sacrificed. Reducing processing temperature, time and pressure can minimize damage to the reinforcements but fully dense

Processing methods 315 Table 14.6 Effect of processing parameters on the theoretical densities of Sic whisker-Al,O, composites pressed at 31 MPa pressure'" Vol. %

Pressing temperature

Densify

("C)

Pressing time (mid

Density

zuh iskers

(g cnz-?)

(% theoreticnl)

0 10 10 20 20 20 20

1500 1500 1650 1500 1650 1725 1800

60 35 60 120 60 60 25

3.95 3.78 3.89 3.68 3.72 3.78 3.81

99.1 96.7 99.5 96.1 97.1 98.6 99.5

or near-fully dense composites cannot generally be produced. Processing by infiltration with a molten matrix would be an ideal way of minimizing mechanical damage and still achieve fully dense structures. The melting temperatures of ceramics used as matrices in composites limits the general use of melt infiltration as a viable processing route. However, by careful tailoring of the matrix and the use of innovative in situ reaction techniques, melt infiltration has been successfully utilized to fabricate ceramic composites. Glass and glass-ceramic matrices have been successfully infiltrated in the liquid form into fiber preforms by 'matrix transfer molding'".The high temperatures required to achieve the appropriate fluidity of the matrix limjts the available fiber-matrix compositions to only those with low mutual reactivity. Other matrix materials such as CaSiO,, SrSiO, and

SrO~Al,O;SiO, were infiltrated into Sic powder preforms with and without Sic whiskers with resulting open porosity of about 1%12. Recently considerable attention has been applied to directed melt gas-metal reactions which produce ceramic matrix composites directly from the liquid metal13.14. Both metal oxide matrix and metal nitride matrix composites have been produced by this technique. Net shape composites can be processed at temperatures of the melting temperature of the metal. The commercial development of this processes is called the DIMOXTMprocess of the Lanxide Corporation. In this process oxidation or nitridation occurs on the surface of the molten metal forming a layer of solid ceramic. The layer thickens as the molten metal wicks up between the grains of the ceramic. A schematic representation of the process is shown in Fig. 14.614.The phenomenon is made

Reinforcement preform Reinforcement preform entrapped in solid reaction product

Fig. 14.6 Directional metal oxidation method for processing ceramic composites4.

316 Ceramic composites possible by dopants which modify the surface energy between the phases. For instance, if the grain boundary energy, yB, is less than twice the energy of the solid-liquid interface, ysI.and the surface energy of the solid-liquid interface, ysL is greater than surface energy of the solid-vapor interface, ysv, thickening of the oxide (or other reaction compound) layer does not take place. By reversing the relative values of the surface energies, (i e, yB > 2ys1. and ysr < ysv) as illustrated in Fig. 14.7the unstable grain boundary permits wicking of the liquid metal through the grain boundaries of the reaction product phase'-'. Application of this technology to ceramic matrix components is achieved by allowing the reinforcement preform to float above the liquid metal bath for infiltration of the molten metal. A growth barrier can surround the reinforcement preform to produce practical net or near net shape component^'^, 16.

I

Molten Metal

I

14.3.3 VAPOR PROCESSING

The infiltration of the reinforcing phase by a gas that decomposes to form the solid matrix phase is generally referred to as 'chemical vapor infiltration' (CVI).Various carbides, nitrides, oxides and boridesI7 as well as unreacted carbon'* have been deposited on silicon carbide-based yam fibers (e.g. Nicalon and Tyranno fibers), oxide based fibers (e.g. Nextel fibers) carbon yam fibers and Sic whiskers. Silicon carbide is one of the most commonly applied matrices using CVI. Methyltrichlorosilane is reacted with hydrogen on the surface of the fiber to deposit silicon carbide. A typical reaction for this process is19. CH,SiC1,

+ H, + Sic + 3HC1 +H,

This reaction can take place by conventional chemical vapor deposition (CVD) at temperatures of 1000-1400°C. Silane-hydrocarbon (SiH,-C,HJ mixtures can be used to deposit Sic at temperatures below 500°C by plasma assisted chemical vapor deposition (PACVD). Table 14.7 lists some of the more commercially important matrix materials that can be applied by conventional CVD. A more complete list of ceramic materials produced by both conventional CVD and PACVD can be found in reference (20). Table 14.7 Ceramic materials formed by CVI processes2'

Matrix ceraTnic ~

I

Molten Metal

I

Reaction temperature ("C)

~~~

Tic Sic B,C TiN Si& BN A1N *1,0,

Fig. 14.7 Mechanism of directed metal oxidation growth14.(a) no growth due to stable grain boundary; (b) oxide growth mechanism with unstable grain boundary.

Reactant gases

SiO, TiO, ZrO, TiB, WB

Tic,-CH,-H, CH,SiCl,-H, BC1,-CH,-H, TiC14-N2-H, SiCl,-NH,-H, BC1,-NH,-H, AlCl,-NH,-H, AlCl,-CO,-H, SiH-C0,-H, TiC1,-H20 ZrC1,-C0,-H2 TiC1,BC13-H, WC1,-BBr,-H,

900-1000 1000-1400 1200-1400 900-1000 1000-1400 1000-1400 800-1200 900-1100 200-600 800-1000 900-1200 800-1000 1400-1600

Design considerations 317 The main drawbacks for processing composites by CVI are the high processing time and costs. Since the deposition occurs most rapidly on the outer surfaces, the internal passages can be blocked off long before full densification is complete. It is usually necessary to interrupt the infiltration process to grind the surfaces in order to reopen the gas access to the fibers or preform in the center of the part. Residual porosity of 10-20% with less than 10% open porosity are typically obtainedz1. Two basic methods of CVI are isothermal processing and forced flow/thermal gradient processing. In isothermal processing the fiber preform is heated by radiation from the walls of the furnace (so called ’hot wall reactor’) or by inductively heating a carbon mandrel on which the preform is placed. In both cases the decomposing gases are allowed to diffuse through the fiber preform. In the forced flow/ thermal gradient method the reactant

gases are forced through the fiber preform retained in a graphite holder with a sharp thermal gradient maintained by water cooling. A schematic diagram of the forced flow/thermal gradient method is shown in Fig. 14.8. 14.4 DESIGN CONSIDERATIONS

The approach to designing ceramic matrix is constrained by the brittle nature for both the matrix materials and reinforcements used. Unlike polymer matrix composites and even metal matrix composites, the rationale for design of ceramic composites is to impart toughness in a structure that would have unacceptable toughness as a monolithic ceramic”. Design methods are unique to the form of the composite, depending on whether continuous unidirectional reinforcements, discontinuous reinforcements or multi-layer, multi-directional reinforcements are being considered. As a starting point in the design of continuous

Exhaust aas Heating element Perforated lid

---

Infiltrated preform Fibrous preform

Reactant gases Fig. 14.8 Forced flow thermal gradient method for CVI processingzo.

318 Ceramic composites strength over the range of reinforcement volume fractions will depend on the relative fracture strain, strength and stiffness of the constituents. The relative fiber and matrix stress-strain curves and strength prediction of a composite consisting of a high stiffness, high strength fiber in a lower stiffness, low strain to failure matrix is represented in Figs. 14.9(a) and 14.9(b),respectively. There are many such fiber-matrix combinations that have this relative behavior as the examination of Tables 14.2 and 14.5 will reveal. For instance if Nicalon is selected as the fiber then the selection of mullite, lithium alumino silicate (LAS) or Pyrex 14.4.1 DESIGN OF CONTINUOUS glass, for the matrix meets the requirement. UNIDIRECTIONAL REINFORCEMENT Selecting Sic monofilament produces this case COMPOSITES for almost all matrix materials listed in Table The Young’s modulus of unidirectional contin- 14.5 with the exception of TiB, and Tic. The uous fiber ceramic composites Ec is composite strength in such a system should satisfactorily predicted by rule-of-mixtures: increase at a rate predicted by the linear ruleof-mixtures based on the strength of the Ec = E , V , + Em V, matrix and the stress on the fiber at the fracture strain of the matrix, a;. At fiber fractions where E , and Em are the Young’s moduli of the to V,,,, failure of the matrix constitutes failure reinforcement and matrix respectively and V , of the composite. The behavior of the composand Vm are the respective volume fractions. ites with fiber fraction below Vcr,thave simple When there is a high bond strength between linear stress-strain behavior to failure. Above the fiber and matrix, prediction of composite this fiber fraction the matrix breaks before the

unidirectional ceramic composites, the rule-ofmixtures can be used to calculate elastic and thermoelastic properties of the composite. Strength properties of the ceramic matrix composites are poorly predicted by the rule-of-mixturessince flaw sensitivity and reinforcement-matrix bond strength are not addressed by these tecluuques. Rule-of-mixtures properties are less important in discontinuously reinforced ceramic composites since toughness is strongly controlled by the interfacial properties.

%

Strain

0

Vcrit

1

Fiber Fraction

(b) Fig. 14.9 Strength prediction for high stiffness, high strength fiber and a lower stiffness, low strain to failure matrix.

Design considerations 319 fibers. The fibers can retain the broken matrix X', is between the range given by in place before the fibers break at a higher load. Composite strength above Vcritdepends upon the fiber strength. A typical stress-strain curve for such a system above V,,, is shown where T is the interfacial shear stress and r is the fiber radius. The value of strain at the end of this process, E ~ isz3 ~ ~ ,

u)

u)

0

L

z V

Emu

E max

Strain Fig. 14.10 Stress-strain behavior for composite with high stiffness, high strength fiber and a lower stiffness, low strain to failure matrix.

schematically in Fig. 14.10. The elastic portion of the curve is followed by a serrated, constant stress portion induced by a matrix failure process. During this process the matrix continues to crack until the spacing between cracks,

The final linear rising portion is the curve is the elastic response of the fiber. Continuous fiber breakage and fiber pull-out can produce the pseudo-ductility of the final portion of the curve. There are many potential continuous fiber-matrix combinations in which the matrix has a higher elastic modulus than the fiber. For instance, matrix materials such as titanium diboride, titanium carbide, silicon carbide and alumina with most of the continuous fibers listed in Table 14.2 would have the relative stress-strain behavior shown in Fig. 14.11(a). The strength of predictions of such systems is shown schematically in Fig. 14.11@).In this case the predicted strength of the composite would decrease with increasing fiber fraction until a minimum fiber fraction, V,, is reached. This behavior is similar in appearance to the

sf"

(D

8

b

v)

Of

(b)

Fiber Fraction

Fig. 14.11 Strength prediction for a high strength fiber and a higher stiffness, low strain to failure matrix.

320 Ceramic composites case of a high modulus, low failure strain fiber transverse elastic moduli E, and E,, respecin a lower modulus, high failure strain matrix tively, composites with aligned short fibers can as is typical of many metal matrix composites. be made by using the Halpin-Tsai relation? However the cause of the minimum behavior is quite different. Below Vmh failure of the matrix still constitutes composite failure where the rule-of-mixtures strength is composed of the matrix ultimate strength, omuand the stress on the fiber at the matrix failure strain, a;. Unlike the case for the lower modulus matrix, the stress 0; is lower than the a,, increasing fiber fraction lowers the rule-ofmixtures strength. Above Vmi,the fracture of tL= 2 l / d , and tT = 2 the matrix no longer constitutes composite fracture as the fibers alone are able to carry the The coefficients of thermal expansion in the load after matrix failure. longitudinal and transverse directions, a, and The above description applies to composaT respectively, can be estimated fromz6 ites with a high fiber-matrix bond strength and neglects the effect of fracture surface a, = (a,E,V,+ amEmVm) / E, energy. These conditions are not typical in real aT = (1+ vf) a,V, + (1+ vm)amVm- aLvLT composites and the simple rule-of-mixtures predictions must be modified to account for where these effects. Aveston et aLZ4accounted for the VLT = vf + YmVm effect of the fiber-matrix bond strength on the matrix failure strain as follows: and vf and vm are the Poisson's ratio for fiber 6zTE,Vf 1/3 and matrix respectively. These calculations will E'f = EkVmrEc] usually overestimate the value of these properwhere r is the fracture specific fracture energy ties because of ineffective bonding between fiber and matrix and deviation from ideal fiber of the matrix.

v,

[

14.4.2 DESIGN OF DISCONTINUOUS REINFORCEMENT COMPOSITES

In polymer and metal matrix composites it is usually desirable to design the fiber lengths to exceed the critical length, 1, given by ufr/t to allow the fiber to carry its full load prior to fracture. In ceramic composites, fiber breakage is rarely the design goal. Instead the role of the fiber is to provide toughness by a combinations of fiber pull-out, crack deflection and crack bridging. Nevertheless, the designer may want to predict the elastic and thermoelastic properties of the discontinuous reinforcement composites' An estimate for longitudina1 and

Fig. 14.12 Microstructure of 20 vel. yo SiC-Al,O, composite fabricated by tape casting.

Design considerations

321

alignment. Figure 14.12 shows the in-plane laminate fracture theories must be employed. microstructure for a 20 vol% Sic whisker-aluExamples of material designs that can make mina composite fabricated by tape casting and use of laminated-composite concepts for hot pressing. This processing method promotes improved performance are illustrated in Fig. fiber alignment in the tape cast direction, how- 14.133z.The magnitude of the surface compresever there is still a considerable deviation from sive stress can be calculated from laminate the predominant fiber directionz7. theory. Figure 14.13(a) shows a laminate For randomly oriented fibers or whiskers the design intended to produce surface compreselastic modulus, E , of the composites can be sive stresses. In this design the layers toward estimated from the results of the Halpin-Tsai the mid-plane gradually increase in coefficient method using the empirical relation: of thermal expansion. The outer layers, con-

ET= (3/8)E, + ( 5 / 8 ) E , 14.4.3 DESIGN OF MULTILAYER, MULTIDIRECTIONAL REINFORCEMENT LAMINATES

The concept of a laminated composite is used effectively in the design of polymer-matrix composites to achieve the high degree of strength, elastic and thermoelastic tailoring. Polymer-matrix composites reinforced with either continuous or discontinuous fibers are fabricated by stacking layers with specific characteristics and orientations in a predetermined sequence to achieve desired mechanical or physical properties. As with polymer composites, the ceramic composite layer properties may be calculated using theoretical and semiempirical method^^*-^^ from the constituent properties such as the elastic modulus of the fibers and matrix respectively, E, and Em, the orientation factor, f, the volume fraction of the fiber, Vf, the fiber aspect ratio, l/d, and the coefficients of thermal expansion for the fibers and matrix, afand am.By selection of the sequence of layer orientations and compositions, various elastic, thermoelastic, strength, physical and chemical characteristics can be produced. Classical laminate plate the0ry3&~~ can be used to accurately predict the elastic and thermoelastic properties of laminated composites from the layer properties. The strength properties, on the other hand, cannot be readily determined by commonly used laminate failure criteria since fracture of these laminates is still strongly controlled by the presence of flaws. Modified

(a) Design with graded composition

I

(b) Design with toughening layers

Oxidation resistant layer Wear resistant layer High toughness core

(c) High temperature wear design

Fig. 14.13 Typical laminate design concepts for ceramic matrix composites. (a) with graded composition; @) with toughening layers; (c) for high temperature wear.

322 Ceramic composites taining increasing amounts of low-expansion whiskers generate compressive residual stresses as a result of the differential contraction during cooling after the high-temperature densification process. A major advantage of laminated-composite processing is that it provides the engineering flexibility to use innumerable material and property combinations that would be impossible with traditional methods involving thermal or chemical tempering. This concept also allows the use of non-equilibrium compositions for greater degree of stress profile variation. For instance, the depth and magnitude of the stress gradient can be independently controlled by selection of layer composition and properties. Maximizing the stress gradient by the introduction of a high-expansion material in the interior of the composite would be impossible by conventional chemical tempering but is quite feasible by lamination. Strengthening can also be achieved by rendering surface flaws ineffective through the introduction of a tougher ceramic layer below the surface (Fig. 14.1303)). This design mitigates surface damage in the outer layers by blunting the cracks when they reach the underlying toughened layer. This layer may contain whiskers, a toughened ceramic, or metallic particles. The use of a toughened ceramic layer as the outer layer would not be as effective since abrasion or impact could produce flaws through its entire depth, thus permitting the crack to propagate through the lower-toughness interior layers with minimum resistance. In addition to increased strength and toughness, high-temperature corrosion resistance can be designed into a composite material by using a corrosion-resistant layer on the exterior surface (Fig. 14.13(c)) and layers tailored for high-temperature strength in the interior. A similar concept may be employed for a material designed as a high-temperature heat exchanger by grading the interior layers for high thermal conductivity. Using composite laminate theory, a materials designer can tailor the grading to minimize the deleterious residual tensile

stresses that are likely in such a construction. Differences in elastic modulus and coefficients of thermal expansion for layers containing different volume fractions of reinforcing whiskers can be used to generate favorable residual stress patterns in fabricated laminates. The thermal stresses o,T,oy' and T~~ in each layer of the laminate at any position through the thickness, z, measured from the midplane, caused by the restraint of the neighboring layer can be determined by Hooke's

[:

]=

y:'

Alll A'l2 A' Af12A'22A':6 "16

Af26

{ q%v]

Properties 323

where t, is the thickness of the kth layer, Q, are the untransformed stiffness coefficients and al are the coefficients of the thermal expansion in the principal material directions. The thermal moments Mx, My and MXyare zero. The residual stresses, oLand oT,in the longitudinal and transverse directions respectively are

For this laminate geometry the residual shear stress, rLT= 0. The compressive residual stresses thus induced in the outer surface of the ceramic composite raise the fracture strength by that amounP3.

(a) Bridging

(b) Pull-Out

14.5 PROPERTIES

The principal objective in design of ceramic composites is to produce enhanced toughness and mechanical reliability. Various energy absorbing mechanisms are produced by the reinforcement depending on the relative thermal expansion coefficients, relative elastic moduli and interfacial bond strength between the reinforcement and matrix. In addition the size, shape, distribution and volume fraction of the reinforcement plays a strong role in controlling the effectiveness of the toughening.

(c) Deflection

14.5.1 MECHANISMS OF STRENGTHENING

The four principal mechanisms of toughening (crack bridging34,35, fiber pull-out, crack deflection and matrix microcracking) are shown schematically in Fig. 14.14. More than one of these mechanisms can be operative at the same time in a ceramic composite but there

(d) Microcracking

Fig. 14.14 Toughening mechanisms for ceramic matrix composites.

324 Ceramic composites is usually a dominant one depending on the resistance. A quantitative treatment of the constituent and interfacial properties. effects of crack deviation on toughness have Bridging, pull-out and deflection are most been provides by Faber and Evans3'. In certain composites the conditions can be effective where the fibers are generally aligned favorable to allow the stress field of the propnormal to the crack surface. In the crack bridging mechanism (Fig. agating crack to interact with the stress field 14.14(a))the fibers remain intact for some dis- around the reinforcements to produce local tance behind the crack front, thus restraining matrix cracking around the reinforcement. the crack opening displacement and reducing Maximum effectiveness of this mechanism the stress intensity at the crack tip. The energy requires a fine dispersion of many reinforceabsorbing processes include fiber fracture, ments as illustrated in Fig. 14.14(d). An fiber-matrix friction and elastic strain energy appropriate mismatch in thermal expansion of the fiber. Thus, strong fibers, high strain to coefficients between reinforcement and fracture fibers and strong fiber-matrix bond- matrix provides the local stress field around ing promote this mechanism. The toughening the reinforcement. An analysis of this mechaproduced by a uniform closure stress of the nism of toughening has been provided by bridging fibers has been estimated by Becher H ~ t c h i n s o n ~ ~ . Typical fracture in whisker reinforced et a1.35to be ceramic (SiC/Al,O,) shown in Fig. 14.15 has elements several mechanisms of toughening. dKc = (3fU[VfrEcG,/6(1 - v2)EfGi]l12 Whisker pull-out and crack defection are eviwhere Gm and Gi, are the strain energy release dent in this example.

"

between the fiber and the matrix. For fibers of critical length the increment in toughness is given as (see Reference 36, for example): AGc = V ,o f u r / 6 ~

Crack deflection mechanism (Fig. 14.14(c)) forces the crack to deviate out of the normal stress plane as it negotiates around the reinforcements. The driving force for this deviation is the residual stress distribution produced by the mismatch in thermal expansion between the fiber and matrix. Reinforcements with higher coefficients of thermal expansion than the matrix will cause the matrix to be in compression near the reinforcement. This state will tend to deflect a crack as it approaches the vicinity of the reinforcement which is a higher crack resistant area to a region of lower crack

Fig. 14.15 SEM photograph of fracture path in Sic whisker-Al,O, composite. 14.5.2 TYPICAL PROPERTIES

The selection of materials for component design requires accurate and reliable mechanical property data. Because of the large variations in processing characteristics and starting material forms, such data are generally sparse. The mechanical properties most signifi-

Properties 325 91 1 cantly affected by reinforcements in ceramics are fracture strength and fracture toughness. 8 Vaughn et ~ 1examined . ~ ~ the effect of processing temperature and reinforcement for alumina 7and Sic whisker/alumina composites. Table 14.8 shows that processing temperature and remforcement have little effect on elastic modulus, but a significant impact on strength, fracture toughness and work of fracture. The noncompliance with the rule of mixtures for Young’s modulus is evidence of the lack of bonding and hence load transfer between fiber and matrix. The toughening of various ceramic matrix materials with increasing Sic whisker content is shown in Fig. 14.1635. A compilation of fracture strength and frac0 0.1 0.2 0.3 0.4 ture toughness for Sic whisker (Silar-SC9) Whisker content (Volume fraction) /A1,0, composites is given in Table 14.92.The Fig. 14.16 Increment in fracture toughness of Sic fracture strength of the composites for the whisker composites with various matrices35.

Table 14.8 Mechanical properties of polycrystalline A1,0, and Sic whisker /A1,0, matrix composites3’ Alumina (1500°C) Young‘s 371 modulus (GPa) Fracture 456k40 strength (MPa) Fracture 3.3d.2 toughness (MPa m1/2) Work of 10 fracture om-’)

Alumina (1659°C)

(1900°C)

Composite (Silar SC-1)

380

375

375

393

385+18

2534

641+34

606+146

5.0~0.2

3.7d.1

8.7M.2

4.6d.2

39

67

21

20

Alumina

Composite (Tateho SCW-1-5)

Table 14.9 Room temperature strength and fracture toughness of Sic whisker (Silar-SC-9)/A120, composites (Adapted from Reference 2 ) Whisker content (Vol. %)

Fracture strength (MPa)

Fracture toughness (MPa mIn)

Xejerence

0

150 253 391 475 540 652 675 680 641 720 640 850

4.3 3.7 3.6 4.0 4.8 4.6 6.1 8.7 8.7 7.0 7.9 6.2

40 39 41 42 42 41 42 40 39 42 42 43

5 10 15 20 30

40

326 Ceramic composites same whisker content can vary fairly significantly among the different sources while the fracture toughness is relatively consistent. Above 30% whiskers some composites can exhibit either a strength or fracture toughness that actually decreases because of the decrease in homogeneity of the whisker distribution. The fracture strength and fracture toughness for Sic whisker reinforced mullite are shown in Figs. 14.17 and 14.18 respectivelp. These composites were fabricated by tape casting and the L-type designation indicates that the crack propagation direction is normal to the tape casting direction and the T-type designation indicates that the crack propagation direction is parallel to the tape casting direction. Greater strength and toughness are achieved for the case of fiber orientation predominantly normal to the whisker axis direction. The maximum strength and toughness for this system is at 40% whiskers. The effect of Sic (Silar S-9) whisker content on strength and toughness of Si,N, matrix composites is given in Table 14.1045.For this material the maximum toughness occurs at 30% wluskers while the strength maximum is 1000

I

I

I

I

I

I

8

20

30

40

50

Fig. 14.17 Fracture strength of Sic whisker-mullite composites with the crack propagating normal (Ltype) and parallel (T-type) to the predominant whisker directionM.

I

I

I

I

I

~

I

I

-

-

z 6Y

E 7

-

-

5-

-

-

2

26 '

Ib

io

' io ' ' 40 ' 5'0 ' SIC Whisker Fraction ( ~ 0 1 % )

Fig. 14.18 Fracture toughness of Sic whiskermullite composites with the crack propagating normal (L-type) and parallel (T-type) to the .predominant whisker direction&.

Table 14.10 Room temperature mechanical properties of Sic whiskers/Si,N, matrix composites45

Whisker content (Vol. Yo)

Fracture strength (MPa)

Fracture toughness (MPa mIn)

0 10 20 30

375 395 550 970

4.0 4.9 7.0 6.4

I

Sic Whisker Fraction ( ~ 0 1 % )

I

a

apparently above this fiber content.An alternative form of Sic whiskers was used to fabricate Si,N, matrix composites by Shalek et ~ 1 . 4 This ~. whisker is formed by the Vapor-Liquid-Solid p r o ~ e s s ~The ~ , effect ~ ~ . of whisker content and processing temperature on the elastic modulus, fracture strength and fracture toughness are shown in Figs. 14.19, 14.20 and 14.21 respectively. There is an apparent critical processing temperature above 1600°C to achieve maximum attainable strength and toughness properties in these composites. Continuous fiber ceramic matrix composites are more likely to obey the rule of mixtures prediction for elastic and strength properties

I I J I I I I I I I I I I IO

I

SiCw-mulIite

Sicw - mullite

2000

*

Properties

327

especially where minimum fiber damage and porosity occurs. The predictable increase in flexural strength for a Nicalon fiber/borosilicate glass composite is seen in Fig. 14.2249.

SIC Whisker -hobpressed Si3 N, matrix composite

SIC Whisker -hot-pressed Si3 N, matrix composite

13

300

: Hot-pressed at 1750°C B

: Hot-pressed

at 1850°C 1600°C

A : Hot-pressed at

-0

1 . . . . . . . . 1 5

10

IS

PO

25

30

35

40

Volume Yo S i c whiskers

Fig. 14.19 Elastic modulus of Sic whisker-Si,N4 composites processed at various pressing temperature~~~.

: Hot-pressed at 1750OC :Hot-pressed at 1850°C A : Hot-pressed at 1600°C 0

4

.

.

.

.

1

5

30

10

15

25

20

35

30

40

Volume % SIC whiskers Si, N, matrix composite 1750°C 1850°C . 1600°C

E3m-

*

03

"

50 -,p

:

t m.

8

A A

I

m

Rule of mixtures prediction for SiC/Borosilicate class

2500

B

4-4

0)

Fig. 14.21 Fracture tougness of Sic whisker-Si,N4 composites processed at various pressing temperatures&.

A

2 2000 E

# ,

,x

F

5 1000 1000-

'7 PI 0

5

10

IS

20

25

30

35

40

Volume Yo Sic whiskers

Fig. 14.20 Fracture strength of Sic whisker-Si,N4 composites processed at various pressing temperatures".

'9 ,

,4 /**4

500-

I

',

, ,

5 1500 1500-

A

,

, 0

/

/

/

O f '

.1

'

.

'

'

'

'

.2 .3 .4 .5 .6 .7 Fibre volume fraction

' .8

'

.9

Fig. 14.22 Rule of mixtures strength for Sic fiber (NicalonTM)-borosilicate glass49.

328

Ceramic composites Crater Wear

14.6 APPLICATIONS

14.6.1 HIGH TEMPERATURE STRUCTURES

The most advanced demonstration of ceramic matrix composites for high temperature structures and components has been for Sic CVI infiltrated carbon and Sic (NicalonTM) fiber. Extensive application of these composites have been made by Societe Europeenne de Propulsion A demonstration twostroke 100 cc engine consisting of Sic-Sic piston, cylinder and cylinder head has run for ten hours at full load and a speed of 1500 rpm without lubricant. A list of fabricated and tested high temperature ceramic composite structures is given in Table 14.11. These components operated successfully under actual- or simulated service conditions. 14.6.2 TRIBOLOGICAL COMPONENTS (CUTTING)

The application of ceramic composites to cutting tool inserts has made a significant impact on machining of difficult-to-cut metals. The typical wear pattern in a cutting tool insert is shown in Fig. 14.23.Uniform wear due to the rubbing action of the metal work piece produces the flank and nose wear. Deeper wear patterns at the depth-of-cut and trailing edge region of the insert are due to the sharp edges

Notch

Nose Wear

Fig. 14.23 Typical wear pattern in cutting tool inserts.

of the chip. The abrasion of the rake by the broad face of the chip produces a crater. Material removal from the rake face can be by dissolution, adhesion and chemical reaction. Low chemical reactivity, hot hardness and wear resistance allows ceramics to minimize all these forms of wear. Unreinforced ceramics, while capable of high cutting speeds, suffer from unpredictable life due to low impact load tolerance

Table 14.11 Typical high temperature applications of ceramic matrix composites

Application

Combustion chamber Turbine blade Turbine wheel Turbine wheel Leading edge Nozzle Radiant burner tube

Composition

Operating temperature ("C)

Environment

Test condition

Reference

Sic-Sic

1200

Oxidizing

14 h running

50

Sic-Sic Sic-Sic C-Sic C-Sic C-Sic

1200 1150 1150 1400 1550

Air-kerosene Air-kerosene Air-kerosene Oxidizing Oxidizing

50 50 50 51 52

Nextel 312-Sic

1150

Air-natural gas

Thermal cycle 400-1200°C 1 h up to 55 000 rpm 70 000 rpm Several hours Maximum thermal gradient 60"C/mm 18 month operation

52

Applications 329 or poor thermal shock properties. Reinforcing ceramics especially by whiskers is an effective technique to defeat these limitations. The best known example of this application is the Sic whisker reinforced A1,0, insert. The commercially available composition contains about 30% whiskers and is designated WG300 by Greenleaf Corp., Saegertown, PA, USA. The range of machining parameters for the Sic whisker alumina compared to other conventional and advances cutting tool materials is shown in Fig. 14.2454.Carbide cutting tools are limited to cutting speeds below 100 m/min while the ceramic cutting tools range from over 100m/min to 450m/min. The Sic whisker reinforced composite can be seen to provide the largest range of machining parameters for the machining of nickel based superalloys compared to other advanced cutting materials such as Sialon and Tic particulate reinforced alumina. The Sic whisker reinforced alumina (WG300) also has significantly greater tool life and allows much greater rates of metal removal for Inconel 718 compared to Tic particulate

reinforced alumina and cemented tungsten carbide as seen in Fig. 14.2555.

2

35 -

r 30c

S

c

'E 25> K .-

+ 200

K ._

3

15-

0

$ - 10m c L

2

50-

Fig. 14.25 Comparison of tool life and metal removal rates between various cutting tool materiais55.

-

0.25

Al,O,-SiC(w) Sialon

C a

0.20

FE

i

-

E 0.15

U

(I)

2 0.10

0.05 0

100

200

300

400

500

Cutting speed (m/min)

Fig. 14.24 Approximate range of machining parameters allowed by various cutting tool materials%.

330 Ceramic composites 15. Mecholsky, J.J., Engineering research needs of advanced ceramics and ceramic-matrix com1. Karasek, K.R., Bradley, S.A. Dormer, J.T., Yeh, H. posites, Am Ceram. SOC. Bull., 1989, 68(2), C., Schienle, J.L. and Fang, H.T., 367-375. Characterization of silicon carbide whiskers, I. 16. Maloney, L.D., Make way for 'Engineered Am. Ceram. SOC.,1989, 72, 1907-1913. Ceramics,' Design News, March 13, 1989, 45(5), 2. Homeny, J., Whisker reinforced ceramics, in: 64-71. Ceramic Matrix Composites (ed R. Warren), New 17. Naslain, R., CVI Composites, in: Ceramic Matrix York, Chapman and Hall: 1992, p.245. Composites (ed. R. Warren), New York: 3. Romine, J.C., New high temperature ceramic Chapman and Hall, 1992, p. 199. fiber, Ceram. Eng. Sci. Proc., 1987,8, 755. 18. Buckley, J.D., Carbon-carbon: an overview, Am. 4. Holtz, A.R. and Grether, M.F., High temperature Ceram. SOC.Bull., 1988, 67(2) 364-338. properties of three Nextel ceramic fibers, 19. Besmann, T.M., Sheldon, B.W. and Kaster, M.D., Presented at 32nd International SAMPE Temperature concentration dependence of Sic Symposium and Exhibition at Anaheim deposition on Nicalon fibers, Surf. Coatings Convention Center, April 6-9, 1987. Technol., 1990,43144, 167-175. 5. Bunsell, A.R., Development of fine ceramic 20. Stinton, D.P., Besmann, T.M. and Louden, R.A., fibers for high temperature composite, Materials Advanced ceramics by chemical vapor deposiForum, 1988,11, 78. tion techniques, Am. Ceram. SOC. Bull., 1988, 67 6. Ishikawa, T., Recent developments of the Sic (2), 350-355. fiber Nicalon and its composites, including 21. Warren, R. and Lundberg, R., Principles of properties of the Sic fiber Hi-Nicalon for ultrapreparation of ceramic composites, in: Ceramic high temperature, Compos. Sci. Tech., 1994, 51, Matrix Composites (ed R. Warren), New York: 135-144. Chapman and Hall, 1992, p35. 7. Okamura, K., Ceramic fibers from polymer pre- 22. van Konijnenburg, J.T., Siskens, C.A.M. and cursors, Composites, 1987,18(2), 107-127. Sinnema, S, Practical designing aspects of engi8. Flock, W.M., Characterization and processing neering ceramics, in: Designing with Structural interactions, in: Ceramic Processing Before Firing, Ceramics, (ed R.W. Davidge and H.M. Van de (ed. George Y. Onoda, Jr. and Larry L. Hench) Voorde), New York: Elsevier Science Publishing New York: John Wiley and Sons, 1978, p.31. Co, Inc., 1991, p.98. 9. Laynge, F.F., Lam, D.C.C., Sudre, O., Flinn, B.D., 23. Pigott, M.R., Load Bearing Fibre Composites, Folsom, C., Velamakanni, B.V., Zok, F.W. and Oxford: Pergamon Press, 1980, p.106. Evans, A.G., Powder processing of ceramic 24. Aveston, J., Cooper, G. and Kelly, A,, Single and matrix composites, Mater. Sci. Eng., 1991, A144, multiple fracture, in: The Properties of Composites, 143-152. Guildford: IPC Science and Technology Press, 10. Kragness, E.D., Processing and mechanical 1971, p.15. behavior of tape cast and laminated silicon car- 25. Halpin, J.C. and Kardos, J.L. The Halpin-Tsai bide whisker/alumina composites, M.Sc. equations: a review. Polym. Eng. Sci., 1976,16(5), Thesis, Pennsylvania State University, 1988. 344-352. 11. Prewo, K.M., Brennan, J.J. and Layden, G.K., 26. McCullough, R.L., Wu, C.T., Seferis, J.C. and Fiber reinforced glasses and glass ceramics for Lindenmeyer, P.H. Predictions of limiting high performance applications, Am. Ceram. SOC. mechanical performance for anisotropic crysBull. 1986,65(2), 305-322. talline polymers. Polym. Eng. Sci. 1976, 16(5), 12. Hillig, W.B., Melt infiltration approach to 371-387. ceramic matrix composites, 1. Am. Ceram. SOC., 27. Schapery, R.A. Thermal expansion coefficients 1988, 71(2), C9W99. of composite materials based on energy princi13. Newkirk, M.S., Lesher, H.O., White, D.R., ples. J. Compos. Mater. 1968, 2(3), 280404. Kennedy, C.R., Urquart, A.W. and Claar, T.D., 28. Kragness, E.D., Amateau, M.F. and Messing, Formation of LanxideTMceramic composite G.L., Processing and characterization of lamimaterials, Ceram. Eng. Sci. Proc., 1987, 8,879. nated Sic whisker reinforced A1203, J. Compos. 14. Newkirk, M. S., Urquart, A.W., Zwicker, H.R. Mater., 1991, 25(4), 416432. and Brevel, E., Formation of LANXIDETM 29. Paul, B., Predictions of elastic constants of mulceramic composite material, J. Mater. Res., 1986, tiphase materials, Trans. Met. SOC.,AIME, pp. 1(1),81-88.

REFERENCES

References 331 3 6 4 1 (February 1960). 30. Pister, K.S. and Dong, S.B., Elastic bending of layered plates, 1. Eng. Mech. Din, ASCE, 1-10 (October 1959). 31. Reissner, E. and Stavsky, Y., Bending and stretching of certain types of heterogeneous aelotropic plates, J. Appl. Mech., 402408 (September 1961). 32. Amateau, M.F., Properties of laminated ceramic composites, The 37th Sagamore Army Materials Research Conf Proc., (ed. D.J. Viechnicki), Watertown, Mass: Materials Technology Laboratory, 317-338, October, 1990. 33. Evans, A.G. and Davidge, R.W., A biaxial stress method for the determination of the strengh of sections cut from glass containers and the size of critical Griffith flaws, Glass Tech. 1971, 12(6), 148-154. 34. Evans, A.G. and McMeeking, R.M., On the toughening of ceramics by strong reinforcements, Acta Met., 1986, 34, 2435-2441. 35. Becher. P, Hsueh, C.-H., Angelini, P. and Tiegs, T.N., Toughening behavior in whisker-reinforced ceramic matrix composites, J. A m . Ceram. SOC.,1988,71,1050-1061. 36. Warren, R., Fundamental aspects of the properties of ceramic-matrix composites, in: Ceramic Matrix Composites (ed by R. Warren), New York: Chapman and Hall, 1992, p. 82. 37. Faber, K.T and Evans, A.G., Crack deflection process, Part I and II., Acta Met., 1983, 31, 565-584. 38. Hutchinson, J.W., Crack tip shielding by microcracking in brittle solids, Acta Met., 1987, 35, 1605-1619. 39. Vaughn, W.L., Homeny, J. and Ferber, M.K., Mechanical properties of silicon carbide whisker/alumina oxide matrix composites, Ceram. Eng. Sci. Proc., 1987, 8(7-8), 848-859. 40. Tiegs, T.N. and Becher, P.F., Sintering of Al,O,-SiC whisker composites, Bull. A m . Ceram. SOC.,1987,66,347-352. 41. Porter, J.R., Langej, F.F. and Chokshi, A.H., Processing and creep performance of Sic whisker reinforced A1,0,, Bull. Am. Ceram. SOC., 1987,66,343-346. 42. Lio, S., Watanabe, M., Matsubara, M. and Matsuo, Y., Mechanical properties of alumina/silicon carbide whisker composites, J. Am. Ceram. SOC.,1989,72,1880-1884. 43. Becher, P.F., Tiegs, T.N., Ogle, J.C. and Warwick, W. H., Toughening of ceramics by whisker reinforcement, in: Fracture Mechanics of Ceramics,

VoI. 7, Composites, Impact, Statistics and HighTemperature Phenomenon, (ed. R.C. Bradt, D.P.H. Hasselman, A.G. Evans and F.F. Lange), New York Plenum Press, 1986, pp. 61-72. 44. Wu, M., Messing, G.L. and Amateau, M.F., Laminate processing and properties of oriented Sic whisker-reinforced composites, Ceramic Transactions, Vol. 19, 1991, Westerville, OH: American Ceramic Society, 665476,1991. 45. Bulgan, S.T., Baldoni, J.G. and Huckabee, M.L., Si,N4-SiC whisker composites, Bull. Am. Ceram. SOC.,1987,66,347-352. 46. Shalek, P.D., Petrovic, J. J., Hurley, G.F. and Gac, ED., Hot-Pressed Sic Whisker/Si,N, Matrix Composites, A m . Ceram. SOC., 1986, 65(2), 351-356. 47. Milewski, J.V., Gac, ED., Petrovic, J.J. and Skaggs, S.R., Growth of beta-silicon carbide whiskers by the VLS process, J. Mater. Sci., 1985, 20,1160-1166. 48. Petrovic, J.J., Milewski, J.V., Rhor, D.L. and Gac, ED., Tensile mechanical properties of Sic whiskers, 1.Mater. Sci., 1985,20, 1167-1177. 49. Dawson, D.M., Preston, R.F. and Purser, A., Fabrication and materials evaluation of high performance aligned ceramic fiber-reinforced, glass-matrix composite, Ceram. Eng. SOC.Proc., 1987,8,815-821. 50. Heraud, L.P. Spriet, 'High toughness C-SIC and Sic-Sic composites in heat engines', in Whiskerand Fiber-Toughened Ceramics, Proceedings of an International Conference, (ed. R.A. Bradley, D.E. Clark, D.C. Larsen and J.O. Stiegler), International Metals Park, OH: ASM, 1988. 51. Melchior, A.B., Pouliquen, M.F., Soler, E., Thermostructural composite materials for liquid propellant rocket engines, Paper AIAA-87-2119, AIAA/SAE/SME/ASEE, 23rd Joint Propulsion Conference, June 29-July 2, 1987, San Diego, CA, American Institute of Aeronautics and Astronautics, Washington, DC. 52. Suacereau, D. and Beaurain, A., Demonstration of carbon-silicon carbide Novoltex reinforced composite nozzle on a LH2-LOx, Engine, Paper AIAA-90-2180, AIAA/SAE/SME/ASEE, 26rd Joint Propulsion Conference, July 16-18, 1990, Orlando FL, American Institute of Aeronautics and Astronautics, Washington, DC. 53. Richards, R.E., Bodkins, D. W. and Copes, S. J., Progress toward a cost effective thin wall RBT, in Energy Technology: Processings of The Energy Technology Conference, Vol. 15, Energy Technology Conference, February 17-19, 1988,

332

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Washington, DC, Government Institute Inc. Washington DC, 749-756,1988. 54. Billman, E.R., Mehrota, P.K., Shuster, A.F. and Beegley, C.W., Machining with A1203-Sic whisker cutting tools, Bull. Am. Ceram. Soc, 1988,67,1016-1019.

55. Rhodes, J.F., Whisker reinforced ceramic composites, in Proc. Fifth Ann. Conf. Maferials Technology, Materials Technology Center, Southern Illinois University, Carbondale, IL, 205-219,1988.

CARBON-CARBON COMPOSITES

15

John D. Buckley

15.1 INTRODUCTION

n

Carbon-carbon (CC) materials are a generic class of composites similar to the graphite/epoxy family of polymer matrix composites. These materials can be made in a 1-D 2-D wide variety of forms, from one-dimensional to n-dimensional, using unidirectional tows, tapes, or woven cloth (Fig. 15.1). Because of their multiformity, their mechanical properties can be readily tailored (Table 15.1). Carbon materials have high strength and stiffness potential as well as high thermal and chemical stability in inert environments. These materi3-D n-D als must, however, be protected with coatings and/or surface when used in an OXi- Fig. 15.1 Multiformity and general properties of dizing environment. carbon-fiber and carbon-matrix composites.

Table 15.1 General properties of carbon-carbon composites Ultimate tensile strength Modulus of elasticity Melting point Thermal conductivity Linear thermal expansion Density

>276 MPa

>40 000 psi

>69 GPa

>lo7 psi

>41OO0C ~11.W 5 m-' K-'

7412°F 6.64 h ft "F

~ 1 .x110"OC

6.1 x W7"F

lOOO°C (1832°F). 15.7 MECHANICAL PROPERTIES Development of advanced carbon-carbon The extreme thermomechanical requirements (ACC) composites has produced a material of the Space Shuttle have been the impetus for that is twice as strong as the CC composite evaluating properties of low-density CC. The first put on the Space Shuttle. The ACC mateuse of CC on the nose cap and leading edges of rial is made using woven carbon cloth. When the Space Shuttle makes it imperative to know unidirectional carbon fiber tapes are interplied as much as possible about all the characteris- with woven cloth to create a hybrid ACC, tics of this material. The effect of temperature strength in at least one direction can be on the ratio of tensile strength to density for increased by >345 MPa (>50 000 psi). Current several classes of high-temperature materials data on thermomechanical and thermochemiis shown in Fig. 15.17. The major advantage of cal properties of some of the advanced CC CC materials for high-temperature applica- systems show that material composition, oxitions is that they do not lose strength as the dation resistance, processing, joining and fiber use temperature is increased. This property is architecture are producing noticeable in contrast to other materials such as superal- improvements in CC materials and structures loys and ceramics. Figure 15.17 shows three (Curry, Scott and Webster, 1979; Buch, 1984; levels of CC strength efficiency. The first, Rummler and Sawyer, 1984; Ransone and and oven cured at 315°C (600"F), liberating all of the hydrocarbons. This procedure leaves silica (SiO,) in all of the microcracks and fissures greatly enhancing the oxidation protection of the CC structure.

Thermal properties 345 Temperature, OC 0

-

,550 I

1100 I

1650 I

High-strengthcarbon-carbon

800

1600 2400 Temperature, OF

3200

**%O

3447 MPa (>500ksi) PAN based graphite fiber High strength >3630 GPa (>SO0ksi) pitch based graphite fiber Composite laminate properties

Teflon fabric where there was insufficient coating material to provide an effective release, resulting in a bond between the Teflon fabric and the composite structure and when using a perforated release film, the perforations were torn rather than pin pricked, allowing excessive resin bleed. Specifications would help limit the inconsistency of some of these products. If one

encounters a problem utilizing support materials, contact the supplier who can provide technical assistance on its products. During the lay-up of composite structures, care must be taken to insure that all areas are covered with a release or separator film. The bleeder and/or breather will bond nicely to composite laminates if there is no separator.

Tooling 361 16.4 TOOLING

for steel, include superior durability, ample tolerance for elevated temperature service and Tooling includes materials, equipment, or good thermal conductivity. forms onto which (or into which) the product Ceramics have favorable characteristics for is made, assembled, or cast. Tooling issues are molds. They have the lowest coefficient of a result of a number of interacting requirethermal expansion and their thermal conducments that are considered when selecting the tivity approaches that of some hardened tool most cost effective tooling. These requiresteels. However, ceramics are brittle at ambiments are as follows: ent temperatures and they must be protected ability to achieve a uniform heat up rate, from processing and handling hazards. One taking into account the mass of the tool; way to get protection is to enclose ceramic allowing sufficient movement of the lami- inserts in a steel case. nate while achieving pressure and Dual steel molds are candidate tools for compaction in all regions; reproducing high quality composites. accommodating resin flow; However, these tools are costly and producfacilitating or allowing for removal of cured tion quantities are often not sufficient to part; amortize tooling costs at competitive prices for realistic tolerance for tool and laminate; the production items so less costly alternatives finishing requirements - coarse finish can are desirable. create lock-on problems; Aluminum molds are less costly. Although adequate area for applying sealant tape for thermal conduction is better for aluminum vacuum bagging and test coupons as than for steel, the tools are less durable and the required; thermal expansion is large. Shallow or flat provisions for vacuum fittings and han- mold plates are usually limited to cures below dling of tool. 177°C (350°F). Other metal tools include sprayed or electroformed molds reinforced Tooling is less expensive for vacuum bag and with cast backings. Alternate types of tooling autoclave/oven molding than for matched die can include composite molds usually based on molding methods. Molds and molding plates high temperature resistant cast or laminated are required to withstand curing conditions epoxy resins. without distorting or degrading and to tolerMaster forms for laying-up composite tools ate handling during the fabrication processes. can also be fabricated using any of the materiThey are not necessarily resistant to unbalals described; a mock-up model of the item anced pressures. The higher costs of composite may be used, or plaster masters can be pretooling can be amortized by taking advantage pared from models. The quality of the plaster of the improved capabilities to mold complex masters depends on the strain compatibility constructions. Composites that may ordinarily between the plaster and its reinforcements and require secondary bonding are often more ecoon the condition of the hardened surface. nomically co-cured. Composite tools can be laid-up using fiber orientations that most closely match the expansion of the items to be produced. 16.4.1 MOLD PREPARATIONS Fiberglass and graphite fibers are the principal Coefficients of thermal expansion for conven- reinforcements. Woven fiber reinforcements tional tooling materials and composites are are the most economical to use. Mold maintegiven in Chapter 25. For metals, the coefficient nance is best relegated to specialized for steel compares most closely with the coeffi- personnel while preparations for the bag cients for the composites. Other characteristics molding processes are assigned to production

362 Hand lay-up and bag riiolding personnel. A successful practice is to provide production personnel with soft tools and solvents that do not degrade the molding surfaces. If the production tools and solvents are inadequate for removing debris and cleaning, the molds are taken out of service for maintenance, repair, or replacement. After they are returned to service, they are solventwiped clean and mold release agents are applied. 16.4.2 RELEASE AGENTS

Release agents for bag molding composites include carnauba paste wax, aerosol dispensed compositions that contain carnauba, fluorocarbon resins, or silicone resins, plastic films and metal foils. On most occasions, the wax or resin mold releases do not contaminate the composite surfaces excessively nor prevent subsequent secondary bonding or coating operations. Prior to bonding, the composite surfaces are cleaned with solvent and lightly sanded to remove resin gloss. The user should be aware of national or regional limitations on solvent usage due to toxicity or ODS (ozone depleting substances) concerns. Peel plies can be used to protect clean surfaces for primary adhesive bonds. Plastic films, metal foils and sprayed metal coatings also serve as release agents when they are integrally laminated to the co-cured composites. Both the polished wax surfaces and the sprayed wax coatings are excellent mold releases for composites cured below 121°C (250°F). However, the wax degrades and discolors the composites at higher molding temperatures. Commercial fluorocarbon mold releases are used for higher cure temperatures. FEP (fluoroethylene propylene) mold releases form a continuous film on mold surfaces. Although the condition of the release film is easy to maintain below 177°C (350"F), the coating degrades at higher temperatures above this. Fluorine which is noxious, corrosive and highly toxic can be released from the polymer above 177°C (350°F).

PTFE (polytetrafluoroethylene) is stable and is often contained in mold releases for service in excess of 260°C (500°F).The mold releases contain suspensions of micropulverized PTFE in a volatile dispersant. Depositions on mold surfaces do not form continuous films, but the PTFE particles provide excellent dry lubrication for the release of the cured composite. Furthermore, the residual particles on the cured composite surfaces are easily removed with a solvent-wipe. Since a variety of commercial mold releases that contain fluorocarbon (or equivalent)are on the market, it is essential that manufacturers' recommendations on uses and limitations be scrupulously followed. Silicone oils and greases are to be avoided since they are the most persistent contaminants of molded composite surfaces. They release secondary bonds and coatings from composite surfaces as effectively as they release the cured composites from the molds. Silicone oils and greases migrate and defeat most removal attempts. They contaminate the solvent wetted cloths and sand papers so that instead of removing the silicones, they spread them. Contaminated surfaces may be salvaged for painting by sandblasting with virgin grit. Table 16.3 summarizes the precautions necessary for successful hand lay-up and vacuum bagging operations 16.4.3 MOLD DESIGN

hlold design is a function of cost and projected life and/or use. A production mold should be made carefully from the best materials. Such a mold will be designed by an experienced designer who will incorporate the necessary thicknesses, materials, structural reinforcement and hardware required for the intended use. 16.4.4 PATTERNS (PLUGS)

A pattern (plug) is a temporary form in the exact shape, contour and finish of that to be molded. (If the outer shape of the items desired the inside contour is used.) Patterns

Tooling 363 Table 16.3 Hand lay-up and vacuum bagging precautions Mold release application Selection of correct mold release ‘Seasoning’ new tool to insure coverage a minimum of three coast Compatibility of resin system and mold release Repeated applications can cause excessive buildup Sealant tape Ease of use, release from backing Double or single strip application Removal after cure Cheapest when suitable Flash breaker tape Check compatibility Useful for holding prepreg during lay-up and heat debulks Separator film Compatability and will release as intended Quality control on perforations Drapability for complex shapes Elongation (%) (high elongation) Bleeder Not all bleeders the same, select for application Ensure no compaction during cure Potential seal off during cure Excessive bleed can saturate bleeder

are made of many materials: wood, plaster, plaster/metal and other combinations. Almost any material can be considered as a pattern material if it holds its shape. It is assumed that when only one large composite structure is required, such as a 23 m (75ft) America’s cup yacht hull, the cost of making a pattern and a mold in order to make the hull may not be justified. However, the construction of a pattern that becomes a male plug can be cost effective for high performance composite structures. To avoid excessive cost with this tooling approach, one must remember that it is the total cost of the end product not just the cost of the pattern (plug) that must be considered. A limited use plug for a large marine hull 14-23 m (45-75 ft) in length would be made as follows: a simple wood frame to

Breather Able to malntain vacuum path Care not to puncture nylon film Bagging film Higher the percent elongation the more forgiving Pliability Defect free Select for temperature performance Thermocouples Through bag/sealant tape, potential vacuum leaks Placed on outside of bag - reliability Vacuum fittings Caution not to allow resin to fill 0 Source for potential leaks Removal from bagging film before disposal 0 Integral with tooling Vacuum lines Ensure fittings do not leak Hose has not been crushed 0 Not pinched off Not filled with resin

the inside contour of the hull is constructed from stanchions and stringers and covered with strips of wood, with a laminate of 2 45”, 90” layers. The tooling plug is finished and covered with mold release. 16.4.5 INTEGRATING INSPECTION AND MACHINING

A machine tool is for machining. If it isn’t making chips, it’s wasting time, so keep non-cutting time to an absolute minimum. That’s the standard philosophy most shops try to live by. In fact, many shops are investing in pallet shuttles, quick-change fixtures, tooling systems, rapid transverse fixtures, programming and scheduling systems to keep spindles turning and cutting tools w o r h g at optimum capacity.

364 Hand lay-up and bag molding

Inspecting the workpieces right on the fiveaxis mills, has the equipment functioning as both machine tool and coordinate measuring machine, so that the production of parts and producing inspection data become equally important. Such a radically different plan means that design, numerical control (NC) programming, machining and inspection cannot be separate functions.Just as each machine tool would have to serve more than one role, one computerized database would have to share the same information with designers, programmers, operators and inspectors. This combination allows a shop to machine, inspect and analyze any surface without removing the work piece from the five-axis machine. This system helps produce higher quality tooling with significant gains in productivity. A large machine bed will accommodate unusually long workpieces and also leave room for smaller workpieces to be clamped on one end while another workpiece is being machined at the other end (see Fig. 16.2).

parts. If the workpiece will not fit into the hard gauge, it has out-of-tolerance features and will not fit mating parts. The workpiece is rejected. A soft gauge can be used to make similar either-it-fits-or-it-doesn’t comparisons. Instead of placing two physical objects together, two CAD models are laid one over the other on the graphics screen. The software version of the checking fixture is the soft gauge. The software version of the workpiece to be inspected is a geometric model constructed from inspection data. Out-of-tolerance conditions will be just as conspicuous in this comparison, but analysis is far more complete and much faster. Moreover, a soft gauge is created directly from the original design data. Because it is created on a computer screen instead of in a tool room, a soft gauge can be constructed quickly and modified easily. It spares the high cost of building and validating a hard gauge.

Closed-loop machining

Closed-loop machining begins with electronic data representing part geometry from the cusSoft gauge tomer. This data describes the outer surface of A soft gauge can be compared and contrasted the customer’s end product. The CAD system with a ’hard’ gauge such as a conventional then creates a 3D model of its surface. Once checking fixture used for inspection and qual- this surface has been established, all manufacity control. If a workpiece drops into the hard turing operations will be derived from and gauge, it is acceptable and will fit with mating related to it. NC tool paths will be generated

Fig. 16.2 Closed-loop machining - mounting various tools. (Courtesy of Coast Composites, Inc.)

Fig. 16.3 Closed-loop machining - touch probe inspection tool. (Courtesy of Coast Composites, Inc.)

Tooling 365 from it. Using dynamic display of the tool path, programmers can visually verify the NC program, check clearances and make sure gouges are avoided. The inspection path will be generated from the same surface geometry. By referencing the soft gauge, the inspection path will be sure to include checks of all critical features. The path of the probe can be visually verified in the same way as the NC program. After executing the NC program, the workpiece can be inspected immediately using the touch probe in the spindle (see Fig. 16.3). This inspection can be considered in-process, because the workpiece is still fixed on the machine tool and can be remachined without being moved or refixtured. This approach is called closed-loop machining. Results of this inspection routine are automatically used to create a 3D model of the features checked. By comparing this model to the soft gauge, any out-of-tolerance conditions can be identified. It will also show where additional machining passes will be required. Final inspection can performed on the machine tool. These results are compared to the soft gauge again to verify that the contoured surface of the graphite tooling will produce the intended part. Using this machining approach for inspection on the machine tool reduces inspection time by 80%. The biggest savings come from eliminating workpiece moves and additional setups and from streamlined programming of the inspection routines. By integrating inspection and machining, overall manufacturing cycle time can be reduced by 30%.

the composite materials. For complex shapes with integral stiffeners, each block of silicone rubber is wrapped on all but one side, the side in which the tooling rubber is removed (see Figs 16.4(c) and (d)). In tooling a thermal expansion molding, it is best to avoid using rubber on both sides of a laminate as illustrated in Fig. 16.4(e)unless straight edges are not critical. Vacuum bag assist (see Fig. 16.4(f)) provides an alternate method. The linear thermal coefficient of most silicone rubbers that have been measured fall into the range of 1-2.1 x This range is consistent over a 23-246°C (75480°F) temperature range. The rubbers are said to have a linear expansion of approximately 17 times that of carbon steel which is why they are used to mold composites by thermal expansion molding techniques. Precautions in mixing some silicone rubber compounds are required if full potential is to be achieved. During prolonged storage, the catalyst tends to separate and settle to the bottom of the container. Mixing the catalyst prior to adding it to the base rubber will allow correct mixtures and long tool life.

(a) Compression molding

16.4.6 THERMAL EXPANSION MOLDING

Thermal expansion molding techniques are utilized for special applications of small complex composite structures and composite tubing with critical outside surfaces. Figures 16.4(a)and (b) illustrate the methods allowing the expansion of the silicone rubber to provide the required pressure for the compacting of

(b) Oven cure critical outer surfaces

Fig. 16.4 Thermal expansion molding. (Continued on next page.)

366

Hand lay-up and bag molding

(c) Enclosed molding

Incorrect mixtures will be light or dark in color and materials such as Silastic J@ will start to crumble within a few thermal cycles. The use of thermal expansion rubber can be hazardous if not planned well. The tooling rubber can exert up to 6.9 MPa (1OOOpsi) during its first few thermal cycles. The tooling rubber requires the minimum of five full cure heat cycles free standing after the initial cure in order to stabilize the expansion characteristics. The tooling is capable of producing 50-75 composite parts before having to be replaced. One other problem with thermal expansion rubber is with its removal from the composite structure. Sharp pointed objects will have a very lasting effect on tool life; once the rubber is damaged, it will continue to tear, needing replacement much sooner than usually required. Silicone rubber is very slow to cool down and extra time must be allowed because the rubber is impossible to remove from the composite part until it has shrunk back to its original size. Putting the thermal expansion rubber tooling into a freezer can accelerate the manufacturing cycle. However, since some of the tools are heavy, due to steel outer encasements, a 12-15 h cooling down period should be planned into the manufacturing cycle. Experimentation is suggested with this molding method since extreme pressures can be generated and undesirable results may occur if the molding method is picked arbitrarily.

3

~~~~

(d)Negative draft molding

r - - - - 1

I

(e) Oven/press cure

1

16.5 BAG MOLDING PROCESS

(f) Vacuum bag assist

Silicone rubber Steel molds

IICompositelaminate Fig. 16.4 Continued.

Molding methods include vacuum bag, pressure bag, oven and autoclave molding. Bags, the thin and flexible membranes or silicone rubber shapes, separate the laid-up constructions from atmospheric pressure during composite cures. The bagged lay-ups in autoclaves are usually vented to pressures lower than those applied to the bag. Consolidations and densifications of the lay-ups are achieved by the resulting pressure differentials across the bag contents. Consolidations are achieved

Bag molding process 367 when the separate plies of prepreg in the layups and other adherents are bonded together. Densifications result in reduction of voids and removal of excess resin. Other results desired of bag molding methods during cure include prevention of blistering in the composites, increased controls of pressure and heat application and control of the fiber/resin ratio. Consolidations and densifications of vacuum bag moldings can be achieved by atmospheric pressure alone as the bagged layups are evacuated throughout the cure cycles. The pressure-bagged and autoclaved-cured composites are pressurized by hot gases. Vents to the atmosphere or vacuum provide for the escape of the volatilized reaction by-products and the entrapped air from the curing composites. If the pressures within the bag are not reduced from those applied to the bag, the membrane remains inert and there is no compaction. Of the three methods, vacuum bag molding is least limited as to the size of constructions that can be processed. On a few occasions, 'wet' lay-up vacuum bag molded composites are room-temperature cured. Most are thermally cured to produce improved properties. Thermal cures are best attained in air circulating ovens/autoclaves, but can also be achieved in infrared heated and passive type convection ovens. Pressure bag molding methods are efficient for producing both deeply contoured structures and shallow composites. Sonar domes, radomes and antenna housings are examples of deeply contoured composites.Architectural panels, door panels and aircraft fairings are examples of shallow composites. Heavy molds are built to reproduce deeply contoured structures. Each specialized mold is constructed to withstand the elevated temperatures and increased pressures required for the cures. Shallow items may often be bag molded in modified compression presses. The lower press platens contain vents and vacuum lines. The upper press platens are made hollow to enclose the mold plates together with the laid

up assemblies. When the presses are closed, the sealed chambers are pressurized and heated to attain molding conditions similar to those of an autoclave. Unlike the specialized pressure bag molds, the modified presses are used to cure many different composite constructions. Autoclave and pressure bag molding conditions to 177°C (350°F) and 1379 kPa (200 psi) are routinely attained. Newer, customized autoclaves attain cure conditions that exceed 260°C (500°F) and 3447 kPa (500 psi). Fire hazards are greatly increased at elevated temperatures and pressures. Pressure vessel fires are minimized by uses of fire retardant processing materials and inert pressurizing gas. Fire prevention measures include uses of silicone rubber, nylon or Tedlar bags. Before cure cycles are initiated, the pressure vessels are purged of all enclosed air. After the thermal cure is completed, the pressure vessels and their contents are cooled to 68°C (150'F) before the pressure is relieved and the autoclave is opened. 16.5.1 EXPENDABLE VACUUM BAGGING TECHNIQUES

Bleed-out systems are devised to maintain reduced pressures within the bags' contents. The bagged lay-up includes the bleed out system designed for the composite part. Bagged lay-ups can be bled in two ways: vertically or edge bled. The classical differences between the two can be seen by comparing Figs 16.5 and 16.6. Many of today's resin systems are mostly 'net resin' and do not require any resin bleed during cure. This allows for better control of the resin content of the composite structures. If a resin bleed sequence is preferred, the following sequence can be used as a guide. 0 0

The surface of the mold is prepared with the release agent. The composite plies are applied and rubbed out to remove entrapped air.


I ATMOSPHERIC OR VACUUM VENT

Fig. 16.5 Vacuum bag edge bleeder - schematic. CAUL PLATF

RAG

NOTE HEAVY PROTECTION AGAINST BAG PERFORATIONS DUE TO INCREASED AUTOCLAVE PRESSURES

Fig. 16.6 Vacuum bag vertical bleeder - schematic.

Bag molding process 369 A perforated release film is applied over the composite laminate and extended approximately 3.2 mm (1.25 in) beyond all edges. A predetermined number of bleeder plies are applied over the release film and extended to the perimeter of the lay-up. A perforated release film is applied over the bleeders and extended 3.2 cm (1.25 in) from edge. One or two layers of a non-woven breather is placed over the lay-up and extended over the release film. Sealant tape is applied around the perimeter of the bleeder. The vacuum bag is positioned and sealed. The contents are evacuated and smoothed and the bag is checked and sealed against leaks. The bagged lay-up is ready to be cured.

In any bagging sequence, the types of release film, bleeder, breather and bagging materials used vary from company to company and from supplier to supplier. Each supplier has typical data sheets on expendable materials to acheve the most efficient use of the materials.

f

In a typical vacuum bag lay-up, there are several methods available; some use double sealant tape side/side, some single and some one on top of the other. The best system is the one that works. The side/side method is used to provide some insurance during cure that the bag will not shrink, pulling an edge off, causing loss of vacuum. The over/under method is used to provide ease of placement of ears to allow some movement of the vacuum bag. No matter which method is chosen, it is important to remember that vacuum bags tend to pull down more than expected and can puncture, if bagging is over a sharp object. During the application of a vacuum bag, 'ears' are required to facilitate the uniform application of vacuum to the composite laminate. Vacuum bag bridging is one of the leading causes of resin rich and excessive voids in corners of composite laminates. Figure 16.7 illustrates this common problem. One method of eliminating bridging of the vacuum bag is presented in Fig. 16.8by means of 'ears' in the bag. Another method to help reduce resin rich

SEALANT TAPE

VACUUM BAG

LAMINATE BEING FORMED

Fig. 16.7 Vacuum bag bridging (Morena, J., Advanced Composite Moldmaking; New York, Van Nostrand

Reinhold, 1988).


fSEAMNT TAPE

VACUUM BAG

4- INCH HIGH PLEAT OR FOLD

LAMINATE BEING FORMED

Fig. 16.8 Elimination of vacuum bag bridging (Morena,J., Advanced Composite Moldmaking; New York, Van Nostrand Reinhold, 1988).

Fig. 16.9 Large vacuum bagged structures. (Courtesy of Richmond Aircraft Products.)

Bag molding process 371 and excessive voids in corners is the place- 16.5.2 REUSABLE VACUUM BAGGING ment of an intensifier over the area, usually TECHNIQUES placed between the separator film and There are material and recurring labor cost breather. The intensifier can be molded rubber disadvantages to the use of expendable vacin the radius desired or some sealant tape to uum bags of plastic films for fabricating fill the corner. 'Ears' may be required in sev- production composites. Expendable bags, laid eral sections of a complex part. Experience will up of plastic films and associated sealants, also determine the height of the ear for a specific incur recurring costs. Expendable bags can be application; 10 cm (4 in) is a good starting laid up only once because of degradation durpoint. Some will be smaller and some will be ing handling and the thermal cures. larger depending upon the complexity of the Use of silicone rubber reusable bags can component being vacuum bagged. reduce fabrication costs and defective comThere is essentially no limitation on the size posite parts because of resulting work of thermoset composite structures. The use of simplification and more positive control of the the thermoset vacuum-bagged composites bag molding cure conditions. Figure 16.11 (autoclave, oven or integrally heated) will con- illustrates an example of a component being tinue to provide excellent composite manufactured utilizing a reusable silicone structures for many years to come. Figures 16.9 and 16.10 illustrate some more complex uses for expendable vacuum bags.

b

I

I

Fig. 16.10 Complex vacuum bagging. (Courtesy of Richmond Aircraft Products.)

Fig. 16.11 Disposable vacuum bagging (top); reusable vacuum bagging (bottom). (Courtesy of The Darner Corporation.)

372 Hand lay-up and bag molding vacuum bag and an expendable vacuum bag. There are often difficulties in having the facilities to handle large reusable vacuum bags weighing several hundred pounds. A cost evaluation must include all aspects of the program. There are some very large aircraft components utilizing expendable vacuum bag materials very successfully. The more complex the composite structure becomes, the more effective the reusable vacuum bagging system is.

are combined when the product is made. The composite designer must consider how the load bearing fibers are placed and ensure that they stay in the proper position during the fabrication. 16.6.1 DESIGN PROCESS

With some large composite structures, potential problem areas can best be identified using scale models. Working problem areas on an individual basis, a major factor in the success 16.6 DESIGN of a program is the amount of planning that The fundamental information needed for any can take advantage of the work force experidesign includes the stresses applied under ence. Successful composite structures are not storage and use and the strength of the mater- fabricated by one person; they require team ial used. Assume that the size, shape, quantity work from all disciplines. It is considerably and rate of production have dictated the use of harder to make a smaller composite structure, open mold techniques. Then the final thick- than make the full scale article. Procedures ness, orientation and quantity of reinforcing developed for the scaled article can, however, fibers are dependent upon the stresses that be easily translated into a full scale structure. must be resisted, how often and for how long. It is essential that the designer find out 16.6.2 FIRST ARTICLE FABRICATION what strength can be built into the laminate. This sets composite structures apart from The first article, also known as tool proof artiother types of materials, since the material is cle, can be used to provide information not made during fabrication of the product; the only on the tool to manufacture to the correct percentages and orientation of the reinforce- tolerances but also to produce an acceptable ment and the types of resins determine the composite structure meeting the design requirements. In addition, it can be utilized as properties of the final laminate. During the initial phase of the development a proof of the documentation of quality conof a composite structure, there is a need for trol inspection requirements, manufacturing design, manufacturing engineering, tooling, procedure's verification and allow design materials and quality control to provide inputs engineering to review overall requirements. so that the selection processes can be established. By coordinating early in a program, 16.6.3 DESIGN DETAILS one can focus on the real problems of design and manufacturing. The preplanning phase Parts with severe contour and thickness will allow for a program to develop at a more variations rapid pace. Confirming materials, manufacturing methods, tooling concepts and design It may seem easy to incorporate variations in requirements early can avoid the extra time contour and thickness into the design of a new and expense to attempt to make the composite product but in open mold products such variations must be made with caution. The fabrication process successful. The design of composite structures, while molding operation requires laying the materessentially similar to conventional design, ial on the mold to follow mold contour. If the does have the added dimension that materials angles are sharp (90' without radius), the lay-up

Design 373 will not follow the mold surface and will develop voids and resin rich areas in the laminate in the vicinity of the angle. For instance, in inside right angle corners without radius, (Fig. 16.12(a))the laminate will not pack into the corner. When there are sharp outside corners, the laminate (Fig. 16.12(b))will not wrap tightly over the corner. The solution to such problems is to design with a generous radius, preferably 4.75-12.75 mm (0.187-0.500 in) inside and out. The laminate will then follow the contour. Abrupt changes in direction are high stress areas and tend to delaminate and crack. They should be avoided and moderate self reinforcing curvatures used.

Changes in thickness To change thickness in open mold construction is to add or remove plies of material. An abrupt change means that the plies must be carefully laid up in a precise pattern. An abrupt change in thickness (Fig. 16.13(a)) results in a stress concentration and should be avoided as delamination is sure to occur at such a point. The solution to this problem is not to have abrupt changes but to gradually change by stepping back or 'shingling' the layup (Fig. 16.13(b)).

Openings The best opening is a round hole; the worst, an opening with sharp, non-rounded corners. The solution to stresses in an opening is to use large radii in the corners, to build up thickness gradually at the sharp corners, or to design a molded in flange around the opening (Fig. 16.13(c)).

1

L Fig. 16.12 Corner lay-up techniques (a) radius corner; (b) no radius corner.

I

Fig. 16.13 Changes in ply lay-up (a) abrupt changes; (b) stepped piles; (c) hole reinforcement.


Joints and bonding Although common practice is to use joints intended for other methods of fastening, structural adhesives require joints of a special design. It cannot be stressed too strongly that the practice of using ordinary joints that have been slightly altered can lead to disastrous results. The type of joint used depends on the basic characteristics of the adhesive since structural bonds act over an entire area and not at a single point as rivet fasteners do. A joint should therefore be designed to minimize concentrations of stress. There are four basic types of stress encountered in structural bonding: tensile, shear, cleavage and peel. 16.6.4 DESIGN CONSIDERATIONS

However, these structures are not generally highly stressed structures. Large amounts of waste can be expensive for any program. As a rule, in programs where is little or no preplanning, the waste factor can be 20-35%. This may be acceptable for low cost materials such as fiberglass but for carbon prepreg materials costing over US$59OO/kg for 827 GPa (120 msi) pitch fiber, the cost for the waste can break an otherwise successful composite program. A well-planned approach to the cutting, kitting, lay-up and inspection requirements can reduce the waste factor to 10-15%. Large structures tend to have less waste than small components. The carbon epoxy central cylinder for a modern communication satellite has less than 3% material waste, the majority being for localized reinforcements. The America's Cup racing yacht was fabricated with less than 2% material waste.

The initial lay-up starts with the preparation of the prepreg ply kits. Individual ply kits will reduce the overall labor requirements during the ensuing lay-up. Hybridization of materials can be achieved on a ply-by-ply basis. Off axis 16.6.5 GENERAL DESIGN PRACTICE - . . - . (* 45") ply orientations can be prepared utilize Attention to ply orientation on strength ing a woven graphite cloth. Interleaf layers of controlled laminates can prevent matrix titanium or fiberglass can distribute high load and stiffness degradation. The 0" ply orieninputs into the laminate. tation is used to carry the longitudinal With large composite structures, i.e. racing loading, the 90" ply orientation the transyachts, the prepreg materials are usually disverse loading and the 45" ply for shear pensed directly onto the lay-up. One must loading. allow for 'fresh' prepreg, with maximum tack, e In order to minimize in-plane shear, place to be applied to the laminate with minimum the +45" and -45" plies together; the inout time of the material during the fabrication plane shear is carried by the tension and phase. compression in the 45" plies. One of the first reactions to using hand layTo minimize warpage and interlaminar up fabrication methods for a composite shear within a laminate, maintain the symstructure is: "this will be too labor intensive". metry about the center line of the laminate. This can be misconstrued to mean that all Stress concentrations can be minimized by composite structures done by the hand lay-up designing tapered or stepped laminate are expensive. Complex, integral, stiffened thickness changes. composite structures may not only be cost The placement of specific ply orientations effective but may not be able to be fabricated can influence the buckling strength and by other methods of manufacture. damage tolerance. The outer ply orientaCompression molding, resin transfer moldtions influence the laminate bending ing and sheet molding processes are cost characteristics more than plies placed at the effective for specific types of material forms. neutral axis.

-

*

Applications 16.7 APPLICATIONS

16.7.1 AEROSPACE

Two aerospace examples that were manufactured utilizing the hand lay-up and vacuum bagging process were an ICBM equipment section structure (300 parts) and a central cylinder (6 parts) for a modern communication satellite. These two primary composite structures illustrate that hand lay-up and vacuum bag procedures can be effectively utilized on limited production programs. For greater production demand or with a less complex structure other methods (machine lay-up or filament winding) could have been used. However, neither alternative process could achieve the results required with materials selected.

ICBM equipment structure The ICBM equipment structure was initially designed to utilize unidirectional graphite/epoxy materials because unidirectional materials were available during the design phase. The development of woven graphite was undertaken in order to reduce the manufacturing costs of the program. Initial prototypes (unidirectional tape) required over 2000 man-hours labor to complete. The first prototype utilizing woven graphite required only 900 man-hours, better than 50% reduction in manufacturing costs. Eventually production labor requirements were reduced to less than 200 man-hours. Satellite central cylinder The graphite/epoxy central cylinder was a unique structure from its inception. The project was undertaken not just because composites would be lighter but that the metallic (beryllium) design required a longer manufacturing time that was unacceptable. The materials of choice were a 520 GPa (75 msi) pitch graphite fiber, epoxy resin and aluminum honeycomb. During the initial

375

manufacturing/design reviews, the decision to fabricate the central cylinder as a one piece monococque structure was made. One quarter scale models were made as initial feasibility trials and the finished cylinder was 108 cm (42 in) in diameter and 231 cm (91 in) high. With that initial success, a full size manufacturing development cylinder required 1700 man-hours to fabricate. By production unit five, the fabrication time was reduced to 500 man-hours per cylinder. The final design incorporated compression molded graphite inserts for hard mounts and was co-cured as a one piece structure. The program set new standards of cooperation between engineering analysis/design, manufacturing and quality control groups. A manufacturing plan including all required inspection points and a detailed fabrication sequence was prepared. The major problem during the fabrication was the out time available with the graphite epoxy resin system. With only 10 working days available, the kit preparation, lay-up, compacting cycles and final cure was on a tight time schedule. This program used various operational procedures to achieve success. Tank inserts were compression molded from chopped carbon/epoxy material prior to the start of the laminate schedule on the cylinder. It was determined early in the development cycle that the unidirectional tape material handled better if it was not precut into kits, but prepared just prior to application to the tool. A combination of heat debulks and pre-bleeding was utilized to maintain the desired resin content, until the supplier could prepare an acceptable net resin unidirectional tape. With limited facilities to autoclave a large cylinder with in-house capability, other methods of manufacture were employed. After the completed hand lay-up was done, a vacuum debulk was applied to ensure all air was removed and the lay-up was compacted. The cylinder was then wrapped with perforated shrink tape (to allow for resin/air bleed during cure), then breather cloth was applied

376 Hand lay-up and bag molding Challenge. A typical America’s Cup yacht utilized over 13 006 m2 (140 000 ft’) of ply surface area of unidirectional graphite tape in the hull and deck structure and 9290 m’ (100 000 ftz)of ply surface area for a one piece graphite mast structure. Due to the overall size and past experience with the boat builder, a male wood plug was fabricated with integral heated wires imbed16.7.2 MARINE APPLICATIONS ded to reduce the heat sink effect during the The fabrication of an America’s Cup racing final co-cure of a complete hull laminate. As yacht (Fig. 16.14) presented another set of the maximum temperature allowed was 90°C requirements. With limited cure temperature, (183”F),wood was a good choice. The hull laminate was then applied, startresin matrix and control density requirements, ing with the inner skin (as on an America’s the challenge was to be able to hand lay-up Cup racing yacht) then film adhesive and honand cure a large composite structure to meet eycomb core was applied. The inner laminate design requirements. Marine applications had reached new heights when the new rules went and honeycomb was vacuum bagged and parinto effect for the 1992 America’s Cup tially cured. The honeycomb core was smoothed and all joints were filled prior to the application of the outer skin laminate. The outer laminate and any local reinforcements were applied to the core and inner laminate (Fig. 16.15). The completed hull laminate was then vacuum bagged and oven cured. The outside of the hull was essentially complete but had a rough surface. The roughness was greatly dependent upon the care of workers during the lay-up of the outer laminate skin. With reasonable precautions, prepreg material can be placed in such a way that there are no overlaps and all gaps have been filled with additional fiber. This effort alone can save hundreds of man-hours during the final finishing. This effort saved over 1200 man-hours compared with a ’wet’ prepreg that did not produce an acceptable outer surface finish. The oven cure achieved maximum mechanical properties required by the design. Laminate

prior to final nylon vacuum bag. Studies were conducted on utilizing a silicone rubber vacuum bag, but overall program costs and difficulties in handling a large (heavy) bag, pushed the utilization of nylon bagging film for the final vacuum bag. The part was then oven cured.

Fig. 16.14 ’Spirit of Australia’.

consisted of filling and grinding the surface in

Applications 377

-

1-

----

-

.k& I

I

*

I

t Fig. 16.15 'Spirit of Australia' laminate laydown (top) 45" ply laminate; (bottom) 0" ply laminate.

two or three stages and culminating with a final coat of good grade epoxy or urethane marine paint. This step is often postponed until the boat is completely assembled and ready for fitting out. Then, the hull was inverted and placed in a fitted saddle; the male plug was removed. Additions of bulkheads, flooring and interior completed the work. The manufacture of the carbon/epoxy mast for the 'Spirit of Australia' presented some unique challenges. The major program challenge was how to obtain a mast that would provide the required performance at affordable

cost. A discussion with design/analysis and manufacturing concluded that it was possible to fabricate a one piece mast within the facilities available and technical expertise within the syndicate. The design of the tooling was aimed at providing a capability to cure the carbon/epoxy resin within the America's Cup rules. The requirements specified the cure temperature 120°C (250"F), cure pressure (3 atmospheres), laminate density /modulus of graphite fiber and overall mast profile. All these objectives were realized within 100 days, from start to completion of the first one piece, 35 m (115 ft) in length, America's Cup composite mast. Performance exceeded all expectations. Additional masts starting with mast number 3 were manufactured with the final weight objective of less than 450 kg (990 lb). Total fabrication time for each mast, from the start of material kit preparation to completion of the cure, required a maximum of 21 days. The selected carbon/epoxy resin system was workable for this limit. The mast, spinnaker pole and rudder stock utilized variations of pressure bag molding with integrally heated fiberglass tooling. These two tooling approaches allowed for some of the largest one-piece structures to date to be fabricated in the commercial marine market. Logistics can become a major role player in the planning required in that not only the support materials are on hand but that power is available; materials handling is taken into account and there is sufficient crew available to complete the lay-up within the material out time limits. As seen in the above examples, the one major factor for all the programs was the material out time. With the development of new resin matrixes, the design of complex structures can be achieved.

MATCHED METAL COMPRESSION MOLDING OF POLYMER COMPOSITES

17

Enarnul Haque and Burr (Bud) L. Leach

17.1 INTRODUCTION

In today’s highly competitive global economy, the need for materials with the right properties to meet the demands of design, environment, durability and economics is growing. Composite materials, with their high strength and stiffness-to-weight ratios, have many advantages and are a desirable engineering material. There is no universal definition of composites. In general, a composite material is a heterogeneous material system consisting of two or more physically distinct materials. In a composite material system, the individual materials exhibit their unique properties and the composite as a whole shows properties that are different from its constituents. In addition to the constituents’ unique properties, the properties of composites are also dependent on the form and structural arrangements of the constituents and the interaction between the constituents. Broadly speaking, composites consist of two components, a binder or matrix and a reinforcement. The matrix functions as the body constituent, serving to bind the reinforcement together and giving the composite its bulk form. The reinforcements are the structural constituents, providing high strength to the internal structure of the composite.

Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

Reinforcement may take the form of fibers, particles, laminate, flakes and fillers. Depending upon the type and orientation of the reinforcement and the manufacturing technology required to produce them, composites with various properties and cost can be fabricated. Polymer composites are composites in which the binder or matrix is a polymeric material and the reinforcement is usually a thin fibrous material. Polymer composites can have either a thermosetting or thermoplastic matrix. In this chapter we will discuss thermosetting matrix based composites. Reinforcing fibers may also be of various kinds, with glass (E-type), carbon, or organic fibers (e.g. aramid) being the most common. Glass fibers are the most widely used type of reinforcement since they offer good strength and moderately high temperature resistance (about 260°C) at a cost effective price. Glass fibers also come in various forms. They can be continuous filaments, cut or chopped strands, roving and yarns, or in the form of cloth, mats or tapes. Thus allows glass fibers to be used in a variety of applications such as lay-up, filament winding, matched die-molding, etc. In this chapter we discuss the matched metal compression molding of thermoset based polymer composites. Matched metal compression molding is a molding process in whch the cure is obtained while the material is restricted between two mold surfaces and the loading and closing of the mold causes the material to conform to the desired configuration. This

Background 379 process enables large scale production of large surface area parts with contour problems and tight tolerances. Matched metal compression molding employs a ’mold’ or match dies. The male mold is matched to the female mold so that when the dies are closed, a controlled space results. A preform charge is placed on the core and the cavity is pressed against it, applying direct pressure on the material. The pressure in this type of molding varies from 1.38 to 6.895 MPa (200 to 1000 psi) and curing temperatures from 125°C to 160°C (260°Fto 320°F).

17.2.1 BULK MOLDING COMPOUND (BMC)

BMC has been defined as ‘a fiber reinforced thermoset molding compound not requiring advancement of cure, drying of volatile, or other processing after mixing to make it ready for use at the molding press’*. BMC can be molded without reaction byproducts under only enough pressure to flow and compact the material. BMC is usually manufactured by combining all the ingredients in an intensive mixing process. Recent advances in BMC technology dictate that both the dry ingredients and wet ingredients be batch mixed separately and then combined 17.2 BACKGROUND together in an intensive mixer. The BMC is usuAdvanced polymer composites are now being ally in a fibrous putty form when it comes out of applied extensively for all types of applications the mixer and resembles ’sauerkraut’. It is usuin the industrial and automotive markets. Table ally compacted and extruded into bars or ’logs’ 17.1’ shows the usage of composites in various of simple cross section. The earliest BMCs were probably made markets during 1991-1993. This section deals primarily with thermoset polymer composites about 1950, employing a process of impregnatused in matched metal compression molding. ing roving strands with blend of resin, filler, etc. The two most popular reinforced molding com- and chopping them to a length in the wet stage. pounds used in the plastics industry are Premix Since wetting glass fibers with a resin containor BMC and SMC (also referred to in modified ing much filler is difficult and slow, these versions as HMC and XMC). Low Pressure premixes had a high glass content. The first Molding Compounds (LPMC),ZMC and TMC high volume commercial BMC was made with sisal fibers and used in molding automobile are also becoming popular.

Table 17.1 US Composites shipments: 1991-1993”

Millions of pounds Markets Aircraft/ aerospace /military Appliance/business equipment Construction Consumer products Corrosion-resistant equipment Electrical/electronic Marine Transportation Other Total

1991 38.7 135.2 420.0 148.7 355.0 231.1 275.0 682.2 73.8 2359.7

1992 32.3 143.2 483.0 162.2 332.3 260.0 304.4 750.0 83.4 2550.8

1991-1 992

1993

1992-1 993

% change

(projected)

% change

-16.5 +5.9 +15.0 +9.1 -6.4 +12.5 +10.7 +9.9 +13.0 +8.1

26.0 146.7 522.0 164.1 336.8 273.0 317.2 810.0 88.0 2683.8

-1.95 +2.4 +8.l +1.2 +1.4 +5.0 +4.2 +8.0 +5.5 +5.2

a Includes reinforced thermoset and thermoplastic resin composites, reinforcements and fillers. Source: SPI Composites Institute

380 Matched metal compression molding of polymer composites heater housings. Improvement in the binder chemistry of glass fibers, development of a chemical thickening system and thermoplastic low profile additives help BMC to attain strength, chemical resistance and to overcome surface irregularities. Consequently, BMC was accepted for use in the electrical, chemical and appliance industries. Today, BMCs are accepted as high performance engineering thermoset molding compounds and used extensively in the electrical, automotive and consumer goods industries. BMC is increasingly injection molded to take advantage of the automation and reproducibility afforded by the process, although it is also both transfer molded and compression molded.

17.2.3 THICK MOLDING COMPOUND (TMC)

TMC was developed by Takeda Chemical Industries, Ltd (Osaka, Japan). TMC is suited to compression, injection and transfer molding and is usually processed on the same equipment as SMC and BMC materials. TMC composites are usually produced up to 51 mm (2in) thick and glass fiber length can vary from 6.4 to 50.8 mm (0.25 to 2 in). In TMC, continuous impregnation and high sheet weight result in complete wet-out of resins, fillers and reinforcing fibers. Better wet-out results in improved mechanical properties and reduced porosity. TMC is usually used in business machine housings, appliance components and other consumer related industries.

17.2.2 Z MOLDING COMPOUND (ZMC)

17.2.4 SHEET MOLDING COMPOUND (SMC)

ZMC was developed in 1979 in France to improve BMC performance. BMC suffers from glass fiber degradation during injection molding and ZMC was developed to keep shear forces as low as possible during moldin$. A special type of injection molding machine developed by Billion in France combines the advantages of both a screw machine and a plunger machine. The ZMC injection machine uses a screw to homogenize and precisely measure the shot. The injection is made like a plunger by the displacement of the screw and inner barrel inside the main barrel. In a ZMC, the different components are mixed in conventional mixers like BMC. The compound viscosity is usually low and adapted to injection machine characteristics. The design of the mold plays a key role in the processing phase and ZMC parts cannot be successfully made unless part design and mold design are combined upfront. Compared to SMC, ZMC parts have lower mechanical properties, but higher performance when compared to conventional injection molded BMC.

SMC is a type of fiber reinforced plastic which primarily consists of a thermosetting resin, glass fiber reinforcement and filler. Additional ingredients such as low-profile additives, cure initiators, thickeners and mold release agents are used to enhance the performance or processing of the material4. The development of SMC started in the early 1950s after the finding that the viscosity of unsaturated polyester resins increases with the addition of Group IIA metallic oxides, hydroxides, or carbonates5. The first published report on SMC was presented at the Cleveland Section of the Society of Plastics Engineers meeting. The report involved work done in Germany using fiberglass mat impregnated with a resin mixture containing magnesium oxide6.At the same time a number of US patent^^,^ were published on the use of Group I1 metal oxides, hydroxides, or carbonates for use on adhesives. The early applications of SMC materials were in electrical and industrial goods. During the next two decades, growth in commercial usage of SMC followed the evolution of continuously improving equipment, low profile additive,

Background catalyst, etc. The automobile industry started using SMC in the early 1970s for producing exterior body components, such as hoods or grille opening panels. With the introduction of high strength SMCs in the mid-l970s, usage of SMC increased to structural components. SMC is currently used extensively in transportation, construction (door panels), appliances (washing machine door, refrigerator housing), furniture (chair, tabletop) and business machines (computer housings). The transportation industry has the highest level of consumption of SMC. For instance, in the North American market alone, the annual rate of consumption exceeds 100 million kg9. Details of SMC manufacturing are available in the literature5.SMC offers many advantages which include variety, part consolidation, lightweight and dimensional stability. With the evolution of flexible backbone polyester resin systems and development of special additives, flexible SMC is becoming very popular and is now competing with thermoplastics for vertical body applications. Special applications SMC is also becoming popular. With the addition of hollow microsphere glass bubbles in a standard SMC formulation, lower density (1.3-1.4) is obtained for weight reduction. High strength molding compound (HMC) is a SMC containing 65% chopped glass fiber instead of the usual 25-35%. HMC uses little or no filler and can be compounded on a standard SMC machine. Directionally reinforced molding compound (XMC) is a directionally oriented moldable resin-glass fiber sheet containing 65-75% continuous reinforcement. XMC is also usually compounded on standard filament winding equipment and has strength five times greater than SMC. Unidirectional molding compound (UMC) is a system of chopped and continuous fibers produced on a modified SMC machine. An advantage of UMC is that different varieties of fibers can be used.

381

17.2.5 LOW PRESSURE MOLDING COMPOUND (LPMC)

Low Pressure Molding Compound (LPMC) is an SMC type material which can be molded at 1.38-2.07 MPa (200 to 300 psi) instead of 5.52 -6.90 MPa (800-1000 psi) required for standard SMC. LPMC is made by replacing the chemical thickening mechanism of alkaline earth oxides (Group 11) with a physical thickening mechanism utilizing a crystic polyester. The material is heated to melt the crystic and then the other ingredients are added, mixed together and run through a modified SMC machine maintaining the elevated temperature. The thickening occurs as the material cools to ambient temperature and the compound is ready to mold at that time. Cooling rolls speed the cooling process and thus the material can be molded right off the SMC machine without waiting for the 2 4 4 8 h thickening of standard SMC. LPMC allows the molder to use lower tonnage presses to mold larger parts and use less steel in building the tools as they do not have to deal with high pressures and corresponding forces. The shelf life of LPMC is much longer than SMC and the physical properties are comparable. 17.2.6 CONTINUOUS IMPREGNATED COMPOUND (CIC)

In 1986, continuous impregnated compound (CIC) was developed in Germany. This is similar to TMC. Like TMC, the impregnation is made between two rolls but the compound is removed by doctor blades and carried by a screw or plunger to boxes or drums. CIC is usually injection molded, but can also be injection/compression molded. Properties are comparable to BMC but processing is easier than BMC. Modified CIC is also known as KMC (Kneaded Molding Compound).

382 Matched metal compression molding of polymer composites 17.3 FORMULATIONS

Polymer composites have the unique and distinct advantage in that their properties can be tailored to meet different applications by designing the formulations. The major components of polymer composites used in matched metal compression molding are resin, low profile additive, fiber, filler, initiator, inhibitor, internal mold release agent and other additives (e.g. viscosity reducer, toughness enhancer, etc.).

butadiene copolymers. Extensive details of LPA mechanism are published in the literature9J8J9. 17.3.3 INITIATORS Ah-D INHIBITORS

Initiators are used to initiate the curing reaction at elevated temperatures. Composites are polymerized or crosslinked by a free radical mechanism in which the double bond of the polyester chain reacts with the vinyl monomer (usually styrene). This copolymerization reaction provides a three-dimensional network that converts the viscous liquid resin 17.3.1 RESIN paste into a hard thermoset solid. Initiators Unsaturated polyesters and vinyl esters are the decompose at elevated temperature and provide a source of free radicals to initiate the principal resins used in polymer composites for compression molding. Epoxies are also used for copolymerization reaction. Peroxyesters and specialty products which require longer cure peroxyketals are the most common classes of cycles and higher strength. Phenolics are being peroxides. Inhibitors are added in small quantities to used for formulating composites, especially prolong shelf life, modify cure rate and magSMC, in applicationswluch require lower flamnitude of exotherm to prevent cracking of mability, reduced smoke generation and higher thick molded sections. Inhibitors are also thermal stability’O. Details of resin chemistry used to improve resin stability. Two general are available in the literaturel’-l4. Styrene is classes of inhibitors are commonly used, subcommonly used for cross-linking of both polystituted phenolic derivatives and the ester and vinyl ester resins. Low styrene quaternary ammonium salts (e.g. hydropolye~ter’~ is becoming popular due to strinquinone, p-benzoquinone, etc). An excellent gent EPA requirements on styrene vapors. New review on initiator and inhibitor chemistry is resin technology is also being considered for available elsewhere y. compression molding. They include hybrids of unsaturated polyester and urethane16,acrylesterol resin with polyi~ocyanate’~, etc. 17.3.4 FILLERS

Fillers are used to improve physical properties, reduce volumetric shrinkage of the resin and to reduce costs. Fillers are typically divided into Low profile additives are thermoplastics that functional and non-functional categories. are added to the formulation in 2-5% (by Examples of functional fillers include alumina weight) of the final product or 10-20% (by trihydrate for flame retardancy, hollow glass weight) of the organic portion of the formulabubbles for lower weights, mica and wollastion to control the shrinkage of the cured tonite for reinforcement. Non-functional fillers composites. Typical thermoplastics include are used for cost reduction and are mineral polyvinyl acetates, poly methyl methacrylate based. Ground limestone (CaCO,) is the most and copolymers with other acrylate, vinyl chloAn excellent review is common type of filler. ride-vinyl acetate copolymers, polyurethane, available on filler use in composite^^^^^^ polystyrene, polycaprolactone, cellulose acetate butyrate, saturated polyester and styrene17.3.2 LOW PROFILE ADDITIVES

Molding 383 17.3.5 FIBERS

17.3.8 OTHER ADDITIVES

Glass fiber reinforcement is used to achieve Pigments are added to produce color in the necessary dimensional stability and mechani- molded part. Common pigments include cadcal properties. E-glass is the most common mium salts, carbon black, titanium dioxide, fiber reinforcement for composites. iron oxides, organic dyes and pigments, etc. Depending on the binder chemistry and Various types of viscosity reducers are used to amount, glass fibers are classified as hard or lower the viscosity of the paste to increase soft type. Other types of fibers include carbon, filler loading and glass wet-out. Other addiaramid (Kevlar), S-2 glass, etc. Glass loading tives include various elastomeric additives normally averages 30% by weight in compres- (e.g. Hycar, Kraton, etc.) to increase toughness. sion molded composites, but can vary from Tables 17.3 to 17.6 show typical formula18-65?” by weight. Table 17.2” shows typical tions for BMC, SMC, ZMC and LPMC. fibers used in polymer composites. 17.4 MOLDING

17.3.6 INTERNAL MOLD RELEASE

Internal mold release agents are added to facilitate part ejection from the mold. Major types of release agents include metallic stearates, fatty acids, fatty acid amides and esters and hydrocarbon waxes. Zinc stearate and calcium stearate are the most widely used internal mold release agents in SMC and BMC. 17.3.7 THICKENERS

The addition of Group I1 oxides and hydroxides to carboxy-terminated unsaturated polyester/vinyl ester resin increases its viscosity. Magnesium oxide, magnesium hydroxide, calcium oxide, calcium hydroxide or combinations of those materials are the most popular thickeners.

Matched metal compression molding is one of the oldest manufacturing techniques in the plastics/composites industry. The recent development of high strength, fast cure, Table 17.3 BMC formulation

PHR Polyester resin Low profile additive Styrene Initiator Inhibitor Mold release Pigment Thickening agent Filler Compound Glass fiber Paste

60.0 40.0 5.0 1.5 Trace amount 4.0 0.25 1.0 50-200 10-25% 75-90%

Table 17.2 Glass fibers used in compositesz1 Fiber

E-glass S2-glass Carbon (graphite) Kevlar 49

Specific gravity

Tensile strength (GPa)

Tensile modulus (GPa)

Tensile failure strain (%)

Coeficient of thermal expansion (x 1 @6/OC)

2.54 2.48 1.76-2.15

3.45 4.30 1.5-5.6

72.4 86.9 220-690

4.8 5.0 0.3-1.2

1.45

3.62

131

2.8

5 2.9 -0.1to -1.2 (longitudinal) 7-12 (radial) -2 (longitudinal) 59 (radial)

384 Matched metal compression molding of polymer composites Table 17.4 SMC formulation

PHR Polyester resin Low profile additive Styrene Initiator Inhibitor Mold release Pigment Thickening agent Filler Compound Glass fiber (25.4 mm) Paste

55.0 40.0 5.0 1.5 250 ppm 4.0 1.0 2.0 150-250 25-30% 70-75%

physical and mechanical properties can be obtained in compression molded parts. Figure 17.1 shows a schematic of a compression molding process. This section addresses the compression molding of composite parts using SMC. BMC molding is similar except for the charge preparation step. The compression molding process can be divided into four distinct steps. Heat and pressure

+

Cavity

Table 17.5 LPMC formulation

PHR Polyester resin Crystic Styrene Initiator Mold release Pigment Filler Compound Glass fiber Paste

65.0 15.0 5.0 1.2 5.0 1.2 220 4

25-30% 70-75%

Table 17.6 ZMC formulation

PHR Polyester resin Low profile additive Styrene Initiator Inhibitor Mold release Filler Compound Glass fiber Paste

65.0 40.0 5.0 1.5 100 ppm 4.0 220

SMC/BMC and advancement in press technology is making the compression molding process very popular for mass production of composite parts. In comparison with the injection molding process, in general, better

I Heat and pressure

Fig. 17.1 Schematic of a compression molding process.

17.4.1 CHARGE PREPARATION AND PLACEMENT

When the SMC has reached its desired molding viscosity, pieces of SMC are cut to pre-specified size after removing the carrier films. The SMC is cut using slitters, pizza type or guillotine type cutters. Several pieces of SMC plied together form the 'charge'. The charge pattern/ply dimensions are chosen so as to cover 20-80% of the mold surface area. The charge pattern and placement on the mold determines the quality of the molded parts, since it influences the length of flow in the mold, fiber orientation,

Properties 385 flowline and other surface defects. In order to reduce cycle time, sometimes the charge is preheated to a temperature below gel point using infra-red or dielectric heaters. 17.4.2 MOLD CLOSING AND FILLING

entire surface. After IMC injection, the press is closed and the curing operation is repeated at or above the SMC molding pressure. Sometimes the IMC is injected at high pressure without mold opening and closing prior to complete cure of the SMC charge.

After proper placement of the charge in the core of the mold, the cavity is quickly closed to 17.4.4 PART EJECTION AND POST-CURE contact the top surface of the charge. The cavAt the end of the molding cycle, the mold is ity is then closed at a slower rate, usually 4-12 opened and the part is ejected from the core mm/s. In most cases the mold is heated to (for with the use of integral ejector pins and example) 150°C, which causes the charge visallowed to cool to ambient temperature. Hot cosity to be reduced. With increasing mold parts are handled carefully and are usually pressure as the mold is closed, the charge placed on a support racks to cool to ambient flows towards the cavity extremities, forcing temperature. As the part cools outside the air out of the cavity. The mold closing speed is mold, it continues to cure and shrink which very important as it induces gelation of the top creates residual stresses due to differential charge surface if the closing speed is slow or it cooling at various sections in the part. After causes trapped air if closing speed is fast. The the part is placed on support rack, it is filling stage is usually completed in 0.5-20 s5. deflashed while still hot and stored in racks for Vacuum molding is increasingly being used secondary operations like punching, drilling, during charge flow to reduce surface porosity bonding, etc. and air entrapment in the part. Vacuum level The compression molding process is comis usually in the range of 7-9 x lo4Pa (21-27 in plex and there are several important variables Hg). The molding pressure based on projected that influence molding. Compression molding part area ranges from 1 to 10MPa (100 to may also produce a variety of surface and 1200 psi). Higher molding pressure causes internal defects which can be eliminated by sink marks, while lower pressure cause scumproper material selection, part design and ming of the mold and porosity. molding technique. Details of the molding variables and the source and remedies of 17.4.3 CURING major molding defects are available in the literature5,'. After filling, the charge remains in the hot mold for the crosslinking reaction to be completed. The curing time is usually between 25 s 17.5 PROPERTIES to 3 min, but depends on several factors, including resin-initiator-inhibitor reactivity, The properties of a polymer composite can be tailored, within limitations, to meet different part thickness and mold temperature. Sometimes in class A or appearance grade applications by designing its formulation. This parts, in-mold coating (IMC) is used to unique characteristic of polymer composites enhance the surface of a molded part. The makes definition of detailed properties diffimost common method of IMC injection cult. The properties are usually used for requires opening the mold by a small amount information and guidelines for preliminary (0.2-0.5 mm) after the curing cycle. IMC is part design, material selection and to underusually a coating of polyester or stand the effect of formulation variables on polyester-urethane hybrid which covers the mechanical properties.

386 Matched metal compression molding of polymer composites 17.5.1 STATIC PROPERTIES

17.5.3 OTHER PROPERTIES

Table 17.7 shows the static and impact properties of SMC, BMC, ZMC and LPMC. In general, tensile and flexural properties are routinely measured and are presented here. Compressive and shear properties are measured only for use in special applications. The static properties of SMC and BMC are highly dependent on the fiber content, length, type and orientation. Tensile strength increases significantly with increasing fiber content; however, the tensile modulus is affected only moderately. Increasing the length of chopped fiber increases the tensile strength, but has no effect on the modulus. Glass fiber type (E-glass or S-glass) has a significant effect on both the tensile strength and modulus. The resin chemistry also influences tensile properties at low fiber content. In general, flexural and compressive properties follow the same trend as the tensile properties. Flexural strength is always higher than tensile strength, though the modulus may be comparable.

Several other tests are now being performed to correlate properties with operation conditions. The dynamic mechanical analyzer (DMA) is used to measure complex and storage modulus at various temperatures and frequency ranges. The effect of environmental conditions on various properties is tested to simulate end-use environment. Creep and stress relaxation tests are also done on SMC/BMC for use in structural applications. Electrical properties are also important, permitting BMC to be used in electrical applications. Arc resistance is important and dielectric strength, dielectric constant, dissipation factor, etc. are also measured.

17.5.2 FATIGUE PROPERTIES

The fatigue properties of SMC and BMC are usually based on tensile cyclic loading of unnotched specimen. A typical S-N diagram is shown in Fig. 17.221.In general, the fatigue strength increases with increasing fiber content and there is no fatigue limit, unlike low carbon steel. Details of such testing are published elsewherez1,".

17.6 APPLICATIONS

Reinforced composites materials offer the maximum design versatility and capability of any material. With the excellent cost/performance characteristics of reinforced composites, the variety and quantity of products being produced with these materials grow annually around the world. Matched metal molded reinforced composites should be considered when the finished product can be enhanced by one or more of the following characteristics. Part consolidation Reinforced composites can be molded in three dimensions in one operation. Complex shapes that require multi-piece assembly using materials, such as wood or steel, may be molded in one step with the use of ribs, bosses and varying wall thickness.

Table 17.7 Static and impact properties"

Property

Tensile strength (MPa)

Tensile modulus (GPa)

Flexural strength (MPa)

Flexural modulus (GPa)

IZOD impact (unnotched)

Specific gravity

Coeficient of thermal expansion (x 1 P P C )

SMC BMC ZMC LPMC a

65-100 30-70 30-70 65-100

Published industry data

9.5-14 8-12 8.5-12.5 9.5-14

130-200 50-150 50-150 120-200

8-14 9-1 7 7-12 8-14

600-1200 100-700 200-500 600-1200

1.3-2.0 1.7-2.1 1.8-2.0 1.8-2.0

8-14 15-20 11-27 7-10

Applications 387 100 -

R = 0.05 80 -

a" z 6 v)

-

60-

cn

E 2 .-E

40-

X

r" 20 -

01 0.1

I

I

I

I

I

I

I

1

10

102

103

104

105

106

Number of cycles, N Fig. 17.2 Typical fatigue !+N diagram for SMC (21) A: at 40°C; 0:at 23°C; Cl:at 93°C. [Reproduced from Composite Materials Technology: Processes and Properties (ed P.K. Mallick and 5. Newman) by permission of

the publisher.]

Light weight Reinforced composites offer a greater strength-to-weight ratio than most non-reinforced plastics and many metals. Dimensional stability Reinforced composites can maintain the critical tolerances required of the most demanding applications. Composites meet the most stringent material stiffness, dimensional tolerance, weight and cost criteria in many diverse applications. High strength Reinforced composites have excellent strength-to-weight properties. By weight, reinforced composites surpass the tensile strength of iron, carbon and stainless steels. Many glass reinforced compounds equal or exceed the flexural strength and impact resistance of metals23.

from most organic chemicals and can be formulated to resist acidic and basic solutions. Electrical resistance Reinforced composites are very poor conductors of electricity. As such, they have a lvgh dielectric strength for application in the electrical and electronic industriesz3. Resistance to minor impact Reinforced composite components have an excellent memory characteristic. Instead of yielding or deforming under minor impact as with steel, a reinforced composite panel will deflect and spring back to its original surface form4(Fig. 17.3).

Surface quality Reinforced composites can achieve a variety of surface textures, from very smooth and glossy to a rough texture. Corrosion resistance Reinforced composites Insignias and alphanumeric characters can be do not rust or corrode, are resistant to attack molded as raised or indented characters.

388 Matched metal compression molding of polymer composites and performance requirements of the product or component. The designer must:

Steel

-

///// Composite

Fig. 17.3 Minor impact.

Molded-in color Color can be added to the reinforced composite compound, often eliminating the need for a secondary painting process. Recycling Most reinforced composites can be recycled either by regrinding or pyrolysis. Reground material can be used as filler or reinforcing material. Pyrolysis reduces the composite into its basic components by heating the material in the absence of oxygen. The process yields gas, oil and solid by-products that can be recycled back into composites, or used in building and agriculture materials 4. Thousands of products are molded each year utilizing reinforced composites: aerospace, automotive parts, sports and recreational equipment, boats and business machines to name a few. This wide variety of applications is indicative of the versatility, capability and cost effectiveness of reinforced composites.

1. establish size and shape limitations based on: 0 basic end use function; 0 aesthetics and marketing; 0 shipping limitations; 0 weight requirements; 0 strength and stiffness requirements; 0 flexibility requirements or limits; 0 process limitations. 2. establish the structural requirements based on: various loads that will be impacted to the part including weight, pressure and dynamic loads; duration of the loads on the part; temperature variations on the part and surface; number of cycles of temperature change; liquid, moisture and vapor resistance requirements; relative significance of strength-toweight ratios.

17.7 DESIGN CONSIDERATIONS

3. establish the non-structural requirements based on: 0 corrosion, weathering, moisture and temperature resistance; 0 moisture and vapor penetration for condensation protection; 0 fire safety relative to combustibility; 0 flame-spread rate requirements; 0 light transmission (transparency,translucency and opaqueness); 0 surface textures, both aesthetic and functional; 0 surface coatings for protection or aesthetics; 0 thermal insulation; noise and sound control; 0 dielectric requirements for electrical insulation24.

Given the wide range of options provided by reinforced composites, it is imperative that the designer accurately establish the functional

With the establishment of the functional and performance requirements the product design can be developed.

Design considerations 389 There are some general design principles which can assist in the development of structurally efficient configurations for reinforced composite components. 17.7.1 SHELL AND PLATE CONSTRUCTION

These are the most common configurations of reinforced composite parts. A reinforced composite component is constructed from layers of reinforced composite materials molded into the shape desired creating a geometric 'shell'. It is good practice to design so that only forces that place a part in tension or compression are applied to any component. Compound curve shapes provide good transmission of uniform loads into tensile and compressive forces within a part. Ribbed configurations are often used to achieve required strength and stiffness in structural components (Fig. 17.4). Corrugated or open ribbed configurations are used to achieve needed structural depth while efficiently using materials and fabrication processes (Fig. 17.5). 17.7.2 DRAFT

Draft is a slight angle introduced relative to the direction of the opening and closing of the mold. It is necessary to design the part so that all side walls, both interior and exterior, have draft. This enables the part to be removed from the mold without hanging or rubbing, which

Fig. 17.4 Rib configuration.

Circular

jf

Pitch

)t '

Rectangular

Fig. 17.5 Corrugated configuration.

can degrade 'appearance' surfaces (Fig. 17.6). Minimum draft angle of 1" for the first 76.2 mm (3 in) of depth, 2" for 76.2-101.6 mm ( 3 4 in) of depth, 3" for 101.6-127 mm (4-6 in) of depth and 1" for every additional 50.8 mm (2in) thereafter is recommended on all surfaces parallel to the mold movement. This pertains to all part details, such as ribs, bosses, elevation changes and holes. A draft angle of 1" on standing ribs and bosses will yield a thickness change of 0.43 mm (0.017 in) per inch per sidez4. 'Zero draft' may be obtained by designing the mold in such a way so the draft-free surface lies at an angle to the mold direction. This will affect the positioning of bosses, ribs and other details of the part.

390 Matched metal compression molding of polymer composites Draft angle (1’ recommended) 1Ae” Minimum Recommended

‘By

Radius Determined Part Thickness

Fig. 17.8 Minimum draft. mold movement

Fiig. 17.6 Draft.

17.7.3 RADIUS

17.7.4 NOMINAL THICKNESS

In mold making, the radius defines the curvature established between two intersecting surfaces. The more generous the radius, the better the flow of molding material for a stronger part (Fig. 17.7). A minimum radius of 1.59 mm (1/16 in) is recommended for all radii for both interior and exterior plane intersections. Radii should be designed to maintain relatively uniform part thickness (Fig. 17.8). Ribs and bosses opposite an appearance surface should have the radii eliminated to reduce the likelihood of warpage or ’sink’ (surface depression).

The nominal thickness is the overall design thickness of most of the part. It is desirable to establish uniform thickness throughout a part, to achieve minimum cure time, uniform cooling and minimize warpage and shrinkage (Fig. 17.9). Nominal thickness for reinforced composites is 2.544.57 mm (0.100-0.180 in). Recommended minimum thickness is 1.53mm (0.060 in). Recommended maximum thickness 25.4 mm (1.00 in). By designing hollow ribs, bosses and elevation changes can achieve intricate part geometry while maintaining nominal thickness throughout the part. 17.7.5 EDGE STIFFENING

Fig. 17.7 Outside radii.

Edge stiffening is a design characteristic applied to unsupported edges to prevent warping or bowing. Edge turning is preferable to edge thickening due to the possibility of porosity at the edge of part caused by the lack of molding pressure on the thick area (Fig. 17.10).

Design considerations 391

Fig.17.9 Nominal thickness. Uniform thickness promotes uniform flow and curing and minimizes the risk of warpage, distortion and telegraphing at thickness changes through the surface. 17.7.6 RIBS

Linear projections 90" from the plane surface of a part are called ribs. The use of ribs will allow the part to meet strength and rigidity requirements, preventing warpage and bowing in large plane surfaces while reducing the bulk and mass of a part. Ribs should be designed to maintain the nominal thickness and follow the guidelines

for draft angles. They should be dimensioned so that their thickness at the juncture of the rib with its plane surface is between 75 and 90% of nominal (Figs 17.11 and 17.12).

4-

0.5"draft

Fig. 17.11 Rib geometry for class 'A' surfaces.

(0.06'') radius

+1.0" draft Fig. 17.10 Edge stiffening (a) Preferred edge flange designs to increase panel stiffness; (b) Thickening

the edge flange may increase cycle times.

Fig. 17.12 Rib geometry for non-appearance surfaces.

392 Matched metal compression molding of polymer composites 17.7.7 BOSSES

17.7.9 MOLDED-IN THREADS

Projections from a plane surface of a part, called bosses, provide attachment and support for related components. They may be solid, hollow or have molded in inserts. They should also follow the guidelines for draft angles and nominal thickness (Fig. 17.13).

It is difficult to mold a thread into reinforced composites and requires highly sophisticated and costly molds and molding procedures. Molded threads should be rounded rather than sharp. Rounded threads will resist chipping and cracking and will also facilitate flow of molding material into all areas of the thread. Molded threads are usually preferred over inserts if the threaded hole diameter is over 12.7mrn (0.5 in), unless the thread is to be subjected to continual fastening and unfasteningz4. 17.7.10 MOLDED SURFACES

Nominal thickness should be maintainedthroughout part

I

Fig. 17.13 Boss design.

17.7.8 INSERTS

Inserts are objects (usually metal) which are molded into a part to facilitate repetitive fastening and unfastening of associated parts and can be provided with male or female threads. Inserts can provide bearing or bushing surfaces, electrical or other mechanical connections. Inserts should have knurls, grooves or shoulders to lock them in place and should be located parallel to the direction of mold travel.

Surfaces exhibited by the part as it comes from the mold have not been subject to any postmolding operation other than the removal of flash. The surface of the mold will reproduce itself as the surface of the part. Many different textures can be produced on the surface of a part. High gloss surfaces can be produced by highly polished molds. Draft is critical when parts are to be textured on a vertical wall. For every 0.254 mm (0.001 in) of texture depth, draft must be increased by 1". Raised or indented characters can be molded into the part. The characters should be rounded and smooth and positioned on the surfaces parallel to the parting line of the mold. 17.8 TOOLING

A good set of matched metal chromed steel tools is required for the optimum conditions when one is molding reinforced composites. Anticipated production quantities expected from the mold or the product end use, or both, should dictate the choice of steel as shown in Table 17.B5. 17.8.1 MOLD STRESSES

It is important to consider the stresses created by the flow of material at typical molding

Tooling 393 pressures from 4.13 to 8.37 MPa (600 to 1200 forced composite parts should be designed so psi). Due to unbalanced flow, narrow mold the height of any projecting mold section does sections that project from the mold surface not exceed two times the width of its base. could bend or break under such stresses. To Angular sections must not be less than 30°4 ensure sufficient strength in the mold, rein- (Figs 17.14 and 17.15).

Table 17.8 Mold steel selection5 A . Production planning volumes

Type of steel Planning volumes

Core

Cavity

5000-20 000 parts/y

AISIa-1045 steel

AISI-1045 steel

20 000-30 000 parts/y

AIS14140 forged steel prehardened to Rockwell C of 28-32


Over 30 000 parts/y


P-20 forged steel prehardened to Rockwell C of 28-32

100 000 parts or less

AISI-1045 steel

AISI-1045 steel

100 000-200 000 parts for mold life



Over 200 000 parts during mold life

AIS14140 forged steel prehardened Rockwell C of 28-32

P-20 forged steel prehardened to Rockwell C of 2&32

AISI-1045

AISI-1045 steel

AIS14140

P-20 forged steel

B. Product end use Structural items where surface appearance is not critical, such as reinforcing panels, truck front ends, etc., where molded surface quality is of secondary importance. High-quality surface appearance decorative items, such as grille opening panels, head lamp surroundings, quarter wheel opening covers, etc. where a high degree of polish is required on the outer part of cavity surface. a

American Iron and Steel Institute

394 Matched metal compression molding of polymer composites

B

Fig. 17.14 Projecting mold section; any projecting Fig. 17.15 Angular mold section: should not be mold section should not exceed two times the 06 SCCTION

05

-

04

.-.

03

Fig. 18.16 Unit cell geometry of plain weave.

02 01

00

The fabric thickness is very close to two yarn diameter, i.e.

0

10 20

30 40 50 60 70 80 90 0()

T=2d

(18'11)

Fig. 18.17 Relationship of fiber volume fraction to fiber orientation for plain weave.

1 - 1 d tan8

(18.12)

18.4 KNITTING

and approximately:

Equation (18.9) is then simplified to

18.4.1 PROCESSING TECHNOLOGY

Knitting is the interlocking of one or more yarns through a series of loops (also called stitches). The lengthwise columns of stitches corresponding to the warp in woven fabrics are called wales; the crosswise rows of stitches Figure 18.17 plots the fiber volume fraction corresponding to the filling are known as against the yarn inclination angle. It can be courses. Knitted structures can be classified by seen that as the inclination angle increases, the basic loop formation mechanism into weft pitch length becomes longer which results in a knits and warp knits. In weft knitting, as lower fiber volume fraction. The woven fabric shown in Fig. 18.18(a),yam feeding and loop has the tightest structure at the inclination formation occur at each needle in succession angle of 60" (when Lld = 3 in equation (18.10)). along the wale direction and all the courses of In this calculation, the fiber packing fraction K loops are composed of single strands of yam. is assumed to be 0.8. In warp knitting, there is a simultaneous yarnThe above analysis is given only for the sim- feeding and loop-forming action occurring at plest of woven structures. Different weave every needle and all the wales of loops are patterns, non-circular yarn cross-sectional composed of single strands of yarn as illusshape, different yarn dimensions and pitch trated in Fig. 18.18(b).

Knitting 409

J

’Djrection

of knitting

Fig. 18.18 Yarn feeding and loop formation: (a) weft knitting; (b) warp knitting (Spencer, 1983).

technology can be found in Spencer (1983)and Raz (1987). Knitted 3-D fabrics are produced either by weft or warp knitting. An example of a weft knit is the near net shape structure knitted under computer control by the Pressure Foot@ process (Williams, 1978). In a collapsed form this preform has been used for carbon-carbon aircraft brakes. While weft knitted structures have applications in limited areas, multiaxial warp knit (MWK) 3-D structures are more promising and have undergone a great deal more development in recent years. Schematic of a MWK LIBA system is given in Fig. 18.20, in which up to six layers of insertion yarns plus one layer of non-woven can be stitched together.

Stitch (loop) formation is similar in both weft and warp knitting. The formation of the stitches in a single wale is illustrated in Fig. 18.19. In Step 1, the needle rises through loop A from its lowest position; in Step 2, yarn slips under the tip of the needle and onto the stem; in Step 3, ascending hook catches the new yarn at the top of its rise and begins to descend; in Step 4, the new yarn slips under the tip and into the hook; in Step 5, the needle moves down until the tip slides under loop A and the hook pulls the new loop through. After the completion of five steps, loop B is formed and the process is repeated. In a knitting operation, each of the needles is controlled by a cam to rise and fall in synchronization with the other needles. Detailed description of the knitting ,Hook

Ti

1

2

3

Fig. 18.19 Stitch formation in knitting machines (Smith and Block, 1982).

410 Textile preforming

Fig. 18.20 Multiaxial warp knit with four layers ( O O , 90" and &) of inserted yarns and (a) chain stitch or (b) tricot stitch.

18.4.2 STRUCTURAL GEOMETRY

ture for the incorporation of 0" and/or 90" insertion yarns. Knitted fabrics are traditionally identified The MWK fabric system consists of warp (O"), with socks, underwear and sweaters. In the yarns held together by search for methods to reduce composite man- weft (90") and bias (d) a chain or tricot stitch through the thickness of ufacturing costs, textile preforms including the fabric, as illustrated in Fig. 18.21. knitted structures are receiving increased Theoretically, the MWK can be made to as interest in the composite industry. While conmany layers of multiaxial yams as needed, but formability and productivity are obvious current commercially available machines only attributes for knitted preforms, the availability of a broad range of micro- and macrostructural allow four layers (the Mayer system) of 0", 90", geometries has only recently been recognized. +O and 4 insertion yams, or six layers (the The non-linearity of knitting loops, severe LIBA system) of 2(90"), 0", 2(+8) and 2 ( 4 ) bending of yams during the knitting process insertion yarns to be stitched together. All layand limited fiber packing density resulting in ers of insertion yarns are placed in perfect the formation of resin pockets within a knit- order each on top of the other in the knitting ting loop prevent kmts from being considered process. Each layer shows the uniformity of the uncrimped parallel yams. The insertion yarns for structural applications. The development of technology for the usually possess a much higher linear density directional insertion of linear yarns in weft than the stitch yarns and are therefore the and warp knits greatly enhances opportunities major load bearing component of the fabric. for knitted preforms for conformable structural composites by combining the 18.4.3 DESIGN METHODOLOGY conformable foundation knit structure with directional reinforcement. As shown in Fig. Similar to the 2-D woven fabrics, the unit cells 18.20, sewing threads (high twist yams) or for the knitted structures are also different, very fine yams are used to form a base struc- depending on the knit constructions such as

Knitting 411

Fig. 18.21 Multiaxial warp knit LIBA system.

stitch patterns and laid-in insertions. To illustrate the use of the unit cell method for relating fiber volume fraction, yarn orientation and processing variables, a plain weft knit as shown in Fig. 18.18(a) is selected as an example. The unit cell geometry identified for the plain weft knit is shown in Fig. 18.22, having a dimension of x (course width), y (half wale width) and z (fabric thickness).

illustrated in Fig. 18.23(b).For untwisted fiber bundles under compression applied during preforming or composite processing, they have a ribbon-like cross-section similar to a race-track with a width-to-thickness aspect ratio off > 1 as illustrated in Fig. 18.23(a).For composite applications, untwisted fiber bundles are usually used in knitting, which have an aspect ratio f slightly larger than 1 at the off-machine state. To increase the fiber volume fraction for knitted structures, very high pressure will be required to reduce the knit thickness. Under the compression status, the yarn aspect ratio f can increase to as high as 12 for untwisted bundles, provided that there are no restrictions applied to yarn edges.

Fig. 18.22 Unit cell geometry of plain knitted struc-

ture. In traditional textile fabric manufacturing, highly twisted fiber bundles are used. These materials can maintain a circular shape with a width-to-thickness aspect ratio of f = 1, as

(a)

f =w/t

1

(b)

f =1

Fig. 18.23 Idealized yarn cross-sections: (a) racetrack cross-section with width-to-thickness aspect ratio f > 1; (b) race-track shape becomes circular whenf = 1.

412 Textile preforming

In this analysis, the knit thickness is assumed to be approximately equal to two yarn thickness (t)for computational purposes, i.e.

z = 2t

(18.14)

The yarn orientation angle (e), which is the angle made by the fabric axis (in x direction) and the yarn path projected to the fabric surface plane ( x - y), is given by: (18.15)

The fiber volume fraction (V,),which is defined as the ratio of volume of total fibers to the overall composite volume, can be derived as:

Y (18.16) 1+tan where k is the fiber packing fraction within yarn bundles and a is the shape correction factor defined as:

relative course width (x/w), relative half wale and yarn aspect ratio (f)under width (y/w) compression is depicted in Fig. 18.24, using the geometric model developed. In the calculation, we use the fiber packing fraction k = 0.8, which is within the range for tightly packed yarn bundles according to experimental observation. Also, to show the processing window of fiber volume fraction in highest region, one can assume yam jamming in the course (x) direction, i.e. x/w = 3 according to equation (18.18). As can be seen from Fig. 18.24, the fiber volume fraction V , decreases with the increase in relative half wale width in the range of y/w = 2-10. When y/w is beyond 10, the fiber volume fraction slightly increases and soon approaches a constant with the increase in relative half wale width. The wale width cannot be smaller than 4 yarn widths, or y/w 2 2 as given by equation (18.19). Knitted yarns have an aspect ratio f = 1 at free-stress status (as made off-machine) and the fiber volume fraction for the knitted preform has a minimum value. Figure 18.24 shows that, for the plain weft knit at its tightest possible structure (x/w = 3, y/w = 2 and k = OB), its maximum fiber volume fraction is only about 0.274. To increase the fiber volume fraction, a compression in the fabric thickness direction is necessary. The effect of the compression is the increase in yam aspect ratio (i.e. yams within the knitted structure become wider in x-y plane

(18.17)

060

The limiting geometry of the knitted structure due to yarn jamming is governed by: X W

-23

(18.18)

020

P LL

f = 1 (circular yarn without compressior 0.10

1

W

2

2

(18.19)

The processing window of fiber volume fraction for knitted structures within the possible ranges of key processing parameters, such as

ooo 10

1000

100

Ylw

Fig. 18-24 Processing window of fiber volume frattion for the plain knitted structure.

Braiding 413 but thinner in z direction).As a result, the yarn coverage over the fabric increases, whereas the volume of the preform decreases due to the decrease in fabric thickness. These two factors, the increased yarn coverage and decreased preform volume, raise the fiber volume fraction to a much higher level. As shown in Fig. 18.24, at a maximum aspect ratio f = 12, the fiber volume fraction can be as high as 0.475. A series of studies on the technology, structure and properties of the MWK preforms and composites have been reported by KO and his co-workers (1980,1982,1985,1986,198813).In a recent study, a unit cell based geometric model of the four-layer MWK structure as shown in Fig. 18.21was developed by Du and KO (1992). Based on the experimental observations, the unit cell geometry of the MWK fabric is identified and a geometric model is developed relating the fiber volume fraction and fiber orientation in terms of structural and processing parameters.

18.5 BRAIDING

18.5.1 PROCESSING TECHNOLOGY

Braiding is an old textile technology, traditionally used for the manufacture of a wide variety of linear products ranging from cables, electrical insulators and shoelaces to surgical sutures. Recognizing the high level of conformability and the damage resistance capability of braided structures, the composites industry had found structural applications for braided composites ranging from rocket launchers to automotive parts to aircraft structures. Two-dimensional braided structures are intertwined fibrous structures capable of forming structures with 0" and & fiber orientation. Although 2-D braids can be fabricated in tape form, the majority of braided structures are fabricated with a tubular geometry. Thickness is built up by overbraiding previously braided layers similar to a ply lay-up process. Braiding can take place vertically orhorizontally, but a majority of the composite braiders are horizon-

tal. A schematic of a horizontal braider is shown in Fig. 18.25.Although braiding is similar to filament winding in many ways, the major difference between braiding and filament winding is that braids are interlaced structures having as many as 144 or more interlacing per braiding cycle (or pick). Three-dimensional braiding technology is an extension of 2-D braiding technology in which the fabric is constructed by the intertwining or orthogonal interlacing of yarns to form an integral structure through position displacement. A unique feature of 3-D braids is their ability to provide through the thickness reinforcement of composites as well as their ready adaptability to the fabrication of a wide range of complex shapes ranging from solid rods to I-beams to thick-walled rocket nozzles. Three-dimensional braids have been produced on traditional Maypole machines for ropes and packings in solid, circular or square cross-sections. The yarn carrier movement is activated in a restricted fashion by horn gears. A 3-D cylindrical braiding machine of this form was recently introduced by Albany with some modification that the yarn carriers do not move through all the layers (Brookstein, 1991). 3-D braiding processes without using the horn gears, including Track and Column (Brown et al., 1988) and 2-Step (Popper and McConnell, 1987),have been developed since the late 1960s in the search for multidirectionally reinforced composites for aerospace applications. A generalized schematic of a 3-D braiding process is shown in Fig. 18.26. Axial yarns, if present in a particular braid, are fed directly Axial yarns,

/Carrier track

Fig. 18.25 Schematic of tubular braider with gantry system.

414 Textile pyeforrning

;;urbanism

,

Forming point

-

Convergence point

Fig. 18.26 Schematic of a generalized 3-D braider.

into the structure from packages located below the track plate. Braiding yarns are fed from bobbins mounted on carriers that move on the track plate. The pattern produced by the motion of the braiders relative to each other and the axial yarns establish the type of braid being formed, as well as the microstructure.

Track and column braiding is the most popular process in the manufacturing of 3-D braided preforms. The mechanism of these braiding methods differs from the traditional horn gear method only in the way the carriers are displaced to create the final braid geometry.Figure 18.27(a)shows a basic loom setup in a rectangular configuration. The carriers are arranged in tracks and columns to form the required shape and additional carriers are added to the outside of the array in alternating locations. Four steps of motion are imposed to the tracks and columns during a complete braiding cycle, resulting in the alternate x and y displacement of yam carriers, as shown in Fig. 18.27(b-e).The formation of shapes, such as T-beam and I-beam, is accomplished by proper positioning of the carriers and the joining of various rectangular groups through selected carrier movements. The track and column braiding machine can also be used to create 2-step braids and other similar 3-D structures by simply adding a certain number of axial yarns and removing most of the braiding yarns (Du and KO, 1993a).

Y

Track direction

Fig. 18.27 Formation of a rectangular 3-D track and column braid, using 4 tracks, 8 columns and 1 x 1 braiding pattern. (a) Initial loom setup; (b) Step 1: tracks move horizontally; odd tracks move to left and even tracks move to right; (c) Step 2: columns move vertically; odd columns move down and even columns move up; (d) Step 3: tracks move horizontally; odd tracks move to right and even tracks move to left; (e) Step 4: columns move vertically; odd columns move up and even columns move down.

Braiding 415 18.5.2 STRUCTURAL GEOMETRY

As with woven fabric, braids can be formed with different yarn interlacing patterns by simply changing relative position of carriers on the track ring. If one bias yarn continuously passes over one yarn and then under one yarn of the opposing group, the pattern is designated as 1/1 braid, or diamond braid as generally recognized. Other simple interlacing patterns in common use include 2/2,3/3,2/1 and 3/1 braids. Figure 18.28 shows the pattern of 2/2 braid with axial insertion. Among all these patterns, the 2/2 braid is the most popular and has been referred to as regular, standard, plain or flat braid. The path of axial yarns is independent of braid interlacing patterns, they are always over one group of bias yarns, but under the opposite group. The formation of shape and fiber architecture are illustrated in Fig. 18.29 which depicts the process of braiding over an axisymmetric shape of revolution. Braiding angle can range

from 5" in almost parallel yarn braid to approximately 85" in a hoop yarn braid, depending on the mandrel dimension, the machine speed ratio and the convergence length (Du et al., 1990). The 2-D braid can be defined as a fabric which consists of only two layers of bias yarns interlaced with each other. In 3-D braided structures, at least three layers of bias yarns go through the thickness in a zig-zag manner along the diagonal direction. Similar to the 2-D structure, longitudinal yarns can be incorporated in the 3-D braid for the enhancement of stiffness and strength in the length direction. Regardless of the difference in the carrier propelling mechanism, there are basically two types of 3-D braiding looms: rectangular and circular. The former is usually used to fabricate solid structures such as panel, I- and T-beam etc. and the latter for making thick wall tubular structures. Figure 18.30(a) shows a schematic of a 3-D braided slab.

Fig. 18.28 Yarn structure in 2-D braid: braiding yarns at & to braid axis, optional axial yarns at 0" to braid axis.

Fig. 18.29 Braid formation over a shaped mandrel.

Fig. 18.30 3-D braided solid slab (and its cross-section as seen on SEM).

416

Textile preforming

18.5.3 DESIGN METHODOLOGY

The unit cell geometry of 2-step braids has been reported by Du et al. (1991). Based on experimental observations, diamond and ribbon shapes for the axial and braider yarns, respectively, are assumed in their analysis. The unit cell was defined; pitch length and percentage of braider yarns were identified as key process parameters which control the braid microstructure and the jamming criterion for the 2-step braid was given. The traditional approach used in modeling 3-D braided composites is to artificially define a unit cell geometry for a 3-D braided structure without providing any relationship between processing variables and geometric parameters. All fibers in the unit cell are assumed to incline in four different diagonal directions, as well as along the longitudinal direction, if any. Fiber volume fraction is assumed to be either known or measured. The approach used in geometric modeling of textile structures is to first determine the dimension, shape and fiber architecture of the unit cell based on process and structural analysis; using the unit cell geometry identified, the relationship between processing variables and key geometric parameters can readily be established. The key geometric parameters of 3-D braids (which affect reinforcement capability and composite processability) include braider orientation, total fiber volume fraction, volume fraction of inter-yarn void and axial fiber percentage of total fibers. Although there are only two simple process parameters adjustable to control the microstructure of 3-D braids (speed ratio between braiding and take-up and linear density ratio of braider and axial yarns), the process-structure model of 3-D braid is complicated. Normally, yarn bundles consisting of numerous continuous filaments are used for fabric preforms, thus, the fabric microstructure has three levels: geometry of interfiber packing in the yarn bundle (fiber level),

cross-section of yarn bundles in the fabric (yam level) and orientation and distribution of fibers in the 3-D network (fabric level). The unit-cell technique is commonly used to establish the geometric relation. In most of 2-D fabrics a unit cell geometry is readily identified, but in complex 3-D fabrics it can be very difficult to define. The fiber volume fraction of a 3-D fabric depends on the level to which yarns pack against each other in the structure and the level to which fibers pack against each other in a yarn, as illustrated in Fig. 18.7. In addition to the level of packing fraction, the fibers also establish the yarn cross-sectional shape, i.e. yarn packing in fabrics. This shape plays a very significant role in determining how many fibers can be packed into a fabric. One good example is the yarn packing in 2-step braided preforms (Du et al., 1991). Due to the use of untwisted fiber bundles and high braiding tensions, cross-section of axial yarns in the 2step braid is deformed to prismatic shapes, giving most the compact yarn packing within the braided structure. For the track-and-column braids, the braiding tensions are lower compared to the 2-step braids and the crosssections of yarns actually have a polygonal shape. The microgeometric model for the trackand-column braid has been investigated by many researchers since the early 1980s (Pastore et al., 1990b; Li et al., 1990). The most recent one was given by Du and KO (1993a), which does not only relate geometric parameters and processing variables but also provides limiting braid geometry due to yarn jamming. In their analysis, the yarns are assumed as rigid circular rods. This assumption is valid when braiding at high yarn tensions. When low yarn tensions are used, yarn crimp will be introduced during braiding or during postpreforming processing due to distortion. This yarn waviness (crimp) may increase the fiber volume fraction of the braid with the sacrifice of directional reinforcing efficiency.

Braiding 417 Figure 18.31shows an idealized braid cross- where IC is the fiber packing fraction (fiber-tosection cut longitudinally at a 45" angle to the yarn area ratio). Due to the bulky fiber and braid surface. There are four groups of yarns nonlinear crimp nature, it is difficult to fabriinclined at angle 8 with the braid axis (z direc- cate the braid with tightest structure. In tion) in different directions; the yarns in each practice, the yarn orientation angle (braid group are parallel to each other within a spe- angle) is determined from the yarn diameter cific plane. Two groups of yarns are parallel to (d) and braid pitch length (kZ).The fiber volthe XI-z plane; the other two are parallel to the ume fraction is controlled by the braiding y'-z plane. The cutting plane is so selected that angle and the braid tightness factor. The govit cuts through the diameter of a group of yarns. erning equations are given below:

8

e

= sin-1

{((k,/d)2 +

4)

(kZ2 2d) (18.21)

where is the fabric tightness factor, which is within the range of 0 to ~ / 4This . tightness factor must be so selected that the required fiber volume fraction is achieved and also that the over-jamming condition is avoided. Figure 18.32 shows the V , 4 relationship prior to and at the jamming condition, based on the governing equations. The fiber packing fraction, K, is assumed as 0.785. As can be seen, there are three regions of fiber volume fraction. The upper region cannot be achieved due to the impossible fiber packing in a yarn bundle. Jamming occurs when the highest

A

C Fig. 18.31 Braid cross-section cut longitudinally at a 45" angle to the braid surface by the Y-y plane ABCD. z is the braid length direction.

The braid has the tightest structure when each yarn is in contact with all its neighboring yarns, in other words, the yarns are jammed against each other. At the jamming condition, the fiber volume fraction V ,as a function of the braid angle 8 was derived by Du and KO (1993a):

v,=-K2K

cos8 1 + cos%

(18.20)

1

0,51= 0.4 0.3 0.2 0.1

0.04 0

I

10

.

,

20

.

Y

,

30

40

50

60

70

80

90

0 (")

Fig. 18.32 Relationship of fiber volume fraction to braiding angle for various tightness factors. Fiber packing fraction K is assumed to be 0.785.

418 Textile preforming braiding angle is reached for a given fabric tightness factor q. The non-shaded region is the working window for a variety of V f 4 combinations. Clearly, for a given fabric tightness, the higher braiding angle gives higher fiber volume fraction and for a fixed braiding angle, the fiber volume fraction is greater at higher tightness factors. Theoretically, the 2-D braid can be considered as a single layer of 3-D braid. For prepreg or tape braiding without much change in yarn width and for the braiding of structures with constant cross-sections, the V f 4 relation is simple. For braiding of dry tows and structures with variable cross-sections wherein a dynamic interaction of the braiding machine and tow geometry takes place, there is a need for a more general representation of the kinematics of the braiding processes which allows for the tow width to vary over a limiting geometry. Two mathematical models have been developed, the first is the kinematics model (Du et al., 1990) which provides the relationship between the braiding angle and the braiding process parameters and the other is the unit cell model (Du and KO, 199313)which relates braiding angle to yarn geometry to predict fiber volume fractions V, along both the braiding and axial directions.

Fig. 18.33 Schematic of the Novoltexa process.

thickness fiber reinforcement. For illustration purposes, our analysis of fiber volume fraction distribution will be focused on orthogonal nonwoven 3-D fabrics. While woven 3-D fabrics have a long history of development and is clearly a product of the textile industry, the class of orthogonal nonwoven 3-D fabrics is a product of the twentieth century, developed in the aerospace industry for specific composite applications. Pioneered by aerospace companies such as 18.6 NONWOVEN General Electric and AVCO, the nonwoven 318.6.1 PROCESSING TECHNOLOGY D fabric technology was developed further by Nonwoven structures are fiber to fabric assem- Fiber Materials Incorporated. Recent progress blies produced by chemical, thermal or in the automation of the nonwoven 3-D fabric mechanical bonding or a combination of the manufacturing process was made in France by above. Starting with discrete fibers or continu- Aerospatiale (Pastenbaugh, 1988), in Japan by ous filaments (mostly tows), the fibers are Fukuta of the Research Institute for Polymers randomly distributed or preferentially ori- and Textiles (Fukuta et al., 1982 and 1984) and ented by dynamic combing (carding) or more recently by Mohammed (1989). Orthogonal nonwoven (ON) 3-D fabrics are hydrodynamic (waterjet) methods. The Novoltex@structure developed by SEI' fabricated by maintaining one stationary axis as shown in Fig. 18.33 (Geoghegan, 1988)is an either by yarn pre-deposition or a spacer rod example of a mechanically bonded structure which is subsequently retracted and replaced wherein multiple layers of oriented or random by an axial yam. The placement of the planar fiber webs are needled together to create an yarn systems is carried out by inserting the integrated structure which has through yarns orthogonal to the axial yarn system in an

Nonwoven 419

Fig. 18.34 Orthogonal nonwoven by direct method.

alternating manner. In Fig. 18.34, the method. 18.6.2 STRUCTURAL GEOMETRY of direct formation of ON 3-D fabric is shown Structural geomeh.ies resulting from the vari(Stover et al., 1971)* By proper Of ous processing techniques are shown in Fig. yarns prior to planar yarn placement, 18.35: (a) and (b) show the single bundle xyz the 3-D fabrics of various shapes and densities can fabrics in a rectangular and cylindrical shape; be produced. (c) demonstrates the multiple yarn bundle possibilities in the various directions.

Fig. 18.35 Orthogonal nonwoven fabrics.

420 Textile preforming 18.6.3 DESIGN METHODOLOGY

A unit cell geometry for the orthogonal nonwoven 3-D fabric is shown in Fig. 18.36, assuming circular cross-section for yarns in all three directions. The fiber volume fraction for

against d y / d x ratios, assuming a fiber packing fraction of 0.8. For all three levels of dJdx ratios, the fiber volume fraction first decreases with the increase in d y / d xratio, reaches a minimum and then increases. As can be seen in the figure, the maximum fiber volume fraction is about 0.63 at either high or low d y / d x ratios, whereas the minimum fiber volume fraction of about 0.47 is achieved when both d y / d x and d z / d xratios are equal to 1. 18.7 SUMMARY AND CONCLUSIONS

In this chapter, we have first discussed the role and importance of textile preforms in composite design processing and design, followed by classifying them into linear (1-D), planar (2-D) Fig. 18.36 Unit cell for orthogonal nonwoven strucand three-dimensional (3-D) fibrous assemture. blies. The objective of this chapter is to the 3-D ON structure can be shown to have the describe the design methodology of the fiber architecture for representative textile preform following form: structures currently used for composite reinn forcements. After a brief introduction to the Vf = -ICx 2 formation technology of each preform, its fabric structure is shown and the geometry of a dx‘(dy+ dZ)+ d$dX + dZ)+ q d x + dy) (18.23) unit cell is defined. The relationship between (dx + dy)(dx+ dz)(dy+ dZ) the engineering parameters (V,,0) and the key where dx, dy and dZ are diameters of the yarns processing variables (such as preform pattern, in x, y and z directions, respectively and IC is tightness factor 9 and linear density ratio etc.) within the range of achievable geometry is the fiber packing fraction of the yarns. Figure 18.37 plots the fiber volume fraction established from the geometric model. A summary of preform fabrication techniques has been given in Table 18.2. Table 18.3 0.65 gives a summary of the engineering and processing parameters. Ranges of fiber 0.60 orientation angle and fiber volume fraction for 5 each fabric preform commonly used for com0.55 posite reinforcements are also included in Table 18.3. It should be noted that although the achievable range of fiber volume fraction is 0.50 restricted by theoretical fabric geometric limits Minimum fiber volume fraction due to yarn jamming, it is possible that a 0.45 1 . ...I.. ._ ._I higher fiber volume fraction can be achieved 10.’ loo 10’ lo2 lo3 lo4 in reality because of the compressible nature of the preforms. A composite having a higher Fig. 18.37 Process window of fiber volume fraction fiber volume fraction can be made simply by squeezing the preform to a smaller mold for orthogonal nonwoven fabrics. ,

,

,

,

Summay and conclusions 421 Table 18.3 Engineering and processing parameters for textile preforms

Preform Linear assembly Roving Yam

Woven 2-D Biaxial 2-D Triaxial 3-D Woven

Non-woven

2-D Non-woven 3-D Orthogonal

Knit

Fiber orientation, 0 (") 0 - yarn surface helix angle 0=0 0=5-10

0.6 0.7

0, - yarn orientation in fabric plane Oc - yarn crimp angle Of = 0/90, Oc = 30 60 Of= 0/90/+30-60, Oc = 30 60 O6 = 0/90, 0- = 30 60

-

Vf

Processing parameter bundle

- 0.8 - 0.9

Bundle tension, transverse compression, fiber diameter, number of fibers, twist level

- 0.5 - 0.5 - 0.6

Ox - fiber/yam orientation along x axis 0, - fiber/yam orientation along y axis O2- fiber/yarn orientation along z axis Oxy - fiber distribution on fabric plane

Oxy = uniform distribution,

Or,O,O,

= 0

-

Fiber packing in yarn, fabric tightness factor, yam linear density ratios, pitch count, stitch pattern

- 0.7 - 0.6

Fiber packing in yam, fabric tightness factor, braid diameter, pitch length, braiding pattern, carrier number

Os - stitch yarn orientation

Braid

Os = 30 Os = 30

- 60 - 60,

Oi = 0/90/+30-60

0.2 0.3 0.3 0.6

-

0 -braiding angle 2-D Braid 3-D Braid

-

0 = 10 80 0 = 10 45

(2-D non-woven) fiber packing in fabric, fiber distribution (3-D orthogonal) fiber packing in yam, yam cross section, yarn linear density ratios

0.2 0.4 0.4 - 0.6

Oi - insertion yam orientation 2-D Weft knit 3-D MWK

Fiber packing in yarn, fabric tightness factor, yam linear density ratios, pitch count, weaving pattern

0.5 0.4

during the process of matrix addition; how- Assuming a tightness factor 7 of 0.573, possiever, a composite with a fiber volume fraction ble braiding angles range from 0 to 40". Young's moduli and Poisson's ratios of fiber higher than theoretical maximum will have a certain degree of fiber crimp and its fiber ori- and matrix are given as E, = 33.5 Msi, Em = 1.3 Msi, 2rf = 0.3 and urn= 0.11. The elastic conentation will also be distorted. The geometric models of textile preforms stants of the carbon-carbon composite was presented in this chapter provide a quantitative obtained from the Fabric Geometric Model communication link between the preform (FGM) (KO et al., 1987).Figure 18.38 shows the manufacturer, composite processors and prod- composite stiffness in different directions uct design engineers. By reducing fiber within the working window of fiber volume architecture and textile preforming processes fraction and fiber orientation. As can be seen, into engineering and processing parameters Young's modulus, Edav in the axial direction Vf, 8 and 17, rational composite design proce- decreases and in-plane shear modulus, Gh- lane, dures and process control guides can be increases with the increase in braiding angfe 8. established. For example, the mechanistic Young's moduli in both hoop and radial direcdesign of a composite product can be demon- tions, Ehmpand Eradial,have the same value at strated using a tubular 3-D braided zero braiding angle, but depart and both carbon-carbon composite as an example. increase as the braiding angle become higher.

422 Textile preforming 0.006 Braid Axis

Pressure drop: 60 psd

s2 -r?

e.-x 0.004c ._

L2

0.002Limiting fiber architecture

6-

2 0.000 0 _1

3

6

9

12

15

Fiber diameter (pm) 0, 0

,

5

.

.

, . , . , , , . , . 10 15 20 25 30 35 40 45 50 55 60 ~

Braid angle, 8 (")

Fig. 18.38 Stiffness properties of 3-D braided carbon-carbon composite. Fiber packing and tightness factor are assumed as IC = 0.785,~= 0.573; Young's moduli and Poisson's ratios of fiber and matrix are given as E , = 33.5 Msi, Em = 1.3Msi, z),= 0.3, urn= 011.

The other example of the application of the fiber architecture models to the composite processing is to predict the permeability of fabric preforms. As suggested by the well known Kozeny-Carman equation, there are two major geometric parameters which greatly affect permeability of fibrous materials, i.e. porosity of fabric preforms E and characteristic dimension of fibers @D,where @ is the shape factor and D is the fiber diameter. Other parameters which also affect the permeability are flow properties, pressure drop and part thickness. These parameters have been shown to be independent of preform fiber architecture. From geometric analysis, one can construct the V f 4 relationship and determine their dependence on the process parameters. The fabric porosity can easily be calculated from V , ( E = 1 - V,), whereas the shape factor of fibers @ is related to the fiber orientation 8 and the flow direction in composite processing. From our experiments, we have observed that @ = 1.5 when most of the fibers are aligned parallel to the flow direction and @ = 0.75 when the fibers are perpendicular to the flow direction. The effect of the fiber volume fraction and fiber diameter on the permeability of air flow

Fig. 18.39 Effect of fiber volume fraction and diameter on preform permeability.

can be quantified using the Kozeny-Carman equation, as shown in Fig. 18.39, noting that the permeability is in the unit of mass flow rate per length of preform. In this example, the shape factor @ is assumed to be 0.75 at a pressure drop across the preform of 60 psi with a fabric thickness of 0.5 in. REFERENCES Brookstein, D.S. 1990. Interlocked Fiber Architecture: Braided and Woven. Proc. 35th Intern. SAMPE Symposium, Society for the Advancement of Material and Process Engineering, Vol35, pp. 746-756. Brown, R.T., Patterson, G.A. and Carper, D.M. 1988. Performance of 3-D Braided Composite Structures. Proceedings of the Third Structural Textile Symposium, Drexel University, Philadelphia, PA. Chou, T.W. and KO,F.K., eds. 1989. Textile Structural Composites. New York Elsevier. Dow, N.F. and Tranfield, G. 1970. Preliminary Investigations of Feasibility of Weaving Triaxial Fabrics (Dow Weave). Textile Research Journal, pp. 986-998. Dow, N.F. 1985. Woven Fabric Reinforced Composites for Automotive Applications. Technical Final Report, NSF Grant No. DMR8212867, MSC TFR 1605/8102, December. Dow, N.F. and Ramnath, V. 1987.Analysis of Woven Fabrics for Reinforced Composite Materials. NASA Contract Report 178275. Dow, R.M. 1989. New Concept for Multiple Directional Fabric Formation. Proc. 21st Intern. S A M P E Tech. Conf., September 2.528.

References 423 Du, G.W., Popper, P. and Chou, T.W. 1990. Process Model of Circular Braiding for ComplexShaped Preform Manufacturing. Proc. Symp. on Processing of Polymers and Polymeric Composites, ASME Winter Annual Meeting, Dallas, Texas, NOV.25-31. Du, G.W., Popper, P. and Chou, T.W. 1991. Analysis of Textile Preforms for Multi-directional Reinforcement of Composites. 1. Mater. Sci. 26: 3438-3448. Du, G.W. and KO, F.K. 1992. Analysis of Multiaxial Warp Knitted Preforms for Composite Reinforcement. Proc. Textile Composites in Building Construction 2nd Inter. Symp., Lyon, France, June 23-25. Du, G.W. and KO,EK. 1993a. Unit Cell Geometry of 3-D Braided Structures. J. Rein. Plus. Comp. 12 (2): pp. 752-765. Du, G.W. and KO, EK. 1993b. Analysis And Design Of 2-D Braided Preforms For Composite Reinforcement. Proc. ICCM-9, Madrid, Spain, July 12-16. Fukuta, K., Aoki, E. and Nagatsuka, Y. 1984. 3-D Fabrics for Structural Composites. 15th Textile Res. Symp., The Textile Machinery Society of Japan, Osaka, Japan. Fukuta, K., Onooka, R., Aoki, E. and Nagatsuka, Y. 1982. Application of Latticed Structural Composite Materials with Three Dimensional Fabrics to Artificial Bones. Bull. Res. Inst. Polym. Textiles. 131(2)2:151. Geoghegan, P.J. 1988. DuPont Ceramics for Structural Applications - the SEP Noveltex Technology. 3rd Textile Structural Composites Symp., Philadelphia, PA. June 1-2. Goswami, B.G., Martindale, J.G. and Scardino, EL. 1977. Textile Yarns, Technology, Structure and Applications. New York John Wiley and Sons, pp. 273-337. Hearle, J.W.S., Grosberg, P. and Backer S. 1969. Structural Mechanics of Fibers, Yarns and Fabrics, Vol 1,New York: Wiley-Interscience. Kaswell, E.R., ed. 1963. Wellington-Sears Handbook of Industrial Textiles. New York: Wellington-Sears. KO, F.K., Bruner, J., Pastore, A. and Scardino, E 1980. Development of Multi-Bar Weft Insertion Warp Knit Fabric for Industrial Applications. ASME Paper No 90-TEXT-7, October. KO, F.K., Krauland, K. and Scardino, F. 1982. Weft Insertion Warp Knit for Hybrid Composites. Proc. 4th Intern. Conf. Composites. KO, F.K., Fang, P. and Pastore, C. 1985. Multilayer Multidirectional Warp Knit Fabrics for Industrial

Applications. J. Industrial Fabrics 4(2). KO, F.K., Pastore, C.M., Yang, J.M. and Chou, T.W. 1986. Structure and Properties of Multidirectional Warp Knit Fabric Reinforced Composites. In Composites '86: Recent Advances in Japan and the United States, eds. Kawata, K., Umekawa, S. and Kobayashi, A. Proceedings, Japan-U.S. CCM-111, Tokyo, pp. 21-28. KO, F.K., Pastore, C.M., Lei, Charles and Whyte, D.W. 1987. A Fabric Geometry Model for 3-D Braid Reinforced Composites. Intern. S A M P E Metals Conference: Competitive Advancements in Metals/ Metals Processing. KO, F.K. 1988. Braiding, Engineering Materials Handbook, Vol 1, Composites, ed. Reinhart, T.J. Metal Park, OH: AMS International, pp. 519-528. KO, F.K., Whyte, D.W. and Pastore, C.M. 1988a. Control of Fiber Architecture for Tough NetShaped Structural Composites. MiCon '86: Optimization of Processing, Properties and Service Performance Through Microstructural Control, ASTM STP 979, eds. Bramfitt, B.L., Benn, R.C., Brinkman, C.R. and Vander Voort, G.F. Philadelphia: ASTh4 pp. 290-298. KO,F.K. and Kutz, J. 198813.Multiaxial Warp Knit for Advanced Composites. Proc. 4th Ann. Con$ Adv. Composites, ASM International, pp. 377-384. KO, F.K. 1989. Preform Fiber Architecture for Ceramic Matrix Preforms. Ceramic Bulletin 68 (2): 401414. KO, F.K. and Du, G.W. 1992. Processing and Structures of Textile Preforms for Composites. Proc. Science and Innovation in Polymer Composites Processing, MIT, Cambridge, MA, July 16-17. Krcma, R. 1971. Manual of Nonwovens. Manchester, UK Textile Trade Press. Li, W., Hammad, H. and El-Shiekh, a. 1990. Structural Analysis of 3-D Braided Preforms for Composites, Part I: The Four-Step Preforms. I. Text. Inst. 81:491-514. Lord, P.R. and Mohamed, M.H. 1973. Weaving: Conversion of Yarn to Fabric. Durham, UK: Merrow Technical Library. Loos, A.C., Weidermann, M.H. and Kranbuchi, D.E. 1991. Processing of Advanced Textile Structural Composites by RTM. Proc. 5th Textile Structural Composites Symp., Drexel University, Philadelphia, PA, Dec. 4-6. McCarthy, S. and Kim, Y.R. 1991. Resin Flow Through Fiber Reinforcement During Composite Processing. Proc. 5th Textile

424 Texfilepreforming Structural Composites Symp., Drexel University, Philadelphia, PA, Dec. 4-6. Mohammed, M.H., Zhang, Z. and Dickinson, L. 1989.3-DWeaving of Net Shapes. Proc. Zst Japan Intern. SAMPE Symp., Nov. 28-Dec. 1. Pastenbaugh, J. 1988. Aerospatiale Technology. Proc. 3rd Textile Structural Composites Symp., Drexel University, Philadelphia, PA, June 1-2. Pastore, C.M. and Cai, Y.J. 1990a. Applications of Computer Aided Geometric Modeling for Textile Structural Composites. Proc. 2nd Intern. Conf. Computer Aided Design in Composite Material Technology, Brussels, Belgium, April 25-27. Pastore, C.M. and KO, F.K. 1990b. Modeling of Textile Structural Composites, Part I: Processing-Science Model for ThreeDimensional Braiding. J. Text. Inst. 81: 480-490. Popper, P. and McConnell, R. 1987. A New 3-D Braid for Integrated Parts Manufacturing and Improved Delamination Resistance - The 2-Step Method. 32nd Intern. SAMPE Symp. Exhib., pp. 92-103. Potter, K.D. 1979. The Influence of Accurate Stretch Data for Reinforcements on the Production of Complex Structural Mouldings. Composites, 10, pp. 161-167, IPC Business Press Ltd, July.

Raz, S. 1987. Warp Knitting Production. Heidelberg, Germany: Melliand. Scardino, F.L. 1989. Introduction to Textile Structures. In Textile Structural Composites, eds. Chou, T.W. and KO, F.K. Amsterdam: Elsevier, pp. 1-26. Scardino, EL. and KO, EK. 1981. Triaxial Woven Fabrics. Textile Research Journal 51(2). Smith, B.F. and Block, I. 1982. Textile In Perspective. Englewood Cliff, New Jersey: Prentice-Hall. Spencer, D.J. 1983. Knitting Technology. New York: Pergamon Press. Stover, E.R., Mark, W.C., Marfowitz, I. and Mueller, W. 1971. Preparation of an OmniweaveReinforced Carbon-Carbon Cylinder as a Candidate for Evaluation in the Advanced Heat Shield Screening Program. AFML TR-70-283, Mar. Svedova, J., ed. 1990. Industrial Textiles.Amsterdam: Elsevier. Tampol’skii, Y., Zhigun, I.G. and Polikov, B.a. 1987. Spatially Reinforced Composites. Pennsylvania: Teknomic. (English translation, 1992). Williams, D.J. 1978. New knitting methods offer continuous structures. Advance Composites Engineering, Summer, pp. 12-13.

TABLE ROLLING OF COMPOSITE TUBES

19

John T. K a m e and Jerome S. Berg

19.1 INTRODUCTION

In the field of composites fabrication table rolling is a major technique for utilizing preimpregnated fibrous tapes in flag or pennant form for tubular structures. The individual flags become part of the total wall thickness by rolling the flags around a mandrel. The hard mandrel provides the support during cure and defines the inside dimensions of the tube. Table rolling is utilized to fabricate a variety of products including straight tubes usually under 7.62cm (3in) in diameter and up to 3.66 m (12 ft) long and small diameter tapered tubes such as fishing rods, golf shafts and ski poles. Flags may consist of a wide variety of fibers oriented either longitudinally (along the axis of the tube) or offset at a bias angle, hence the terms ’longitudinal and bias flags’. The resin content (RC) and the fiber areal weight (FAW) of the prepreg define the ply thickness. Since external molds are seldom used for table rolled tubes, a variety of polymer compaction tapes are used to apply an external pressure. These tapes provide the external pressure necessary to debulk and prevent flag unravelling before cure and to provide some heat driven compaction during cure. 19.1.1 FIBERS AND RESIN

Carbon fiber form 234 M P (34 ~ msi) to 620 M P (90 msi), glass ’E’ or 8s’ aramids, polyethylene and boron are some of the common fibers

we,


in table rolled tube manufacture. The most common resin used to coat the fibers is the epoxy blend family, which is formulated for specific product purposes. The resin and fiber are combined and advanced slightly to a selected ’tack‘ (stickiness) level. This fiber and resin combination is called ’prepreg’. Prepreg surface tack has an important adhesive quality in table rolling which permits composite flags to adhere to one another or the mandrel without slipping during the table rolling operation. Prepreg is offered by specification of resin type and roll width. FAW from 130-160 g/m2 and RC from 30-36% are common in table rolling. Higher modulus fibers favor a lighter FAW to ease rolling. The epoxy in prepreg is catalyzed, so care must be exercised in following the prepreg vendor’s storage and handling recommendations. Freezer storage can extend the shelf life; therefore, the material generally arrives on a freezer truck. Because air and moisture are detrimental to the surface tack, it is important to cut and consume the flags as soon as possible after opening and unrolling the prepreg. Dry and low tack prepreg can influence and aggravate flag wrinkles and ply slippage, leading to voids and dimensional problems. Address tack with the prepreg supplier to find a suitable resin formulation for the table ~ roller’s manufacturing environment. equipment and A Partial list of tooling suppliers commonly utilized in table rolling is presented in Table 19.1 at the end of this chapter.

426 Table rolling of composite tubes 19.1.2 DESIGN

Figure 19.1 illustrates the relationship of the mandrel, prepreg flags and diameters in table rolling. 19.2 EQUIPMENT

19.2.1 SHEAR/SHEETER

A number of commercially available power shears have hardened steel blades and include automatic feed mechanisms for the material rolls. Safety guards with interlocks are needed to prevent finger and hand injury (Fig. 19.2).

Fig. 19.2 Photograph of prepreg sheeter. (Courtesy of Century Design Incorporated.)

The tooling used in this machine is the 'Steel Rule Die'. 19.2.2 ROLLER PRESS The steel rule die, the prepreg and an impact The roller press is a machine used to press sev- sheet (usually a soft plastic like polyethylene) eral stacked layers of prepreg tape into are passed through the rotating pressure individual patterns such as tapered pennants. wheels of the roller press. After compaction,

' Fig. 19.1 Diagram of tapered mandrel, bias flags and longitudinal flags.

-p

Of

longIt@Ml

ma,

Equipment 427 the impact sheet is removed to expose a stack of ready-to-assemble flags which are now nested between the blades of the die. The roller press and dies act similarly to a kitchen cookie cutter (Fig. 19.3). A sharp knife and straight edge can act as a prototype or for small scale production.

---'-.-'

---

,.-.. ._."

I

-.

.-

Fig. 19.4 Photograph of table rolling press. (Courtesy of Century Design Incorporated.)

Fig. 19.3 Photograph of a roller press including loading and unloading racks. (Courtesy of Century Design Incorporated.)

19.2.3 ROLLING TABLE

Originally, in the 1940s, prepreg flags were hand cut and hand rolled like cigars to produce the tubular structure. The fishing rod industry was probably first to commonly use rolled tubular structures. In the 1960s several devices with a mobile lower platen were developed. After activating the machine, a pivoted upper platen is lowered down upon the mandrel and a linear motion activator in the lower platen rolls the mandrel into the prepreg flag. These machines permit pressure ranges to be established, gaining maximum compaction and increasing the speed of rolling. The pivoted upper platen permits the combinations of parallel tube or tapered (cone-like)tube rolling. If the mandrel is parallel, then the pivot function of the upper platen will be unnecessary. Figure 19.4 illustrates a commercially available rolling table. Current rolling tables include temperature controlled platens and platens with piano key-

like fingers for achieving uniform pressure on tapered parts. Both flat bed and segmented bed versions are covered with canvas. This pad provides sufficient resiliency and friction to permit flags to roll without slipping, yet conform to the mandrel surfaces. A slight dusting of talcum powder can be used to prevent prepreg from sticking. Table rolling provides tighter and a more uniform compaction of plies than hand rolling. 19.2.4 VERTICAL TAPE WRAPPER

A variety of plastic and/or cellophane tapes, 1.27-2.54 cm (051.0 in) in width are used to compact the table rolled plies of prepreg. Machines used to apply these tapes must permit tape tensioning to debulk the product as the tape is applied. Some applications call for multiple passes through the tape wrapper to increase the tape pressure for better compaction. Additional wraps of tape are needed for thicker wall structures. Frequently, two types of tape may be used: a release tape and a secondary compaction tape. Some tapes have a release backing which can allow a single pass of tape. However, these tapes are generally more expensive. Apply the tape as soon as possible after table rolling to prevent the flags from loosening. Figure 19.5 illustrates a vertical tape wrapping machine.

428 Table rolling of composite tubes

Fig. 19.6 Photograph of horizontal tape wrapping machine. (Courtesy of Century Design Incorporated.) 19.2.6 MANDRELPULLER

Fig. 19.5 Photograph of vertical tape wrapping machine. (Courtesy of Century Design

Incorporated.)

19.2.5 HORIZONTAL TAPE WRAPPER

This machine represents an alternate to the vertical tape wrapper. It is used frequently for longer, heavier parts and also for very flexible mandrels such as fishing rods. The mandrel is affixed to a chuck or mechanical coupling which rotates the parts while tape is applied. The rollers provide support for the part while motion is in place. The single or even dual tape feed spools move with the tape carriage and return to restart position (Fig. 19.6). After cure, the wrapping tapes are removed by slitting the tape longitudinally and peeling the tape away from the cured part. Wrapping tapes are then discarded.

Tubular parts which have been cured over a hard mandrel are all subject to mandrel extraction. The mandrel puller generally connects to a bolt on the larger shank end of the mandrel. The end of the composite tube rests against a stationary block shaped to permit passage of the mandrel but blocking the tubular part. Mandrel pullers are generally hydraulic or pneumatic. Hydraulic pullers offer a controlled extraction speed, while pneumatic pullers are faster and useful in high volume environments. Mandrel withdrawal is generally done prior to tape removal. Figure 19.7 shows a pneumatic mandrel puller. The type of mold release used, correct size of the stationary block and the mass plus integrity of the cured part must be carefully evaluated or end crushing of the part will occur. Also, thin walled tubes or tubes with a high degree of longitudinal plies can crack during mandrel extraction. 19.2.7 CURING OVENS

Ovens used for curing the composite tubes can be either electrically or gas heated and of batch or conveyorized design. Temperatures ranging from 121-191°C (250-375°F) are most common for roll forming prepregs. Consult

Materials 429 19.3 TOOLING 19.3.1 MANDRELS

Fig. 19.7 Photograph of pneumatic mandrel puller. (Courtesy of Century Design Incorporated.)

the prepreg supplier for recommendations on appropriate cure profiles. Fine tuning of the cure profile is often needed to optimize particular roll forming operations and specific products. Ovens with thermocouples are useful in determining hot and cold spots, which may indicate oven regulation for uniform temperature control. This ensures a uniform gel within the part. 19.2.8 CENTERLESS SANDER OR GRINDER

The wrapping tapes can leave a series of spiral indentations approximately 0.5 mm (0.002 in) deep in the composite tube surface. If a smooth surface is desired for cosmetic reasons or for geometry requirements the part can be surface sanded or ground. A centerless sander basically removes a user defined controlled amount of surface material. A centerless grinder provides a more accurate finish dimension. Centerless grinders are common for the high precision required for the tip ends of golf shafts in which a tolerance of f 0.5 mm (a.002 in) is not uncommon. In carbon fiber golf shaft manufacture, these surface finishing techniques are also used to tailor the product stiffness by incrementally removing material along the shaft length. This changes the shaft stiffness characteristics.

The mandrels used for table rolling are usually hardened steel, sometimes aluminum or even composite. The mandrels are designed to support the prepreg during rolling and curing and provide the inside dimensions for the part. Recalling that mandrels must be extracted in 'mandrel pulling', some negative taper is beneficial. Mandrel makers are skilled in the art of metallurgy. They can select the materials and heat treatments necessary to create a mandrel resistent to permanent bending. However, in many instances the mandrels can be restraightened if damaged. Hard plated mandrels generally provide a longer life since scratched or dented surfaces will hamper mandrel removal. 19.3.2 STEEL RULE DIES

These dies incorporate multiple blades embedded in a rigid backing (usually marine grade plywood) which cut the material in the roller press. The prepreg tape (up to 20 layers) is cut between the cutting die blades and a polyethylene sheet sandwiched between the rotating press wheels. Dies with one piece blades provide the best and most continuous cuts (Fig. 19.8(a)). Dies which include weldments (as in a triangular shape flag) generally dull faster since the weldments soften the cutting edges (Fig. 19.8(b)) Ramps can be used between the cutting blades in die designs to keep the roller pressure off the blade ends. The die builder can recommend blade height and cutting edge type best suited for the task. 19.4 MATERIALS 19.4.1 MOLD RELEASES

Generally, mold releases for table rolling mandrels consist of two components: a primary mold release which provides a polymer bond

430 Table rolling of composite tubes 19.4.2 FIBERS

a

L\ A

Rubber Dads (Stacking reference)

C blades utting\-

U

The prepreg tapes can be made from longitudinal tows or woven tows of the following fibers: aramid, glass, carbon and boron. All these fibrous composite tapes can be cut into flags and pennants needed for the table rolling process. However, because of the brittle nature of boron prepreg, boron is most often cut for longitudinal flags. Refer to the appropriate chapter for specific properties of these fibers.

A

19.5 TYPICAL PROBLEMS

19.5.1 VOIDS

Cutting blades Ramps

Fig. 19.8 (a) Diagram of steel rule die for rectangular shaped flags. Ramps prevent blade damage. @) Diagram of steel rule die for a triangular shaped flags. Ramps prevent blade damage. Weldments can cause blade dulling.

Voids are caused by entrapped air which is not evacuated before resin gelation. The presence of voids reduces the strength bearing capabilities of the part, creates stress risers and can contribute to surface finishing and cosmetic problems. Voids are first minimized by working with the prepreg supplier to assure a high quality material with uniform resin content and good 'wet-out' of the fibers. Also, the material suppliers (prepreg and wrapping tapes) must play a key role in developing a cure profile for the specific process and products. Voids are increased by flag wrinkles which are indicative of rolling problems. The capability to perform void content checks (ASTM D3171) and photomicrographs of the laminate can be extremely useful to develop and improve tube processing. Laminate photos are also very useful in operator training. Few laminates are completely void free but void contents lower than 1%are possible with table rolling and tape wrap compaction.

to the mandrel surface to prevent adhesion; and a secondary mold release which acts as a slip agent. The secondary mold release is most beneficial in straight or slightly tapered mandrels and is reapplied between subsequent mandrel turns. A primary release can lose its effectiveness after several hundred turns and must be 19.5.2 DRY AND DIFFICULT TO ROLL stripped off and recoated. A variety of quality MATERIAL mold releases are on the market. The fabricator should work with the release supplier to Prepreg dryness (lack of tack) can be due to low develop a coating program for the specific resin content, resin formulation, ambient condiapplication. Silicone based releases should be tions of the manufacturing environment, the avoided if the tube is subjected to subsequent age or out time of the material. Insufficient tack can cause flag movement during assembly, bonding and painting.

Typical problems 431 wrinkles, voids and parts with a poor surface finish. Resin content and formulation can be adjusted to suit the manufacturing environment. Temperature and humidity control are very helpful in maintaining consistent material tack in the manufacturing shop. Avoid leaving cut patterns exposed since moisture in the air greatly affects the material surface tack and sometimes renders it useless. Consuming the material within two days is a good rule to follow.

Warm lay-up and rolling tables can help increase material rolling ability and are generally adjusted for slight material and environmental changes. Off angle plies are difficult to roll adjacent to the mandrel and the difficulty is magnified by the higher modulus fibers. Tack tape is a narrow strip of reinforced adhesive designed to aid the adhesion of bias plies to the mandrel. Also, solvent based 'tack resins' can be applied to the mandrel to ease application of the first ply. Once the first ply is tightly rolled, however, the material tack is sufficient for subsequent flags.

Table 19.1 Table rolling equipment, material and tooling suppliers in USA Equipment

Tooling

Century Design Incorporated.

Mandrels Lynco Grinding Corporation 5950 Clara Street Bell Gardens, CA 90201

3635 Afton Road San Diego, CA 92123 (619)-292-1212

(213)-773-2858

Materials

Prepreg Newport Adhesives and Composites 1822 Reynolds Avenue Irvine, CA 92714 (714)-253-5680

Fiberite 4300 Jackson Street Greenville,TX 75403 (903)-457-8554

Mold release Frekote Products Dexter Adhesives and Structure Division One Dexter Drive Seabrook, NH 03874 (603)-474-5541

Wrapping tapes Flexicon Pacific, Inc. 856 North Elm Suite J Orange, CA 92667 (714)-%33-9820

Toray 5729 Lakeview Drive, NE Kirkland, WA 98083-2548 (206)427-9029

Cytec Engineered Materials, Inc. 1440 North Kraemer Boulevard Anaheim, CA 92806 (714)-666-4349

Chemlease P.O. Box 540083 Orlando, FL 32854-0083 (407)-425-2066

Dunstone Company, Inc. 2104 Crown View Drive Charlotte,NC 28227 (704)-841-1380

Steel rule dies Ontario Die Company of America 2735 20th Street Box 610397 Port Huron, MI 48061-0397 (810)-987-5060

432 Table rolling of composite tubes 19.5.3 PART SLIPPAGE DURING CURE

19.5.4 EXPOSED SURFACE VOIDS

The viscosity of the resin drops as the heat of cure begins. Occasionally, tapered mandrels and the constriction of the wrapping tape during the cure can force a part to slip down the mandrel. Golf shaft design is highly dependent on mandrel reference position for proper stiffness and geometry requirements. Slippage can first be minimized by designing a short semi-parallel section in the mandrel (as in the butt section of the golf shaft). Slippage is also reduced by overwrapping the tapes onto the mandrel at both ends to secure the part. In addition, the cure profile or the mold release can be adjusted to limit slippage.

Exposed surface voids after sanding or grinding are indicative of poor rolling practices, insufficient lamination pressure and questionable material. Exposed surface voids are sometimes referred to as 'fiber pulls', which have a wood grain appearance on parts with longitudinal surface plies. 19.5.6 LONGITUDINAL PLY WAVINESS

Tapered parts with longitudinally oriented fibers are prone to zones with a wavy or 'fiber wash' appearance. The problem is amplified with multiple taper mandrels and very low viscosity prepregs. Cure profile modifications or alternate resins can reduce the tendency of 'fiber wash'.

RESIN TRANSFER MOLDING

20

Lihwa Fong and S.G. Advani

20.1 INTRODUCTION

ready for its removal from the mold when sufResin Transfer Molding (RTM) is a closed ficient green strength is attained. Processes mold process in which matched male and that are based on similar principles include female molds, preplaced with fiber preform, Structural Reaction Injection Molding (SRIM) are clamped to form composite components. and different versions of vacuum assisted Resin mix is transferred into the cavity RTM (Figs. 20.1 and 20.2). RTM offers the promise of producing low through injection ports at a relatively low prescost composite parts with complex structures sure. Injection pressure is normally less than and large near net shapes. Relatively fast cycle 690 kPa (or 1OOpsi). The displaced air is times with good surface definition and allowed to escape through vents to avoid dry appearance are easily achievable. The ability spots. Cure cycle is dependent on part thickto consolidate parts allows the saving of conness, type of resin system and the temperature siderable amount of time over conventional of the mold and resin system. The part cures in lay-up processes. Since RTM is not limited by the mold, normally heated by controller, and is the size of the autoclave or by pressure, new Mixing Head

r -------1 I 1

I

I I I I

I

I

I

I

I I

I

Pumpunit

---------

Fig. 20.1 Schematic of the RTM process. Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

Vent Port

-

434 Resin transfer molding ISOCYANATE

POLYOL

HYDRAULICS

MACHME MODE

HEAD

Fig. 20.2 Basic construction of a SRIM machine.

tooling approaches can be utilized to fabricate Advantages are: large, complicated structures. However, the 0 Class A surface: Surface definition is supedevelopment of the RTM process has not fulrior to lay-up. In addition, using matched filled its full potential. For example, the RTM tools for the mold, one can improve the finprocess is yet to be automated in operations ish of all the surfaces. such as preforming, reinforcement loading, 0 Close tolerance: Parts can be made with betdemolding, and trimming. Therefore, RTM ter reproducibility than with layup. can be considered an intermediate volume 0 Design tailorability: Reinforcement and molding process (Krolewski, 1990). combination of reinforcements can be used Several unresolved issues in RTM encounto meet specific properties. tered by composite engineers are in the areas 0 Fast cycles: Production cycles are much of process automation, preforming, tooling, faster than with layup. mold flow analysis and resin chemistry. 0 Filler: Filler systems can be used to reduce During the last decade, rapid advances in cost, improve fire/smoke performance, surRTM technology development have demonface appearance, and crack resistance. strated the potential of the RTM process for 0 Gel coat: One or both mold surfaces can be producing advanced composite parts. The gel-coated to improve surface performance. advantages and associated disadvantages of 0 Good mechanical properties: Mechanical the RTM process are summarized. As the properties of molded parts are comparable development of this process is rapid, some of to other composite fabrication processes. the disadvantages may be overcome by the 0 Good moldability: Large and complex advances made in this technology. shapes can be made efficiently and inexpen-

RTM process 435

0 0

0

sively. In addition, many mold materials can be used. Inserts: Ribs, bosses, cores, inserts and special reinforcement can be added easily. Labor saving: The skill level of operator is less critical. Low tooling cost: Clamping pressure is low compared to other closed mold operations. Low volatile emission: Volatile emissions are low because RTM is a closed mold process. The worker is not exposed to chemical vapors as with the lay-up process.

Disadvantages are: Mold design: The mold design is critical and requires good tools or great skill. Improper gating or venting may result in defects. Mold filling: Control of flow pattern or resin uniformity is difficult. Radii and edges tend to be resin-rich. Properties are equivalent to those with matched-die molding (assuming proper fiber wetout, etc.), but are not generally as good as with vacuum bagging, filament winding or pultrusion. Reinforcement movement during resin injection is sometimes a problem. In the following sections, the resin transfer molding process is discussed in terms of the unit operations involved, to familiarize readers with the basic steps of the RTM process. The discussion covers details such as materials of construction, mold design, preforming, curing, and demolding. Processing issues are mentioned in each individual unit operation. Relevant variations of RTM such as vacuum assisted resin transfer molding and flexible molding tools are summarized. Process physics is described with emphasis placed on the principles that govern the RTM process; these are applied in the use of computer simulations. Through the design tools such as simulation codes for mold filling analysis, engineers are able to predict or diagnose the problems in gating and venting in the

design stage. The usefulness of such design tools is discussed in detail, giving the relevant advantages and disadvantages. 20.2 RTM PROCESS

The RTM process can be viewed as seven unit operations. The general practice and processing issues are described for each unit operation. 20.2.1 FIBER REINFORCEMENT

Selection of the proper reinforcement type should take into consideration loading condition, part geometry (size, thickness), mechanical properties and surface finish. The quantity of parts demanded also determines the selection. The reinforcement normally carries 90% of the load in a composite and provides over 90% of the stiffness. The reinforcement in a composite can be designed to match the strength requirements of the part. The following characteristics should be considered when selecting fiber reinforcements: Volume fraction: ratio of the volume of a given mass of reinforcement to the volume of the same component after molding; Wash resistance: ability of a reinforcement to withstand movement due to fluid motion or solvation of the reinforcement binder by the resin; Wettability: ability of a reinforcement to reach a condition wherein all voids in the reinforcement are filled with resin; Sizing: most fibers are coated with size for better wettability and bonding but the size may influence the cure kinetics during the manufacturing. Most standard reinforcement materials for composites can be used, but fiberglass, carbon and aramid are the most common in RTM. One requirement is that the reinforcement should hold its shape during the injection

436 Resin transfer molding

phase. Therefore, the reinforcements are generally stitched, woven or bonded together. Reinforcement build-ups in certain areas can easily be included. For example, woven roving and fabric can be combined with continuous strand mat and chopped strand mat in applications where higher strengths are required. Hybrid systems composed of high performance reinforcements such as carbon fiber and aramid fiber can also be incorporated in RTM laminates. Surfacing materials called veils can be used in the preforms to hide the imprint of fibers, for improved surface finish. Another application of surfacing veil is to achieve a resin-rich skin to improve corrosion resistance. Stitched fabrics (Fig. 20.3(a))reduce stresses inherent in the woven roving design and lead to higher compressive strengths in the composite. However, other constructions such as 8-HS style of weave (eight-harness satin weave) in Fig. 20.3(b)have been used because of improved wetting characteristics and compressive strength compared to bidirectional woven fabrics. Continuous strand mat is multi-stranded, laid in swirled configuration. The mats nor-

mally have 4-6 wt.% of thermoplastic binder added. They are thermoformable and can therefore be used for highly complex shapes or when the anticipated volume of production makes them economical. Different sizings can be obtained on many reinforcements. Sizings can be tailored to the type of resin system. Sizings are available that are compatible with epoxy, vinyl esters or polyesters. The strength variation with type of sizing can be as much as 20%, so this factor needs to be considered in the choice of reinforcement. 20.2.2 PREFORM

For a flat part, the preform can be as simple as a stack of reinforcements that fit in the mold cavity. As preforms become more versatile, various means of producing preforms are available. Currently cut-and-sew is commonly used to assemble preforms of various shapes for aerospace applications. Other near net shape techniques include braiding, spray-up and thermoforming (Fig. 20.4).

I

I

Fig. 20.3 (a) Stitched fabric; (b) eight-hamess satin weave.

I t

1 1 1 1 1 1 1 1 1

1 1 1 1 1 1 1

I

RTM process 437

blank preparation zone

heating zone

stamping zone

unloadinghrimming zone

heatingcuring resinbinder

1

1 cooling A

t+

Vayum source

I

demolding

1

N

directed fiber preform

Fig. 20.4 (a) Four station thermoforming preformer; @) spray-up process.

M

'

forming screen


If prefabricated preforms are not used, then some means must be found to hold the layers of reinforcement together as they are built up on the tool surface. For example, unidirectional reinforcement is subjected to washing (washing is unplanned reinforcement movement due to resin movement) if proper precautions are not taken to prevent it. To improve conformance of fibers, a tacky resin (e.g. epoxy), dissolved in suitable solvent (e.g. acetone), can be used as a spot glue to hold the reinforcement layers together. The tacky resin will be washed out during the resin injection cycle and will not interfere with the cure of the part. Sometimes veil can be used to hold the layers and prevent washing.

Edge definition: The edges of the composite will be resin rich if the preform is not cut to fit closely to the edge of the cavity or inserts; Fiber distribution:Uniformity of fiber content in preforms, without excessive thinning, wrinkles or folds, is important; Permeability:A measure of resin distribution into the cavity. This quantity is also affected by fiber volume fraction.

Prefabricated preforms can be further bonded together, with or without a core, to achieve part consolidation. For structural composites, this eliminates the need for fasteners and adhesives to assemble discrete parts. New thermoformable reinforcement mats can be Advantages are: used for highly complex shapes or when the 0 Fast loading: Preforms allow fast loading of anticipated volume of production makes them the mold. economical (Carvalho, 1991). 0 Precise fiber placement: Preform placement Design of preforms should go hand in hand can be made precisely without misalignwith part design. For example, preform corners ment. This allows high quality, close are sensitive to radii of the shape. Figure 20.5 tolerance composites for advanced applicashows the thickness reduction of preform over tions molded by the RTM process. different radii. The preform thickness does not 0 Net shape preforms: If thermoformable change appreciably compared to those around reinforcements are used, the stamped prethe corner when the radius is made larger than forms have excellent dimensional stability. a critical value. However, if the radius is less 0 No additional tool: For low production volume of the composite, the tool for preforming can be the same as the tool for molding.

The only disadvantage with use of preforms is that there is an additional unit operation. With the obvious advantages, use of preforms is advisable when volume of production allows their economical use. When designing fiber preforms, following issues should be considered: Corners: The fiber in the bent corner of a preform tends to move to the inside of a radius. This can cause channeling of flow that leads to poor mold filling patterns; Drapability:This characterizes the ability of a fabric or reinforcement mat to conform to contours of the tool;

1 .o

WRm Fig. 20.5 Effect of comer radii on preform thickness (Rmis the recommended radius).

XTM process 439 than this critical value, dramatic movement of the fibers to the inside of the radius occurs. As a result, channeling becomes dominant in the mold filling stage and induces irregular flow patterns. The edge of a preform is another source of the race tracking of resin. In order to avoid the channeling effect, the preform edge should be cut to fit the edge of the mold cavity. The task of obtaining a good edge definition is normally difficult because of the bulkiness of the layers and inter-layer movement (sliding, rotating) during the mold closing when prefabricated preforms are not used. Preforming of fabrics over tool geometry other than simple flat type will induce shear deformation in the fiber reinforcements. For a biaxial reinforcement, shearing of the weave

(Fig. 20.6) is necessary to conform to the contours of the tool. This drapability problem, therefore, has a two-fold significance in RTM. Because of the fiber rearrangement, the nonuniformity of fiber distribution should be accounted for in the design of the composite. Fiber volume fraction and orientation are no longer that of the unreformed reinforcement. Further, such preforms exhibit different characteristics to resin flow. Designers should account for this change in determining the location of vent ports relative to an injection port. In practice, to modify the permeability of preforms, various flow inducing media or mechanisms have been suggested. Application of such high porosity to the preform or inclusion of a runner system in mold design can alter the mold filling pattern.

Fig. 20.6 Draping fabric on a mold causes shearing of the weave.

440 Resin transfer molding Low viscosity: High viscosities can cause mold pressures that are too high in both the The resin used in the RTM process forms the mold and the injection unit. Raising the temmatrix in the composite after solidification. perature of the resin system is effective to The solid structure is a result from polymerlower its viscosity, but pot life may be ization. To select the resin system, one must adversely affected. take into account of the rheological change Sufficient pot life: This is the time it takes and resin cure kinetics. The formulation of the resin system the resin system's viscosity to reach a level depends on many factors. For example, the that no longer be comfortably handled by the resin system can be combined with promoters, equipment. fillers, internal mold releases, pigments, etc. Tg point: The glass transition temperature Typical fillers, such as clay or calcium carbonshould be as high as possible. As a rule of ate, may reduce cost. The optimum viscosity thumb, the glass transition temperature should for RTM should be less than 500 cP s. Mixing is be at least 30°C (50"F), and preferably 55°C normally required to form a suspension. (100"F),higher than the service temperature Properties requirements (mechanical, chemical, fire retardancy, etc.) can also affect Toughness: Toughness in a resin system is resin selection; the resin mix can be formu- exhibited by its tensile elongation. If sufficient lated to meet specific needs. Attributes to look damage tolerance is required, the elongation for in resin systems are: should be at least 3%. 20.2.3 RESIN SYSTEM AND INJECTION

e consistent reactivity;

ability to wet out the reinforcement; e rapid cure after gel.

0

The ester-type resin mix is combined with an appropriate catalyst, such as emulsified BPO, MEKP, cumene hydroperoxide, at the mixing head and transferred into the RTM mold. Low profile additives have been developed especially for polyester resins to improve surface appearance. In addition, epoxies, urethanes, vinyl esters, nylon and other hybrid resins are available for RTM. The newer resins may require modifications to the pumping/injection unit to meter and condition the resin mix prior to injection. These new systems offer a range of cost and performance options for the RTM process. Influencing parameters are viscosity, pot life, tensile modulus, glass transition temperature, tensile elongation and moisture absorbance. In considering a new resin system, the choice of the proper resin system for RTM must satisfy the following system criteria. Failure to meet these criteria usually means that the resin system is impractical for RTM.

Young's modulus: This modulus must be over some threshold value or the composite compression strength will be less than the optimum value. A high tensile modulus is required to adequately support the fiber reinforcement and prevent premature buckling. The effect of the resin system on hot-humid performance is important in the composite part. The modulus of a typical resin remains essentially constant until the temperature is close to the ultimate T , when it falls off to zero. Under wet conditions, the strength of the resin usually falls off at the same rate as the modulus because of the effect of absorbed moisture. Absorbed moisture plasticizes the resin matrix and lowers the strength of the composite in non-fiber dominated directions. The amount of moisture absorbed by the resin matrix should be small, normally less than 2%. This limits the amount of mechanical performance degradation at elevated temperatures. One final topic to consider is the injection of the resin system (schematics shown in Figs.

Mold materials 441 20.1 and 20.2). Items to control in the resin mix to assure a consistent, smooth running process include: 0 0 0 0

resin mix temperature; ratio of catalyst or curatiire to resin; resin mix viscosity; amount of air entrained in the resin mix. Presence of air in the mix can lengthen the gel time/induce porosity in the composite and/or affect the mix viscosity.

(a)

Most successful production resin transfer molding operations are now based on the use of resin/catalyst mixing machinery using positive displacement piston-type pumping equipment for accurate control of the resin to catalyst ratio. Back pressure at the mix head may change when a mixed resin is injected into a cavity filled with the fiber reinforcement. Static mixers greatly simplify the process and are easily cleaned at the end of the injection cycle. A static mixer sends the proportioned resin and catalyst through flexible Fig. 20.7 (a) Matched mold with rigid halves; hoses to an injection head employing a (b) matched mold with a flexible mold half. motionless mixer to thoroughly blend the materials together immediately prior to injecGate and vent: This critical part of the mold tion step. design should allow complete wetout with minimal resin wastage. 20.2.4 MOLD

RTM mold design and construction is the most critical factor in successful resin transfer molding. The mold must be constructed so that resin reaches all areas. RTM molds require special considerations compared to other composite tooling. Figure 20.7 shows two possible configurations in RTM processing. The mold must be designed to account for the following factors:

Mold sealing: A perimeter gasket is necessary to keep void content low. Tight sealing is important when vacuum is used. Heatingkooling: A typical RTM cycle consists of a wide range of temperature for initiating the chemical reaction, curing and final demolding. Hence, proper heating/cooling channels need to be designed.

Mold materials: The material of construction 20.3 MOLD MATERIALS dictates life cycle of mold, temperature control The low pressure requirements of RTM allow and press requirement. the use of more types of mold materials than can be used in other composites manufacturCavity design: The RTh4 mold should consoliing. The choice between metal molds and date as many assembly steps as possible. A good polymeric composite molds is chiefly one of design should take advantage of this ability. volume and processing temperatures. High


volume and high temperatures dictate metal molds. Steel, the most suitable mold material, provides superior face life. Aluminum is good for construction of prototype molds since the metal is easy to machine, is lightweight and has a reasonably high heat transfer rate, but also galls easily. Cast aluminum and spraymetal tooling are currently available and can be used for higher volume applications. Cast copper alloys are being considered for use in RTM molds due to the potential for increased throughput via heat management and better durability. Composites, for example reinforced polyester and epoxies, are most frequently used for making RTM molds. They can be expected to last for approximately 2000 parts (Isorca, 1992).Higher production volumes may justify the use of higher cost spray-metal or metal tools. In some cases, the mold must be backed up in order to maintain its shape. Conventionally the backup can be done cost-effectively with core material or steel frames to add rigidity to the cross section and to support composite mold faces. The closure of the mold is achieved by mating of the mold surfaces against a perimeter gasket. Therefore, guide pins are usually employed to align the mold halves both laterally and vertically to keep resin from leaking. Advancement in adapting composite tooling to the needs of RTM is underway. For example, lengthening the life of the composite tool face is desirable and effective to maintain quality while keeping costs low. The factors that cause deterioration of the mold face are temperature fatigue and attack by solvents or mold release agents. An electrolytically or vapor deposited nickel shell is a new technique that will extend face life.

steps. Therefore, the mold designer should incorporate this rule in the design of the mold cavity. Instead of joining several substructures or onto a major structure after molding, it is structurally more effective and efficient to incorporate them into the part before fabrication. This can be easily achieved by joining substructure preforms when practical. In production the number of molds or cavities required is determined by needed throughput. This should take into account the cycle time. For small parts, the designer can incorporate several cavities in a mold. High surface quality with excellent dimensional control can be achieved by electroplating the mold face with nickel. The appearance surface of a part is usually placed on the bottom of the mold. Pinholes are more likely to collect on the top surface. Mold preparation is similar to that used for hand lay-up. Anew tool must to be treated with several coats of release agent. New mold materials provide flexibility in mold design for RTM. For example, to demold a part with vertical sides, it is common to allow several percent draft in the vertical dimension. Flexible silicone rubber has been used for RTM molds in the form of a bladder mold half which is capable of being inflated or deflated depending on the process requirement. During mold filling, the flexible mold wall is pressed against the rigid wall by inflating the bladder with a pre-determined pressure. During the injection cycle, the mold can deform to enhance resin flow. Upon completion of mold filling, the flexible tool can be further inflated to consolidate the composite component. Part removal in this case is easy since the flexible half can be deflated. %s technique allows fabrication of complicated parts that are not ordinarily possible to demold.

20.4 MOLD CAVITY DESIGN

20.5 INJECTIONPORT AND VENT DESIGN

One of the most important design rules for RTM parts is to reduce the number of assembly

The injection port allows the resin to be transferred into the mold (Fig. 20.1) and its design

Heating and cooling design 443 may be critical. The location of inlet ports must allow the resin to reach all areas without bypassing part of the reinforcement. Air vents help control internal pressure, bleed out air and provide a visual indication of mold filling. Race tracking, or channeling, in the mold is usually the reason why the resin bypasses areas of the reinforcement. Since the resin will not flow backwards, this tends to create dry patches. The engineering way to ensure complete initial wetout is to gate the mold correctly in the design. This may be difficult even for an experienced mold designer. Use of computer simulations as a design tool has become popular in conventional injection molding. Without an engineering design tool, gates and vents can be put in the mold after molding some trial parts, but many trial runs may be prohibitive in some applications. In the next section, new engineering tools adapted for RTM mold filling will be discussed to overcome the problem. Mold designers have found that RTM molds must be vented to allow the air within the mold to be pushed out by the resin. Gate at the lowest point and vent at the highest point is generally a good design practice. Experienced designers may use symmetry to design the inlet ports and outlet vents to remove entrapped air. Venting ports must be placed to draw the resin through sections of the part that are difficult to wet out. They are best placed at dead ends where the resin would not flow by itself. After the resin has finished bleeding, both injection and venting ports must be sealed off. This allows pressure to build up in the mold, and forces the resin to further wetout other sections of the part. This packing stage allows the part to gel under pressure, decreasing void content in the finished part. 20.6 SEALING THE MOLD

The perimeter gasket seals the edges of the mold to prevent loss of resin and injection pressure. In addition, it is an absolute necessity

when vacuum is used. Sealing the mold to achieve cavity pressure of 690 kPa (100 psi) or higher is necessary if the void content of the part is to be kept low. The only practical way to accomplish this is to use O-rings. Machining the face of the mold to close tolerances is prohibitively expensive. It is also usually impossible to maintain the mold absolutely flat to achieve a metal-tight seal. O-ring design is well established. The slot has to be cut so that the O-ring can deform when the mold is closed and maintain a seal. Either square or round O-ring grooves can be used. The type of O-ring material used depends on the maximum temperature the 0ring will experience during the fabrication cycle and the type of solvent used to clean the mold. Nitrile rubber material can be used satisfactorily up to 120°C (250°F). Over 120°C silicone rubber can be used to temperature approaching 177°C (350°F). If help is needed in sealing around inlet or outlet tubes, tacky sealant can be used. This type of sealant is useful for making an O-ring where grooves do not exist. 20.7 HEATING AND COOLING DESIGN

The mold should have good temperature control. The RTM mold should be able to heat and cool the part during the fabrication cycle. Most resin systems cure faster at elevated temperatures. During demolding, lowering the temperature is sometimes helpful in removing the part. Even molds that are intended for room temperature-cured resins should be well insulated so that environmental conditions do not change the gel times and viscosity of the resin. Some molds are heated or designed to go into ovens to achieve faster cures at higher temperatures. Normally, the mold is heated and cooled using either hot water or oil. The mold is constructed to allow the heating/cooling fluid to flow through channels (Fig. 20.8) in its interior. The fluid is heated and cooled by conventional means, such as a gas-fired heater and heat


heating channel

-

////////////// TCO

topmold platen resin flow

hh

V/HB la

=

oT -oOoOo Ko0o 0o 0

Fig. 20.8 Heating/cooling by flow channels in the RTM mold.

exchanger. For larger molds, the heating and cooling times will be longer if the heat transfer area does not increase in proportion to the weight. At some point, the production cycle time becomes limited by the rate at which heat can be added or removed, and becomes independent of the curing characteristics of the resin system. Under development is low thermal inertia technology that allows the tool face to be heated by electric wires buried in the face. The construction of the mold face is such that the heat flows into the mold face and not outward toward the mold support structure. This is accomplished by use of a foam core that insulates the bulk of the mold from the tool face. This novel technology, if successful, will allow a more instantaneous transfer of heat where it will do the most good - at the mold face. 20.8 MOLD FILLING

Resin injection is to pump the base resin system to a mixing head through either a single or two pot system. Impingement mixing of the components occurs in the mixing head. The catalyzed mix is then pumped through a static mixer which completes the mixing of the two components. The injection nozzle is attached to the injection port on the mold and the resin system is injected into the mold to pack the mold to a predetermined pressure. When the

mold is filled, the pumping system is shut off and immediately flushed, and the part is allowed to cure. Successful configurations demonstrated in the industry show a common factor: that is, the flow of resin is symmetrical about the vent ports, in a manner such that the volume of air left in the reinforcement decreases. This compression effect helps sweep the remaining air out of the part. When the flow path is arranged in such a way that the resin flows into a configuration with increasing volume, there is a tendency to bypass part of the reinforcement. This situation can happen when core material is used. For example, when there is reinforcement on either side of a core, it is possible that slight misalignment in the core thickness will cause dry spots in the part. To overcome this problem, the resin must be introduced on either side of the core simultaneously. Holes may be drilled through the core to allow the resin system to flow to the other side. When this is done, the core floats on the wet reinforcement and equalizes itself. When the injection pressure is too high or reinforcement tends to move in the mold, the following remedies must be considered:

Multiple gates: partition the mold along the flow path such that travel distance for resin is reduced.

Curing 445 Runner system: allows the delivery of resin to various parts of the reinforcement quickly without using high injection pressure. Flexible mold wall: allows the deformation of the bladder wall to facilitate mold filling. There are several techniques to modify the flow patterns. Application of high porosity media on the preform or inclusion of a runner system in mold design can alter the mold filling pattern. This is helpful in reducing injection pressure or displacing air. All resin movement must be accomplished within the time allowed before the onset of gelation. Additionally, the resin injection process should not cause movement of the reinforcement and should be done at low pressure so that the mold will maintain its shape without requiring massive backing. Vacuum may be used to facilitate filling the mold and simultaneously assist in removing air from the laminate. This requires good mold sealing and the use of a vacuum pump. Vacuum up to 740-760 mm Hg (29-30 in Hg) has been reported in assisting RTM mold fill-

ing (Mosher, 1995). Note that the tooling must be large enough to accommodate the perimeter gasket, air vents, injection ports and guide pins. 20.9 CURING

To convert a resin system into useful products it must be cured or cross-linked by chemical reaction into a three dimensional network. The reaction usually involves either a step growth polymerization, a chain growth polymerization, or a combination of both. The accompanying rheological change in the process is shown in Fig. 20.9 (Macosko, 1989). The curing step constitutes a major portion of a typical RTM cycle. During curing, rheological property changes of the resin system and heat transfer between the mold wall and the resin dictate the cure cycle. Simultaneously, modulus and strength begin to build up at a rate depending on the type of resin and catalyst used and the chemical kinetics of the resin system. Curing can continue after the part is demolded.

Matrix

Time

Fig. 20.9 Rheological change during the curing process. (Reproduced from Macosko.)

446 Resin transfer molding Cure cycle is dependent on part thickness, If the adhesion to the mold face is too strong, the ratio of catalyst or curing agent to resin even exceeding the strength of the composite, and the temperature of the mold and the resin it can be reduced by spraying release agents, system. In some cases, the part is removed normally fatty ester soaps or waxes, on the from the mold immediately after gel occurs. mold surface. After the two mold halves separate, the part The part must develop sufficient green can be removed from the cavity. Part removal strength for handling prior to its removal from methods range from the use of plastic/ the mold. Green strength is the strength a comwooden wedges and rubber mallets to the use posite exhibits after the resin gels, but prior to of knock out pins. A mold designed for low complete cure. Gel time is the interval of time throughput with hand operated clamps probetween introduction of catalyst or curing ducing a relatively simple, lightweight part agent to a thermosetting resin and the formawould most likely be removed using a wedge tion of a gel. Typical gel times range from several minutes to about an hour depending and mallet. Sophisticated hydraulic ejection systems can be used for high volume, complex on the factors mentioned above. The glass transition temperature, Tg,for an or heavy parts. To be pushed out, the part RTM resin system depends on thermal history. needs enough green strength to survive conFor a given temperature, the Tg increases dra- siderable bending stresses. The most common test for sufficient bendmatically with time until it levels off. As the curing temperature is raised, the T reaches a ing strength is to fold over a corner of the part steady-state value at a faster rate. &e steady- immediately after demolding. If the corner state value for Tg is a function of the curing survives the bending without cracks or a temperature, and usually approaches the cur- crease, the part is accepted. Otherwise, meaing temperature. However, the limit is sures to improve its green strength include bounded by the degradation temperature of any of the following steps: the resin system. 0 allow the part more time to cure in the mold; 20.10 DEMOLDING AND POST PROCESSING 0 increase the mold temperature; 0 modify or change the resin system, e.g. The minimum the curing step must accomincrease the catalyst level. plish is to develop sufficient green strength so that the part can be removed from the mold. There is often excess resin at the edges of the While cost is an important factor, it is not the part and in the vents. Considerable trimming, only criteria in choosing a method to remove a part of the post processing, is common when part from an RTM mold. For example, part reinforcement is clamped in the parting line. weight and complexity, and throughput are Trimming is required for almost all items important considerations. In many ways, the made by the RTM process. Accurate preform choice of ejection methods parallels the choice placement and precise alignment can reduce of clamping methods. the labor in this step. A few precautions are required to facilitate Postcure, one of the post processing operademolding. Before opening the mold halves, it tions, is used for various reasons. A molding is necessary to release the part from one mold cycle including postcure can increase producsurface. The force required is approximately tion throughput. While postcuring in an oven, that to overcome the adhesive force between the temperature is not restricted to that the mold and the composite. Typically, tears of allowed for the mold materials. Therefore a surface skin or flash, both resin rich, can be higher conversion of reactive groups can be found around the comers or edges of the part. achieved. It can also prevent the reaction

Process physics and use of simulations as a design tool 447 exotherms of a resin system from damaging a composite tool. It is important to hold the part shape during the process of postcuring and cooling to prevent distortion or warpage. 20.11 PROCESS PHYSICS AND USE OF SIMULATIONS AS A DESIGN TOOL

The processing defects addressed in the previous section are often caused by lack of a systematic treatment in RTM part design and process planning. Among the unresolved issues in RTM encountered by composite engineers, those related to the physical processing have developed rapidly during the last decade. The advancement in RTM technology demonstrates the potential of RTM becoming a primary process for producing many composite parts. In this section, the issues in reinforcement preforming, alternative tooling, mold flow analysis, and cure kinetics are revisited. The focus is on the use of models to describe and enhance the understanding of the physical phenomena. The models are built on the experimental evidence and observations, the goal being to reduce the scope of experiments in the engineering applications. Reducing engineering experimentation is achieved by combining three elements: mathematical models, numerical methods and computer software, into a simulation. One example of such software is LIMS (Liquid Injection Molding Simulation) (Advani et al. 1993) which has been developed specifically for mold filling of complex structures in RTM and can be used also as a design tool for manufacturing of complex structural parts as shown in Fig. 20.10. The topics will be presented in the order found in the unit operations of RTM. Draping of reinforcement plays the role of distributing fibers in a way that depends on the tool geometry. Simulation of reinforcement draping allows an engineer to estimate the fiber content distribution. This distribution can change the volume fraction as well as the orientation

Fig. 20.10 Complex structure manufactured by RTM.

in the molded part and therefore is of extreme importance. Tooling and mold construction are critical factors in successful RTM. By considering several alternative configurations, both the injection pressure and the filling time can be reduced. These alternative designs are valuable as the injection pressure tends to rise rapidly when inhomogeneous fiber distributions are present as a result of preforming. 20.11.1 PHYSICS GOVERNING RTM PROCESSING AND NUMERICAL SIMULATIONS

Darcy’s law for flow through porous media is conventionally used to describe the resin flow in the fiber reinforcement. The generalized form expresses the superficial velocity of resin flow in terms of a factor, which is permeability divided by fluid viscosity, multiplied by pressure gradient. This expression together with the mass conservation in the mold are solved together using various numerical methods. A typical example of this method combines finite element method with control volume method (Bruschke and Advani, 1990a, 1990b, 1991, 1994; Advani, 1994; Young et al., 1991a, 1991b). The solution is moved forward in time after


surface. The length of the cell segment can be changed as a result of slippage to accomodate this effect. A dome shaped part will serve as an example of this draping simulation. First, a square bidirectional mat is draped. The workpiece is initially configured so that warp and weft tows are perpendicular to each other. Then draping starts at an arbitrary point on the tool. The initial constraints used in this case study are prescribed along the central tows in both the warp and weft directions. The length of the cell segment is assumed to be constant. In the draped configuration shown in Fig. 20.11, the degree of deformation varies from 20.11.2 PREFORMING cell to cell. The minor angles in the preform For bidirectional mats, woven or stitched, range from 90" to a minimum of 35". The shear draping an arbitrary tool surface depends on also results in fiber volume fraction increase two deformation modes: shear deformation up to 70% for the dome. This information can and inter-yarn slip (Potter, 1979).A mat of this assist a designer in material selection, setup of nature is treated as a net that consists of many processing conditions and part design: a cells (Van West, 1990).Therefore, draping over process engineer can use this information to a surface of double curvature requires the net find out where to make necessary cuts in order to map on the surface by changing the internal to accommodate for induced deformation. As angles in each cell. The four sides of a cell are a rule of thumb, formability of preform mat made up of fiber tows. These tows, under the relies on absorption of such deformation by preforming condition, are inextensible. At the reinforcement material. A good material high deformation regions in a reinforcement, can withstand high deformation without slippage may be necessary to drape the tool wrinkle formation.

the pressure field is obtained during the filling process. The pressure solution obtained from the mold flow analysis can be used to position the gate and vent. This lends a design engineer 'infinite' options when facing the task of mold design. The design rules are no longer restricted to the rule of symmetry used by experienced designers to position the inlet and outlet ports. Instead, a composites engineer would be able to optimize the overall design based on criteria such as minimizing the injection pressure.

Fig. 20.11 Draped dome.

Process physics and use of simulations as a design tool 449 20.11.3 ALTERNATIVE TOOLING

or the preform is driven by the pressure difference. Therefore, the equation of motion is a One benefit of this process is that it can con- function of the bladder as well as the reinsolidate several complex three dimensional forcement material (Lucey, 1992). On the parts into one molded piece. The key to preform part, the compressibility of the reinaccomplish this is tool design. From the design forcement in the thickness direction plays a point of view, a flexible mold wall is very major role. On the bladder part, factors such as desirable to mold certain parts with difficult- inertia, damping, and rigidity of the elasor impossible-to-demold geometry. While a tomeric material can also be included when hard tool makes clamping and demolding dif- they are significant. ficult, the flexible mold provides a convenient From lateral compression tests, the alternative for mold design of these types of load-deflection curve of the fiber reinforceparts. Figure 20.10 shows an example of possi- ment material behaves like a nonlinear spring. ble features which may be molded using this The elastic constant of the preform depends on concept. In this part, one can easily see the its state. Preform permeability is a function of small draft angle and the stiffeners which can fiber architecture and porosity. Since the make demolding difficult. Moreover, the porosity of the preform changes with the beads and the flanged opening in the bulk- thickness, the permeability can be expressed in head of this frame are features that are terms of the cavity thickness. impossible to mold using conventional rigid Figure 20.12 was obtained from the numerimolds. cal simulation of two cases: one with rigid walls To avoid unconstrained wall movement, and the other with a flexible mold wall. In the the bladder pressure is higher than the injec- case with rigid mold walls, the pressure drops tion pressure. The motion of the flexible wall linearly with respect to the flow distance. This is

0

0.2

0.4

0.6

0.0

1

X

Fig. 20.12 Pressure distribution in the 1-D mold near the end of mold filling for flexible and rigid tool.


caused by the constant permeability of the pre- pressure drop and overall filling time which is form inside the mold. The pressure curve for the impossible to attain simultaneously in convencase with a flexible mold wall reflects the fact tional tooling (Fong and Advani, 1995). that the fluid flow in the filled region exhibits a smaller pressure drop. This reduction is benefi20.11.4 GATING, VENTING AND VOID cial to the molded parts as it causes less fiber CAPTURE washout and preform deformation due to the In this section, computer simulations for RTM resin. Figure 20.13 shows the results of computed mold filling are discussed to overcome the gatgap thickness of the 1-D mold with a flexible ing and venting problem. Mold filling mold wall. The straight line shows the thick- simulation is an effective way of positioning ness in a rigid tool. From this distribution, one injection and vent ports. Gating and venting can see that the gap height is a function of time are critical in the mold design because they during the filling process as well as a function determine whether complete wetout is achievof pressure. Near the injection gate, the resin able under normal operating conditions. A gate designed at the lowest part and vent pressure balances the applied pressure from at the highest point is generally a good practhe bladder and increases the gap thickness to tice to allow the air within the mold to be its maximum in the 1-D mold. As a result, the resistance to the incoming flow has reduced pushed out by the resin. Experienced designsignificantly as shown in the previous figure. ers may use symmetry to design the inlet Through the numerical study, the potential ports and outlet vents. However, the picture is of the flexible tool design has been demon- often complicated by the geometry or the strated. It has the advantage of reducing the presence of inserts. The engineering way to

0.005

0.004

0.003

0.002

0.001

0 0

0.2

0.4

0.6

0.8

X Fig. 20.13 Gap thickness variation in the 1-D mold near the end of mold filling.

1

Process physics and use of simulations as a design tool 451 ensure complete wetout initially is to gate the the design tool. The situation would be much mold correctly in the design. more complicated if mold filling is coupled Figure 20.14(a) shows a square plate with with phenomena such as preform deformation two cutouts in the part. The injection port is and channeling in the corners and along the first positioned at the center of the lowest part. edges. The flow fronts corresponding to the gate At the microscopic level, heterogeneities design are indicated by the curved contours. always exist in the preform media. For examContours in this figure indicate different time ple, the fiber tow may have a permeability steps. For example, the contours closer to the several orders lower than that of the intergate represents area that is filled first and the stices. Therefore, micro-voids form when the contours closer to the vent the last filled orientation of the fiber tow does not allow the region. As a result of colliding flow fronts in displacement of the air inside the tow. A novel the middle and top portion of the part, the fig- approach in mold filling analysis is reported ure demonstrates the capturing of dry patches by modifying the equation of mass conservaor macro-voids. These voids can degrade the tion to account for the fluid absorbed by the properties of the molded composite signifi- fiber tows (Fong and Advani, 1994). Void cantly. Void capturing is important in the entrapment inside tows is found to be depenprocess simulation to avoid formation of such dent on the microstructure, the vent pressure, defects. Figure 20.14(b) shows an alternative and the ratio of the difference in the permedesign that eliminates venting in the middle of ability of the tows and the permeability of the the part. As a result of injection in the corner, preform (Pillai and Advani, 1994; the vent port has to be positioned differently. Ranganathan et al., 1994). This demonstrates the power and simplicity of

Vent Port

4

I Injection port

(b)

I

Injection Port

Fig. 20.14 Design of injection ports: (a) central injection; (b) corner injection.

452 Resin transfer molding 20.11.5 SENSOR CONTROLLED INJECTION

Sensor controlled injection is multiple injection in an 'intelligent' way without involving a complex control algorithm. It requires placement of gates along the flow path at a number of locations. The injection gate is also a sensor capable of detecting the arrival of resin. These gates are then activated or deactivated in the order of first on, first off, and, therefore, allow the mold to fill in a series of steps. For example, Fig. 20.15 is a simple mold which has four injection gates. To help visualize the concept effectively, a 1-D mold is used. The T-column represents different time stages in the filling process. In this example, only one gate is allowed to open at a time. As the injection starts at T1, the first gate is open and the remaining three are closed. As the flow front progresses through the mold, it hits the second gate location at time T2. The injection unit shuts off the resin to gate one and opens the second injection gate. Instead of having the resin flow through the whole length in the mold, the length is divided up into a number of intervals. Therefore, the overall flow resistance decreases as the effort required is for the resin to flow from one gate to the next closest gate in the flow path.

Figure 20.16 shows the pressure calculation from the mold filling simulation. The pressure drops linearly in the one dimensional flow. As the flow front progresses from the inlet toward the vent under a constant flow rate boundary condition at the inlet, the pressure build-up looks like the schematic shown in the lower left figure. For the multiple gate with sensor controlled injection, the pressure at the first gate increases up to a limit when the flow front hits the next sensor. When the next gate is open, the previous gate is shut off. So the pressure build-up is only limited by the length of the interval. Therefore, the maximum pressure seen in the mold is only a fraction of the pressure compared to the lower left figure.

n q--\n E

a

Flow Length

Flow Length

CONTROL SCHEME T1 T2 T3 T4

ON OFF OFF OFF

OFF ON OFF OFF

/

OFF OFF ON OFF

OFF OFF OFF ON

Mold

Fig. 20.16 Representation of pressure during mold filling.

Table 20.1 shows the results from two sets of computer simulations. For either case, only one gate is open at any time during the mold filling stage. The first column uses constant flow rate and the second column uses constant Table 20.1 Comparison of processing parameters

Gate with flow front sensor

PP

Fig. 20.15 Control scheme for a 1-D model.

Single-gate injection Sensor-controlled injection

Slnglegnfr

100% 48%

s2nglegate

100% 36%

Process physics and use of simulations as a design tool 453 pressure. If only one injection gate is used, the pressure under the constant flow rate boundary condition will reach a maximum. Compared to the sensor controlled injection with four gates, the pressure at the gate is only 48% of the pressure reached by the single gate injection. In terms of filling time, the two molds are subject to a constant pressure boundary condition. Results show that the mold filling for the single gate injection takes almost three times that for the sensor-controlled injection. An example is shown in Fig. 20.17, which elaborates on how one can utilize a sensor to eliminate a dry spot during molding. In Fig. 20.17(a), where no sensors are implemented and the injection gate is at the location as shown, a dry spot will appear in the middle of the part. However, an extra gate in the middle as shown in Fig. 20.17(b), if triggered at the point the fluid reaches the midframe, can prevent this void, as indicated by Fig. 20.17(c). This feature is incorporated in a numerical simulation such as LIMS and can be systematically studied for a given geometry to decide the best strategy when in situ sensing capabilities are incorporated in the fabrication phase (Liu et al. (1995)).

t Vent

20.11.6 MOLD FILLING WITH RESIN DELIVERY SYSTEM

Conventionally, an injection port serves as a point source where fluid is pumped. The drawback of a point source is that the pressure value tends to rise rapidly to an extent that could be detrimental to the preform. By extending the point source into other forms proves to be effective in reducing the pressure build-up. To implement this concept, one can use multiple point sources as discussed previously. A line source has been popular in vacuum assisted RTM because of its ability to fill the mold using 1 atm of pressure. A line source may be modified to serve as a runner by allowing more fiberfree space in this delivery system. This is the channeling effect, now used to advantage in mold filling. Further extending the fluid source may possibly yield a 'plane' source. The actual implementation of a plane source may include a high-porosity layer in the stack-up of the reinforcement mats. The layer can possess a permeability several orders higher than that of the fiber preform. The result of this is a three-dimensional mold flow with fluid propagating rapidly through the spreading plane or surface first

1Vent

Fig. 20.17 Use of sensors to eliminatedry spots: (a) no sensor; (b) extra gate sensor; (c) void prevented.

454 Resin transfer molding followed by percolation of the resin through the thickness of the preform. For three-dimensional flow, venting the mold may become less intuitive. In practice, vacuum assistance can provide part of the solution. 20.12 CONCLUSIONS

Resin transfer molding is a practical process for much of the composite industry. The quality of RTM molded parts can equal that by conventional autoclave processes and its economic advantages are obvious. Although the underlying principles of RTM appear at first to be simple, this is often not the case. The challenge for RTM is to bring together the disciplines of preforming, mold design and process development with existing fibers and resins. This can be best achieved through an understanding of the physics governing RTM and by current simulation technology. REFERENCES

Advani S.G., Bruschke, M.V. and Parnas, R., 1994, Resin Transfer Molding, in Flow and Rheology in Polymeric Composites Manufacturing (Ed S.G. Advani,) Amsterdam: Elsevier Publishers, Ch 12, pp. 465-516. Advani S.G., Bruschke, M.V. and Liu, B., LIMS 3.0: Liquid Injection Molding Simulation, User Manual, CCM Report, University of Delaware, Newark, DE 19716. Bruschke, M.V. and Advani, S.G., 1994, A numerical approach to model nonisothermal, viscous flow with free surfaces through fibrous media, Intern. J. Num. Methods Fluids, 19,575-603. Bruschke, M.V. and Advani, S.G., 1991, RTM: Filling simulation of complex three-dimensional shelllike structures, SAMPE Quarterly, 23(1), 2-11. Bruschke, M. and Advani, S.G., 1990, A finite element/control volume approach to mold filling in anisotropic porous media, Polym. Comp., 11, 398-405. Bruschke, M.V. and Advani, S.G., 1990, Mold filling of generalized newtonian fluids in anisotropic porous media, Transport Phenomena in Material Processing, ASME Trans. HTD 132, 149-158. Carvalho, R.L., Personal communication, Manager

of Application Support Laboratory, Fiber Glass Reinforcements Division, Vetrotex CertainTeed Corporation, 1991. Chou, T.W., 1992, Microstructural Design of Fiber Composites, Cambridge: Cambridge University Press, UK. Isorca Inc., 1992, Introduction to Resin Transfer Molding, Society of Plastics Industry, Composites Institute. Fong, L., J. Xu, and Lee, L,J., 1994, Analysis of thermoformable fiber mat preforming in liquid composite molding: study of deformation modes and reinforcement characterization, Polym. Comp., 15, 134. Fong, L. and Advani, S.G., 1994, The role of drapability of fiber preforms in resin transfer molding, Amer. SOC.Comp., Proc. 9th Tech. Conf., 1246. Fong, L., and Lee, L.J., 1994, Analysis of fiber mat preforming in liquid composite molding, preforming induced effects on mold filling, J. Reinf. Plas. Comp., 13, 637. Fong, L., Varma. R.R. and Advani, S.G., 1994, Use of process models and simulations as design tools in molding polymer and polymer composites, The Pacfic Conference on Rheology and Polymer Processing (PCR’94), Kyoto, Japan. Fong, L. and Advani, S.G., 1994, The role of dual permeability of fiber preforms in mold filling simulation of resin transfer molding, Proc. Zst Intern. Conf. Comp. Engng, New Orleans, LA, 17. Fong, L., Liu, B. and Advani, S.G., 1995, Modeling and simulation of resin transfer molding with flexible mold walls, 50th Ann. Conf., SPI, Comp. Inst., Session 3-A. Krolewski, S. and Busch, J., 1990, The competitive position of selected composites fabrication technologies for automotive applications, Proc. 35th Intern. SAMPE Symp., pp. 1761-1771. Lee, S.M. International Encyclopedia of Composites, 1991, New York: VCH, 1991. Liu, B., Bickerton S. and Advani, S.G., 1994, Modeling and simulation of RTM - venting and void formation, Proc. Intern. Conf. Comp. Engng, p. 17. Lucey, A.D. and Carpenter, P.W., 1992,J. Fluid Mech., 234, 121. Macosko, C.W., 1989, RIM, Fundamentals of Reaction Injection Molding, New York: Karl Hanser Verlag. Mosher, I?, 1995, An introduction to vacuumassisted resin transfer molding (SCRIMP), 50th Techn. Conf., SPI, Comp. Inst., Session 8.

References 455 Pillai K. and Advani, S. G., 1994, The role of dual permeability of fiber preforms in resin transfer molding, Proc. 9th Am. SOC.Comp., p. 17. Potter, K.D., 1979, Composites, lg 161. Ranganathan, S., Wise, G.M., Phelan, F.R., Jr., Parnas, R.S. and Advani, S.G., 1994, A numerical and experimental study of the permeability of fiber preforms, Proc. 10th ASM/ESD Adv. Cornp. Con$, 309. Scheidegger, A.E., 1974, The Physics of Flow through Porous Media, University of Toronto Press. Tucker, C.L., 1989, Fundamentals of Computer Modeling for Polymer Processing, New York: Karl Hanser Verlag.

Vanwest, B.P., Pipes, R.B., Keefe, M. and Advani, S.G., 1991, The draping and consolidation of commingled fabrics, Comp. Manufng, 2, pp. 10-21. Young, W.B., Rupel, K., Han, K., Lee, L.J. and Liou, M.J., 1991a, Polym. Comp., 12, 30. Young, W.B., Han, K., Fong, L., Lee, L.J. and Liou, M.J., 1991b, Flow simulation in molds with preplaced fiber mats, Polyrn. Comp., 12,391.

FILAMENT WINDING

21

Yu.M. Tarnopollskii, S.T. Peters, A.I. Bed’

21.1 INTRODUCTION

A winding operation is the basic fabrication technique for forming load-bearing structural elements made of polymer matrix-based fibrous composites, which have the shapes of bodies of revolution. A semifabricated product (uncured preform) of previously impregnated filaments, strands, tapes and fabrics is wound layer by layer with controlled tension onto the mandrel or previous layers. By varying the angle of filament or tape placement, it is possible to control the reinforcement fiber angles within the same layer and through the thickness of the composite wall. During winding, fiber tension generates pressure between layers of uncured composite; this pressure influences the compaction and void content of the article and contributes to more complete utilization of the strength and stiffness of the reinforcing fibers. If the contact pressure is insufficient for compaction of the material, additional layer-by-layer compaction of a semifabricated product must be employed. The wound article must be converted by chemical and/or thermal means to the finished article. With heat treatment, the usual method, the temperature can be constant or can vary with time. The mandrel defines the internal shape of the article. It is removed after curing if the mandrel is not an element of the structure. The winding process is illustrated in Fig. 21.11,*.


Filament winding is a natural way to combine two-dimensional reinforcement and, with additional processes and devices, threedimensional reinforcement. Advanced processes, combining filament winding and braiding, allow fabrication of spatially sewn structures. The most important groups of wound articles are: thin-walled shells (their thickness is negligible compared to their radius); compound structures, including three-layered and multi-layered shells with a light foam or honeycomb plastic filler; and thick-walled structural elements. For thin-walled shells, it is most important to optimize the reinforcement configuration. For compound surfaces, contact pressure on the interface is a design parameter. Also, there is the problem of monolithicity for thick-walled structural elements, which is closely related to the problem of control of residual stresses. Most composites cannot sustain a significant internal pressure over time without leaking (weeping) through the otherwise sound composite wall. The use of internal liners made from rubber, plastic or metal can provide a leak tight structure. The semifabricated (uncured) article has extremely low strength in the radial direction. Thus, the ratio of elastic modulii along and across fibers can reach lo3. As additional circuits are wound with tension onto the earlier ’soft’ laminate the previous circuits are compressed with a redistribution of preset tension. The process is typical for all types of filament wound articles with specific processing approaches for each3.Controlling the winding

Introduction 457

q=t(t)

CURING (HEAT BUILDUP, CURING COOLING EXTRA 6OMPACTiAN

TENSIONING

REMOVAL FROM MANDREL

ARTICLE

W

SEMIFABRICATED MATERIAL

INITIAL STRESSES

Fig. 21.1 Stages of winding process.

process may include specifymg fiber tensioning and extra compaction pressure (internal or external), or geometric parameters (configuration of reinforcement lay-up along the length and through the thickness of the article). The reinforcement configuration is determined by the design requirements of the article and can be achieved with high accuracy. Winding, as a way of obtaining preset reinforcement configurations, is restricted in terms of available winding angles. Filament lay-up on geodesic lines on an external surface is easy. (A geodesic line is a line linking two preset points of the surface along a shortest path.) In non-geodesic winding, filament tension can shift the filament

to the geodesic line, against the static friction or viscous resistance of the binder, or mechanical obstacles, such as pins, to keep the filaments on non-geodesic curve. Larger components will result in changes to the system of residual stresses. To calculate, analyze and control the values of residual stresses, it is necessary to find the interrelation between the winding parameters (mainly tensioning) and the properties of the finished product. By controlling filament tensioning and the reinforcement configuration, it is possible to control strength and stiffness of the material, residual stresses and pressures on the mandrel or previous layers.

458 Filament winding 21.2 HISTORY OF WINDING

Historically, the developments in winding technology can be traced back to 1370 BC. There is an ancient Egyptian flask in the British Museum made by winding glass fibers from the melt onto a dissolvable mandrel. Embalmed mummies were wrapped with tapes soaked in resin that were the natural precursors of polyester resins. Rope wrappings were used for reinforcing the bamboo powder rockets in ancient China. Guns with wire wound barrels for reinforcing and proper stress distribution were fabricated in Russia as early as 1869 and in Germany from 1900. In the development of fiber winding technology there have been three important stagesP6:the adoption of the first patint in 19637; the creation of NOL-ring sample4for strength testing in 1955; and the publishing of the first monograph wholly devoted to composite winding technology by Rosato and Grove in 1964R.All these works addressed the development of thin-walled articles rather than the task of converting a semifabricated or uncured material into a monolithic article, which is the primary objective for modern composites technology. The history of analytical development includes the paper by L. Euler on thread equilibrium on a rough surface, papers by P. Laplace and G. Lame on equilibrium of shells loaded by pressure, a paper by A.W. Gadolin on compound cylinders with preliminary tightness and his formulation of the problem on tension wrapping of pressure vessels by wire and papers by A.P. Minakov on thread equilibrium on an arbitrary rough surface. The winding mechanics of an extendible wire have been described by R.V. Southwell. The state of the art of the winding mechanics of composites has been reviewed in detail2.

following stages: winding of a preform; heating of a preform with the mandrel up to temperature of binder polymerization; curing at constant or variable temperature; cooling to below the glass transition temperature and then to room temperature; and removal of the article from the mandrel. As a rule, thermosetting binders are commonly used in winding, although recently progress has been made with advanced thermoplastic binders. Among the reinforcing fibers, the most widely used are fiberglass, carbon, organic (aramid and p~lyethylene)~,~. They can be used in the form of single filaments, rovings, woven and unwoven tapes, strands and woven fabrics. 21.3.2 MATERIALS FOR WINDING

There is a distinction between winding with prepreg, when the reinforcing material is impregnated with a binder by an outside vendor and subsequently partially polymerized to a ' B stage and 'wet' winding, and when the processes of reinforcement impregnation with the binder and the winding operation are performed on the same equipment simultaneously. The rarely used process of 'dry' winding involves winding of the reinforcing material without any binder on the mandrel, followed by vacuum or any other impregnation of the article with the binder and curing. This latter process will see increasing attention since it is a convenient and inexpensive method for providing a stable preform without edges for resin transfer molding. The winding operation of thermoplastics can involve, in the case of in situ consolidation, dilution of the high viscosity of the matrix material by heat or organic solvents. It is physically analogous to the effect of impregnation in 'wet' winding or preliminary heating of a 21.3 TECHNOLOGY OF WINDING prepreg. Although a 'curing' stage is missing in the winding operation of thermoplastics, 21.3.1 STAGES OF PROCESS heat treatment and compaction, as a rule, will The production cycle for most filament lead to significant changes in macromolecular wound composites can be subdivided into the structure. These changes can result in physical

Technology of winding 459 shrinkage analogous to chemical and thermal shrinkage in the curing and cooling of thermosetting plastics. The analysis of winding techniques and general engineering winding theory for thermosetting plastics are surveyed in detail6.

Winding of more complex shapes aggravates the problem of pattern closure. To obtain a continuous layer, it may be necessary to vary the bandwidth during winding. Also, it may be advantageous to vary the fiber volume and thickness by alternate compaction techniques.

21.3.3 GEOMETRY OF WINDING

21.3.4 EQUILIBRIUM CONDITIONS OF FILAMENT ON THE SURFACE

The shape of the article to be wound is determined by the mandrel. The thickness of the article is controlled by the number of circuits wound (taking into account their compressibility during winding and shrinkage during cooling). The reinforcement configuration is determined by controlling the wind angle, which is the angle formed between the tangent to the filament and tangent to the intersection line of the surface to be wound with the plane parallel to the running axis of winding. For a body of revolution it will be the angle between the filament and the wind axis. The formation of a continuous layer with finite thickness and bandwidths, and with proper fiber geometry, requires a coverage solution. If the bandwidth, the number of circuits to close and the winding angle are carefully chosen, the adjacent circuits along the axis will be linked in butt joints, forming a smooth continuous covering. If the bandwidth is greater, the adjacent circuits will form a lap joint resulting in a rough surface. If the band width is smaller, there will be a gap between the adjacent circuits along the axis. This gap can be partly or completely filled up by either fiber or matrix of subsequent layers. Some structures may be amenable to an incomplete winding pattern. Winding of subsequent circuits is feasible either directly on the preceding ones (winding of ribbed and honeycombed structures, Fig. 21.1), or on the rough outer surface formed from several preceding layers. In the second version, the reinforcing fibers undergo additional bending and interweaving and the structure obtained may have more voids and lower fiber volume than in the case of winding with butt joints.

Winding along geodesic lines is easier than along non-geodesic lines. However, to obtain a continuous covering on h e surface it may be necessary to use non-geodesic winding. This presents the problem of permissible deviations from the geodesic lines. The angle of geodesic deviation can be designated by v, i.e. the angle between the main normal v to the curve, along which the filament has been laid, and the normal to the surface n at the same point. The curve on the surface is defined by the radius of curvature Rc and the radius of twist R, and the radius of normal curvature Rn (projection of curvature radius of the curve on normal n to the surface): l / R n = cosv/Rc and radius of geodesic curvature Rg: l / R g = sinv/R, The equilibrium equation for a filament on the surface has the form (Fig. 21.2): ~

~

dN ds

N Rn

+ FT+dQr =

0

+ F,, + p Q , - P = 0

N -+F

+ p Qg = 0

(21.1)

(21.2) (21.3)

RP

where N is tension of filament, s is the arc coordinate on the curve, F represents surface distributed forces, related to a unit length, p is the linear mass density of filament, Q represents mass distributed forces, referred to a mass unit, including dead weight and inertial forces,

460 Filament winding

Fig. 21.2 Condition of filament equilibrium on the surface.

t is

the basis vector directed along the tangent to the curve and P is the normal surface reaction related to a length unit of filament. Equation (21.l) describes the redistribution of tension along the filament length due to friction, viscous resistance from lower-lying circuits, viscous resistance of above-lying circuits and inertial effects. Equation (21.3) represents the force, which tends to shift the filament onto a geodesic line and viscous resistance of both the lower-lying and abovelying circuits. In the simplest case of the friction law I Fg I 5 kP where k is the friction coefficient, from equation (21.2) (at Fn = Q, = 0) and equation (21.3) (at Q = 0) the following relationship can be derive& -Rn =tanv 0.3

Efforts were made to describe the d a l d t versus a data with a modified Arrhenius type equation. The proposed empirical equations are

where

26.2.2 FIBER DEFORMATION

K , = A, exp (-AE,/RT)

AE, = 5.66 x lo4J mol-'

The main contribution from Gutowski's model is the description of fiber deformation behavior. Instead of treating fibers as separate layers, a network concept is introduced. In other words, fiber-to-fiber contact is assumed within a fiber assembly, even in the case of aligned fiber bundles. Thus a fiber filament span between the neighboring contact points becomes a small bending beam. During a consolidation process when fibers are pushed closer, more and more fiber-to-fiber contacts take place, and the span length reduces. Thus the bending stiffness of these small fiber beams increases rapidly, resulting in nonlinear elastic deformation response. The nonlinear elastic response of a fiber assembly under a compressive load has been also studied in the textile field, and an empirical formula was proposed (van Wyk, 1946). A proposed fiber deformation model for aligned fiber bundles considers the deformation status variable, the fiber volume fraction V f ,as a function of the consolidation pressure (Gutowski, 1985).The expression is

As can be seen from the discussion, all the constants involved in the model are determined experimentally through a specified process. Similar treatment can be used for other types of resin systems, and experimental investigation results have been reported, including Hercules HBRF-55 Resin (Bhiet aI., 1987)and Fiberite 976 Resin (Dusiet al., 1987).A similar process model has also been discussed by Roylance (1988).

where V , is the maximum obtainable fiber volume fraction for a given fiber network configuration, and V , is the fiber volume fraction below which the fiber network carries no load. The empirical constant As is obtained from curve fitting on available measurement data. A typical fiber deformation curve for

K2 = A, exp (-AE,/RT)

K, = A, exp (-AE,/RT) A,, A, and A, are the pre-exponential factors, AE,, AE, and AE, are the activation energies, R is the universal gas constant, and T is the absolute temperature. The constants in the expression are found as:

B = 0.47 A, = 2.101 x lo9 min-'

A, = - 2 . 0 1 4 ~ l O ~ r n i n - ~ A, = 1.960 x lo5 min-l AE, = 8.07 x lo4J mol-' AE, = 7.78 x 104 J mol-'

Consolidation models 581 well aligned graphite fibers is shown in Fig. 26.5 with the- co-mparison of measured data points.

In

,Q 700

b

In v)

? 500

Data Point-

al

-1 300 e

:200

.c

100

n-

0.4 0.5 0.6 0.7 0.8 0.9 Fiber Volume Fraction ( V f )

Fig. 26.5 Typical fiber deformation curve for wellaligned XA-S and A S 4 graphite fibers (Gutowski et al., 198%).

26.2.3 CONSOLIDATIONMODELS

As discussed above, in Springer’s model, it is assumed that there is no fiber-to-fiber contact. Thus a dynamic fluid pressure exists between the consolidated layers. The consolidation time, which is crucial to the cure process, is related to the permeability of the fibrous preforms, resin viscosity, and the applied consolidation pressure. In Gutowski’s model, the fiber reinforcement and the fluid state resin are considered as a system. Both fiber network deformation and fluid resin flow are solved together. Both models are presented here with a laminated composite structure as the example. The example for Springer’s model is the laminate consolidation with flow in the laminate transverse direction, or z direction. A bleeder ply is assumed to be placed on top of the composite. Figure 26.6 shows the setup for the model. At any instant of time the liquid velocities in the bleeder Vb and in the composite V care given by Darcy’s law. For a constant viscosity liquid, the integrated forms are:

K c (Po - P,)

The proposed relationship between the compressive fiber stress ofand fiber volume fraction V , provides a tool to estimate the finished consolidation status of the composite products. If the time window for the consolidation is long enough, and excessive resin is completely squeezed out from the structure, the consolidation pressure is then balanced by the fiber stress. However, because of the dramatic change of the resin viscosity and preform permeability during a consolidation process, resin flow may not be complete. Thus, developed consolidation simulation models are needed for the process analysis and improvement. During the compression of fibrous preforms, structural relaxation has been observed (Gutowski, 198%). Thus the deformation to some extent is not elastic but viscoelastic. This issue has been addressed by using a Maxwell type model (Kim, McCarthy and Fanucci, 1991).

9, =

7

(26.7)

hC

where p , and pb are the pressures at the composite-bleeder interface and in the bleeder respectively, po is the consolidation pressure and is related to the applied force or pressure, is the instantaneous thickness of the liquid in the bleeder, and hc is the thickness of the resin starved layer, or the thickness of the layers through which resin flow takes place, and K, and Kb are the permeability of the composite layer and bleeder respectively. If the compacted composite layer thickness is h,, then hc = nh,

(26.9)

where n is the number of layers or plies already compacted.

582 Consolidation techniques and cure control

L

Resin Flow

Fig. 26.6 Illustration of the consolidationmodel proposed by Springer (1982).

layers. The final status of the composite is dependent on the compaction of each individual layer. -~ d(hA) = Aq, = Aq, As a comparison, Gutowski’s consolidation (26.10) dt model combines the flow of resin through where A is the surface area of the composite porous media and the fiber deformation laminate, and h is the total thickness of the behavior. Similar treatment has been precomposite laminate. The second equation sented in studies of other fields including soil expresses the fact that at any instant of time, mechanics (Biot, 1941,1955,1956; Gibson and the flow out of the composite is equal to the Hussey, 1967).In general, consolidation occurs flow into the bleeder. The pressure po is related in only one direction, but flow may take place in all three directions. Thus an element is to the applied force as: deformable in the z direction. A new variable 6 f (26.11) is used to represent the deformation, and 6 = z = + pa + w where w is the local displacement of the where F is the applied force and pa is the fiber network. The laminate setup for the atmospheric pressure. By combining these model is illustrated in Fig. 26.7. If the initial equations, the consolidation equation fiber volume fraction for the composite is V, and the fiber volume fraction at any instant is becomes: Vf, the fiber continuity condition states The equation of continuity gives the rate of change of volume of the composite as:

vo=-v, 36 Therefore for each individual layer, the consolidation time can be calculated. The total consolidation time is the summation for these

aZ

(26.13)

Resin flow continuity condition requires:

Consolidation models 583 Here it is assumed that the inertial effects in the process are small. Therefore the applied pressure is balanced by a combination of the average resin pressure and the fiber stress. In (26.14) other words, any load which is carried by the fibers is then unavailable for pressurizing the With the application of Darcy's law, a consoli- resin. dation equation using the fluid pressure p , and Since both the permeability and the fiber fiber volume fraction Vf as variables can be stress are expressed as functions of fiber volwritten as ume fraction V , with the given initial and boundary conditions, the variables V ,and p , as a function of time and location can be solved. In general numerical calculation procedures have to be developed for solving the partial differential equations. In some simplified cases, analytical solutions are possible. This equation gives a relationship between the spatial and time-varying nature of the presExample problem 1: One-dimensional flow sure in the resin and the fiber volume fraction in compression molding of the composite. The equilibrium statement for the consolidation is: A simplified example of composite consolidation is the compression molding of a flat (26.16) rectangular laminate. The composite part is A pressed between two solid dies. Therefore only in-plane flow is possible. In other words, flow components are in the x and y directions only. If the initial fiber volume fraction is uniform, the equation of the resin flow and fiber deformation becomes:

J2Pr + K - a2pr + -~ P av, = 0 (26.17) K

X

~y

ay2

v,

at

ho Here it is also assumed that there is no signifi-

a€

Fig. 26.7 Illustration of the consolidation model proposed by Gutowski et al. (1987a).

cant pressure gradient in the z direction, and the viscosity p does not vary spatially. In some cases, K Z / a 2>> Ky/b2 where a and b are the dimensions of the laminate in x and y directions respectively. The compression molding results in primarily one-dimensional flow in the x direction. Then the equation can be solved analytically. With the assumed boundary conditions of p , = 0 at x = M and ap,/ax = 0 at x = 0, the result is a parabolic pressure distribution as


The solution for the fiber volume fraction Vf as a function of time is:

Po = Of(Vf) 3 pa' K _ Vdvf ,T +

*

(26.19)

L

This expression shows how the applied pressure p, is carried by the fiber stress G~ and the average pressure in the resin. The load sharing in a composite is directly analogous to how the load is shared in a parallel spring and damper set. For example, initially if Vf is less than V,, then there is no deformation in the spring (fibers) and the entire load is carried by the resin. On the other hand, at long times and finite viscosity, if the rate of change of Vf is close to zero, then the pressure in the damper (resin) goes to zero and the total load must be carried by the fibers. Figure 26.8 shows an example of the one-dimensional flow in compression molding with the comparison of computer simulation results.

Example problem 2 Compression molding with two-dimensional flow Here the case of compression molding of a rectangular laminate with an isotropic in-plane permeability is considered. In other words, Kx = Ky = K. This may correspond to a quasiisotropic lay-up. The flow equation becomes Poisson's equation, which can be solved by the separation of variables technique. The solution for the pressure distribution in a laminate with zero pressure at the boundaries is:

With the applied load balance condition, the final result is:

It can be seen that the result is analogous to the previous case except for a geometry effect term which is shown in the bracket. 600

- PR,Theory

- 400 -

0

PR,Measured

'3 0

Time ( m i d

Fig. 26.8 Example of one-dimensional flow in compression molding and computer simulation results (Gutowski, et al., 198%).

Example problem 3: Bleeder ply molding This has been presented with the Springer's model. In this case, a porous bleeder ply is placed on top of the composite, and flow is principally in the z direction. With the introduction of a new variable, the void ratio e = (1- Vf)/V, one may obtain the nonlinear one-dimensional consolidation equation. An equation similar to this was first derived by Gibson et al. (1967) for the consolidation of saturated clays. The expression is:

(

")

de = (e, + 1)2- a - Kz 'Of at 3.z p ( l + e ) ' e az

-

(26.22)

Consolidation models 585 The void ratio e or the fiber volume fraction V, is a function of both time and location. An equivalent equation using variables Vf and p , can be written as

With similar pressure equilibrium conditions, the distribution and time history of Vf or e can be solved numerically. Figure 26.9 shows an example of the bleeder ply molding measurement setup, and the comparison of the computer simulation results with the measured data. It is interesting to see that, with Gutowski’s consolidation model, the final status of the composite in terms of the average fiber volume fraction can be estimated from the proposed fiber deformation model if the consolidation process is complete. The consolidation time for a particular setup can be solved through numerical simulation. As can be seen from the analysis, the total consolidation time for a composite structure is strongly dependent on the dimension in the resin flow direction. For laminated composite structures, usually the dimensions in x and y directions (directions within the laminate structure) are much larger than that in the z direction (direction transverse to the laminate plane). For example, many aerospace structural parts range from a few inches to several feet in x or y direction, but only have a thickness of a fraction of an inch in the z direction. Thus the bleeder ply molding process is preferred and is widely used in many part fabrication processes. However, for the socalled thick composites, for example with lay-up of 64 or 96 plies, the consolidation time required increases dramatically in the bleeder ply molding cases. With the selected cure cycle for thin composites, complete consolidation may not be achieved for thick composites. Thus the final fiber volume fraction of the thick composite tends to be relatively lower. This has been observed in experiments involving thick

DT TRANSDUCER

600

/Applied

Pressure

doto -theory o

500 -

-.-

Modified Cormon-Kozemy

400-

v)

a v 300

-

3

L

a

\o

IO0

\O

P 0 0

10

20

30

40

50

60

Time (rnin) Fig. 26.9 Example of bleeder molding and computer simulation results (Gutowskiet al., 198%).

composites (Kim, Jun and Lee, 1989). It can also be seen from the comparison of the two models that with relatively low fiber volume fraction, fibers carry almost no load. Thus the consolidation process is dominated by the resin flow through the fiber network. Then the difference between the two models is very minor. Springer assumes the consolidation is done layer by layer, while Gutowski treats the fiber network as a whole system. However, in both cases the top layers are consolidated first. When the fiber volume fraction becomes high, then the predictions from the


two models show significant different results. On the other hand, the numerical schemes of the two models are different. Springer’s model requires only the solutions of a series algebraic equations, while in Gutowski’s model nonlinear partial differential equations have to be solved. A comparison study has been presented by Smith and Poursartip (1993). 26.3 CURE CONTROL

Fiber reinforced thermosetting resin composites manufactured in autoclaves are made by forming the uncured fiber-resin mixture into the desired shape and then curing the material. Curing requires the application of heat and pressure. Heat is used to facilitate and control the chemical reactions of the resin, and pressure is used to consolidate the composite, squeeze out the excess resin, and minimize the void content. A cure cycle usually means the magnitude, duration, and profile of the temperature and pressure applied during a curing process. Selection of the cure cycle directly affects the quality of the finished composite product, such as fiber content, fiber distribution, and void percentage.

Viscosity

Specifically, a selected cure cycle must ensure that:

the temperature inside the material does not exceed a preset value at any time during the cure; 2. at the end of the cure the resin content is uniform and has the desired value; 3. the material is cured uniformly and completely; 4. the cured composite has the lowest possible void content; 5. the cured composite has the desired thermal and mechanical properties; 6. the curing is achieved in the shortest time. Figure 26.10 shows schematically the overall cure process model structure. In an early study, Loos and Springer (1983a)proposed a thermochemical model. Heat transfer from the environment to the composite material determines the temperature distribution, the degree of cure of the resin, and the resin viscosity within the composite structure. The temperature inside the composite can be calculated using the law of conservation of energy. By neglecting the energy transfer by convection, the energy equation can be expressed as:

b

Flow

/

Reaction kinetics

\

Heat transfer

Residual stress

Fig. 26.10 Schematic of overall cure process model (Dave et al., 1990).

Cure control 587 resin viscosity, the degree of cure a and the rate of the cure da/dt can be characterized using a modified Arrhenius type equation, with relevant constants in the model deterdH -k+p(26.24) mined experimentally (Lee, Loos and dt Springer, 1982; Bhi et al., 1987; Dusi et al., 1987; where p and Cv are the density and specific Roylance, 1988).Figure 26.11 show an example heat of the composite, kx, k and k, are the ther- of the rate of heat generation and rate of mal conductivities, and ?is temperature. In degree of cure of the 3501-6 resin system as the case of relatively thin composite structure, functions of time and temperature. conduction heat transfer is mainly in the z It is noted that the densityp, specific heat Cy, direction. Thus terms in the x and y directions heat of reaction Hr, and thermal conductivity k can be dropped. The rate of heat generation are all dependent on the instantaneous and dH/dt is defined as: local resin and fiber contents of each ply, and

i ~zE) (

(26.25) 01

c3501-6

400K

1 05

450K

1

where H, is the total heat of reaction depending on the resin type. The rate of the cure reaction is a function of temperature and the cure status, and can be expressed symbolically as: da dt

- = f(T,

4

(26.26)

The degree of cure is then determined as:

)$(:I

(26.27)

dt

a =

o f \

It is assumed that for an uncured material, a = 0, and for a completely cured material, a approaches unity. As discussed earlier the

0 0

,

02

,

Iojo\J

06 0 02 DEGREE OF C U R E , a

04

04

06

TEMPERATURE ( K l I

I

1

1

I

I

3501 - 6

-

-1

0 5 r n coI/sec

-

s \ I

,

,

,

,

~

,

,

l

,

l

l

l

l

Fig. 26.11 Rate of heat generation and rate of degree of cure of the

588 Consolidation techniques and cure control can be handled using rule of mixtures (Loos and Springer, 1983c) or proposed approximate formulas (Springer and Tsai, 1967). The solution to these equations can be obtained once the initial and boundary conditions are specified. The initial conditions require that the temperature and degree of cure inside the composite be given before the start of the cure. The boundary condition requires that the temperatures on composite surfaces in contact with the tool be known as a function of time during cure. Therefore the boundary condition is related to the specified cure cycle and the equipment setup. The objective for the cure control scheme is to achieve the desired composite quality. Some of the main targets are reasonable temperature distribution, complete consolidation, minimum thermal stress and minimum void content. With a developed numerical scheme, the temperature distribution inside the laminate is calculated as a function of position and time. A good cure scheme should realize the two main targets: (a) the temperature is reasonably uniform inside the material and (b) the temperature does not exceed a preselected maximum at any time. For a given cure temperature and cure pressure, the time window for the consolidation is then specified. From the consolidation models, the compaction status of the consolidated composite can be obtained. In Springer’s model, the result is the total number of compacted plies, while in Gutowski’s model the result is the V ,distribution across the layers. If the consolidation cannot be completed with the selected cure cycle, proper modifications are then made. The compaction issue becomes crucial to the cure process of the thick composite structure. A multiple stage heating process may be designed to defer the cure reaction of the resin and thus prolong the consolidation time window. Voids within the composite material are harmful to its mechanical Performance. Experimental study shows that the resin

pressure early in the cure cycle and the initial resin moisture are crucial considerations in producing void-free laminates (Kardos et al., 1983, 1988). Since the driving force for diffusion rises with temperature, in order to prevent the potential for pure water void growth by moisture diffusion in a laminate at all times and temperatures during the curing cycle, the resin pressure at any point within the curing laminate must be higher than the minimum resin pressure required, which is a function of the relative humidity and temperature (Dave et al., 1990).Figure 26.12 shows a void stability map for pure water void formation in epoxy matrices. A similar pressure requirement also holds for small air/water voids after an initial growth period. It has also been observed that the void content is reduced (1 ATM I 101 kPI)

(RH), = 1ooO/o

300

400 1,K

(RH),= 50%

500

Fig. 26.12 Void stability map for pure water void formation in epoxy matrices (Dave et al., 1990).

Efects of tooling and part shape 589 significantlywhen the applied pressure is sufficiently high to collapse the vapor bubble before the gel point is reached. Therefore, after the time-temperature cycle is determined, it is possible to obtain a profile of the minimum pressure versus cure time. The boundary pressure is then maintained greater than the minimum pressure throughout the cure cycle. During the cooling stage after the cure of the composites, residual thermal stress is related to the difference between the cure temperature and ambient temperature, and the thermal expansion behavior of the composite material. For a laminated structure, calculation of the thermal stress has been discussed and formulated by Tsai and Hahn (1980). Since the material shows viscoelastic behavior, stress relaxation has been observed over time. A post-cure process is usually applied to the structure to relieve the induced thermal stress. For large complex-shaped composite structures, non-autoclave curing methods are used. Compared with traditional autoclave curing methods, the component size restrictions are eliminated, energy consumption is reduced, and capital equipment cost can be cut down. The non-autoclave processes use an oven, integrally reinforced tools, and presses. Major issues related to non-autoclave curing are the effective compaction of the composite plies, and the elimination of the trapped interlaminar or intraply air. 26.4 EFFECTS OF TOOLING AND PART SHAPE

Properly designed tools that produce acceptable parts on a reproducible basis are a must when fabricating composite structures. The tool design requires the consideration of as many factors as are studied in the design of the part itself. The main requirement for the tools is to maintain proper geometric dimensional stability and surface profile during the compression and thermal cycling processes. On the other hand, the tool must also be

heated to a specified temperature at a specified rate under controlled conditions in the autoclave. Tooling materials may be metal (steel, nickel, nickel alloys, and aluminum), graphite-epoxy and elastomer, depending on different composite part shape, size, volume of production and curing method. Selection of the tooling material often reflects a compromise among these considerations. Thermal behavior of the tooling material is also crucial in the design and fabrication. Table 26.1 lists the coefficient of thermal expansion of different composite and tooling materials. The values for the composites are dependent on the ply orientation and fiber volume fraction, and typical values are shown there.

Table 26.1 Coefficient of thermal expansion (CTE) for various materials (Borstell and Turner, 1987)

Material

CTE ( I P / K )

Structural composite material Boron-poxy Aramid+poxy Graphiteepoxy Fiberglass-epoxy

3.6-10.8 -2.0-5.8 1.8-9.0 7.2-9.0

Tooling material Graphiteepoxy Cast ceramic Tool steel Iron (electroformed) Nickel (electroformed) High-temperature cast epoxy Aluminum Silicone rubber

4.1-9.0 0.81 11.3 11.9 12.6 19.8 23.2 81-360

26.4.1 TOOLING FOR AUTOCLAVE MOLDING

The traditional autoclave molding process uses a vacuum bag to impose a pressure difference on the composite lay-up. A typical bagging system consists of the following steps (Schwartz, 1983).

590 Consolidation techniques and cure control 1. Cover the lay-up with a perforated parting film or separator. Then lay up a layer or layers of bleeder material. The requirement of the bleeder layers should be such as to ensure adequate bleeding of air and excess resin out of the part. 2. Place a strip of jute (vent material) just beyond the edge of the lay-up and put bagsealing compound along the outside perimeter. 3. Cover the lay-up, jute, and sealing compound with a flexible-film diaphragm and seal the diaphragm to the mold with the seal compound. 4. Connect the vacuum lines and slowly apply the vacuum pressure while working the wrinkles and excess air out of the lay-up, bleeder material, and vacuum bag. 5. Check system for vacuum leaks. 6. Keep the part under vacuum while it is waiting to be cured in the oven or autoclave.

Graphite-epoxy laminate

Mold form

-/

cdp\

Mold half-

Angle caul plate

Caul plate stop

{Resin reservoir

.Mold half

To prevent surface irregularities on the bag side (untooled surface) of the parts, a caul plate may be used. The sole purpose of caul plates is to improve the visual appearance of I the parts. They do not control part thickness. A flexible caul plate with a thermally stable rub\Cao ber such as silicone or a fluoroelastomer is often used to accommodate the surface geom- Fig. 26.13 Example of autoclave tooling (Borstell and Turner, 1987). etry. Figure 26.13 shows examples of autoclave tooling setups with caul plates. The three issues related to the tooling introducing a thermal strain. As the part and design (Borstell and Turner, 1987) are thermal tool cool down from the gel temperature, the expansion correction, coordinating the loca- tool usually shrinks more than the part. As an alternative, graphiteepoxy molds are used in tion of partial plies and use of caul plates. Because of the low coefficients of thermal some applications. Although some data has expansion of composites when compared with been published, not all composite materials metal tooling materials, thermal strain or have been measured. One empirical method is stress must be considered for a curing process. to cure a representative panel on a plate of the In the autoclave, the temperature at which the specified tooling material using the specified resin solidifies is the gel temperature. At that cure cycle. Corrections can be estimated by specific temperature, the part is the same size comparing the difference between the mold as the thermally expanded mold. At a temper- and part dimensions. Another recommended ature above the gel temperature, the tool empirical correction method is to correct steel expands more than the partially cured part or nickel tools by making the tool 0.999 of the

Effects of tooling and part shape 591 engineering dimension, and to correct aluminum tools by 0.998. For example, a 2540 mm (100 in) dimension is tooled to be 2537 mm (99.9 in) for the steel tool. These corrections are needed to ensure an acceptable fit of mating composite parts. Most parts contain partial plies to accommodate local areas of increased stress. Several techniques are used to control the location of partial plies, including polyester film templates, slotted templates, and rails and banking surfaces. These tools serve as supplemental guidance to position the partial plies in the lay-up process. Typical cases of applying a caul plate are to control the edge of a panel or the flanges of channels. The design of the metal caul plates must take into account the fact that the matrix resin melts in the autoclave to a very low viscosity. The caul plate performs by pushing excess resin sideways. Thus the rigid metal caul plates must have high rigidity so that they do not deflect under autoclave pressure at curing temperature. The thickness of the caul plates can be calculated by use of the equations for unsupported bending beam analysis. The deflection of the caul plate can be estimated using the balance condition of resin pressure and applied force (Gutowski and Cai, 1988).The caul plate deflections should be limited to half the tolerance permitted in the part.

aluminum. During autoclave curing of composite parts, the thermal uniformity is excellent with rapid heat-up and cool-down rates. It is easy to handle and transport because of its light weight. It also offers outstanding durability because the mold surface resists cutting or impact damage and is not thermally degraded. When damaged, it is easy to repair by welding, soldering, silver-soldering, or selective plating. It can provide complex contours without expensive machining. With most resin systems, it shows good release properties. Figure 26.14 shows the procedures of making an electroformed nickel tool. As in some other types of tooling, constructing a model of the part surface is the first step in creating an electroformed mold. The models are the same net dimensions as the required nickel mold. Compensation may be required when the coefficient of thermal expansion of the composite part differs greatly from that of the nickel mold. Models are made from plaster, epoxyfaced plaster, fiberglass, fiberglass-epoxy, wood or other materials. From the model a reverse mandrel 'splash' is generally fabricated from epoxy-faced fiberglass or plaster. The mandrel to be used in electroforming is then copied from the 'splash', although the model can be used as the mandrel if it is prepared correctly. The comers of the mandrel should be designed to have radii in excess of 0.76 mm (0.030 in) to avoid thin spots in the 26.4.2 ELECTROFORMED NICKEL TOOLING deposit. Draft and taper should be designed An electroformed nickel tool consists of a into the mandrel to facilitate its removal from 4.6-6.4 mm (0.18-0.25 in) thick electrode- the electroform. Sharp corners or narrow, deep posited mold surface that is supported by a grooves should be avoided if possible. The simple steel substructure. The mold surface is mandrel can be fabricated from epoxy-faced produced by the electroplating process fiberglass, rubber, or other materials. The surface of the mandrel is made conductive by (Sheldon, 1987). The electroformed tooling concept offers proper coatings. The back of the mandrel must numerous advantages. The size of the mold is be reinforced to keep the mandrel from disrestricted only by the size of the electroform- torting during the electroforming process. Electroforming is the process of producing ing tank. The cost of producing duplicated tools is low. The mold surface is very smooth an article by electrodeposition of a metal onto and scratch resistant. The coefficient of ther- a conductive mandrel surface. An anode susmal expansion is approximately 40% less than pended in an aqueous electrolyte is connected


---t

Splash

Model

-

Fiberglass plating mandrel

Mold electroformed

I

I

i

Plated mold and tool upport structure

Mold and structure joined

Plating mandrel removed

Fig. 26.14 Example of electroformed nickel tooling (Sheldon, 1987).

to the positive pole of a DC electric source, and parts. These include low coefficient of thermal the mandrel (cathode) is connected to its neg- expansion, ease of preparation, low density, ative pole. The flow of electricity or electrons and thermal stability (Harmon, 1987). Their results in the oxidation of a nickel anode to disadvantage is that they are less durable than nickel ions and the reduction of nickel ions to metal tools. Composite tool making starts with a master nickel metal at the cathode (mandrel). The typical rate of growth is approximately model, usually built with plaster or hard0.013-0.025 mm (0.0005-0.001 in) per hour. wood. The master models require proper When the electroform is removed from the drying, sealing, and coating with mold release. mandrel, its surface is a mirror image of the Then lay-up can be done directly on the plaster surface of the mandrel. A natural physical or wood master. Liquid gel coats are required characteristic of electrodeposition is that elec- to obtain a high fidelity surface on tools cured tric current will tend to localize the deposit on by the vacuum bag process which does not all edges and corners, causing an uneven generate enough pressure to ensure a void-free thickness on the electroform. However, there surface, but may not be required on tools are a variety of techniques to offset this effect. cured by the autoclave process which does After the desired mold thickness is provide sufficient positive pressure. Prepregs obtained, the mold is removed from the tank, with light weight fabrics are used directly cleaned and the steel back-up structure is against the tool surface, while prepregs with attached. The nickel mold is then polished to heavier fabrics are used to build up the thickness. During the lay-up, care should be taken the required finish, and ready for use. to work each ply into all radii and corners and to remove all entrapped air. Debulking is 26.4.3 GRAPHITE-EPOXY TOOLING applied after the lay-up, either with a vacuum Composite tools have definite advantages bag setup or with -assistance of an autoclave over metal molds for large or highly contoured for a pressure debulk, to consolidate the plies

Eflects of tooling and part shape 593 and remove all entrapped air. The curing process is done with a vacuum bagging system or with an autoclave. With the tool still on the model, the support structure, either a solid laminate or an ’egg-crate’ panel is attached to the tool by means of locally applied fabrics, room-temperature curing, and high-temperature resistant resins. Once the support structure is cured to the laminate shell, it is removed from the master. Care should be taken to avoid damaging either the tool or the master. Figure 26.15 illustrates the graphite-epoxy tooling making process. Compositetools are being used successfully throughout the aerospace industry to produce parts that are structurally reliable, reproducible, and dimensionally accurate.

In thermal expansion molding, two basic methods are employed: the trapped or fixedvolume rubber method and the variable-volume rubber method. Figure 26.16 shows the setup for both methods. The fixedvolume method exploits the large difference between the coefficient of thermal expansion of the elastomer and that of metals. The elastomer is confined within a closed metal tool

Rubber tool sized to fill the cavih, in the pan

,Pan Teflon separator film

Breather cloth

/

Vacuum bag

.---

Floating-plate pressure control

-

’

Rubber tool projects above the pan 30 excess pressure is vented by forcing the floating plate to the bag.

‘PFP

master

Fig. 26.15 Example of graphite-epoxy tooling

(Harmon, 1987).

M

,Outer

26.4.4 ELASTOMERIC TOOLING

Elastomeric tooling or rubber tooling can be used to generate molding pressure or to act as a pressure intensifier. In thermal expansion molding, elastomeric tooling is constrained within a rigid frame to generate consolidation pressure by thermal expansion during the curing cycle (Foston and Adams, 1987).

box

w

Fig. 26.16 Example of elastomeric tooling (Foston

and Adams, 1987) (a) fixed volume method; (b) variable volume method.

594 Consolidation techniques and cure control Impregnated Composites, Proc. 9th Int. Cod. cavity. When heated, it expands into the cavity, Composite Mater. (ICCM-9), 1993, 3,575-583. exerting the pressure required to compact a composite laminate. The variable-volume Dave, R.S., Kardos, J.L. and Dudukovic, M.P., A Model for Resin Flow During Composite method offers more flexibility and control than Processing, Part 1: General Mathematical the fixed-volume method because a precisely Development, Poly. Composites, 1987, 8(1), calculated volume of rubber is not normally 29-38. required. In most applications, the rubber is Dave, R.S., Kardos, J.L. and Dudukovic, M.P., A Model for Resin Flow During Composite simply 'set back' to allow for the bulk factor of Processing, Part 2: Numerical Analysis for the molding material during assembly of the Unidirectional Graphite/Epoxy Laminates, tooling details. A floating plate is used for the Poly. Composites, 1987,8(2), 123-132. pressure control. Dave, R.S., Mallow, A., Kardos, J.L. and Dudukovic, Thermal expansion molding with elasM.P., Science-based Guidelines for the tomeric tooling has been successfully used on Autoclave Process for Composites commercial aircraft parts such as rudders and Manufacturing, SAMPE I., 1990,26(3),31-38. spoilers (Schneider and Carroll, 1987). This Dusi, M.R., Lee, W.I., Ciriscioli, P.R., and Springer, G.S., Cure Kinetics and Viscosity of Fiberite 976 reduces the number of detail parts fabricated Resin, J. Composite Mater., 1987,21(3),243-261. and the need for bonding and mechanical fasFoston, M. and Adams, R.C., Elastomeric Tooling, tening on assembly, thereby effecting in Engineered Materials Handbook, Vol. 1: significant reductions in production time and Composites, ASM International, 1987, pp. cost. 590-594. Gibson, R.E. and Hussey, M.J.L., The Theory of One-Dimensional Consolidation of Saturated REFERENCES Clays, Geotechnique, 1967,17,261-273. Batch, G.L. and Macosko, C.W., A Model for Two- Gutowski, T.G., A Resin Flow /Fiber Deformation Model for Composites, S A M P E Quarterly, 1985, Stage Fiber Deformation in Composite 16(4),58-64. Processing, Proc. 20th Intern. SAMPE Tech. Gutowski, T.G., Morigaki, T. and Cai, Z., The Conf., September 1988, pp. 641-650. Consolidation of Laminate Composites, J. Bhi, S.T., Hansen, R.S., Wilson, B.A., Calius, E.P., Composite Mater., 1987,21, 172-188. and Springer, G.S., Degree of Cure and Viscosity of Hercules HBRF-55 Resin, Proc. 32nd Intern. Gutowski, T.G., Cai, Z., Bauer, S., Boucher, D., Kingery, J. and Wineman, S., Consolidation SAMPE Symp. Exhib., Vol. 32., 1987, pp. Experiments for Laminate Composites, J. 1114-1118. Composite Mater., 1987,21,650-669. Biot, M.A., General Theory of Three-Dimensional Gutowski, T.G. and Cai, Z., The Consolidation of Consolidation, J. Appl. Pkys., 1941,12, 155-164. Composites, in The Manufacturing Science of Biot, M.A., Theory of Elasticity and Consolidation Composites, Proc. Manufacturing International for a Porous Anisotropic Solid, J. Appl. Phys., 88, Vol. IV,(ed T.G. Gutowski), 1988, pp.13-25. 1955,26(2), 182-185. Biot, M.A. , General Solutions of the Equations of Halpin, J.C., Kardos, J.L. and Dudukovic, M.P., Elasticity and Consolidation for a Porous Processing Science: An Approach for Prepreg Material, J. Appl. Meck., 1956, March, 91-96. Composite Systems, Pure Appl. Chem., 1983,55(5). Borstell, H. and Turner, K.T., Tooling for Autoclave Harmon, B.D., Graphite-Epoxy Tooling, in Molding, in Engineered Materials Handbook, Vol. Engineered Materials Handbook, Vol. 1: Composites, 1: Composites, ASM International, 1987, pp. ASM International, 1987, pp.586-589. 578-581. Kardos, J.L., Dudukovic, M.P., McKague, E.L. and Lehman, M.W., Void Formation and Transport Cai, Z . and Gutowski, T.G., Fiber Distribution and Resin Flow in the Molding Process, Proc. 7th During Composite Laminate Processing: An Initial Model Framework, in Composite Int. Conf. Composite Mater. (ICCM-7), 1989, 1, Materials: Quality Assurance and Processing, 76-82. ASTM STP 797, (ed C.E. Browning), 1983, pp. Connor, M., Gibson, A.G., Toll, S. and Manson, J.A.E., A Consolidation Model for Powder 96-109.

References 595 Kardos, J.L., Dave, R. and Dudukovic, M.P., Voids in Composites, in The Manufacturing Science of Composites, Proc. Manufacturing International '88, Vol. IV,(ed T.G. Gutowski), 1988, pp. 4148. Kim, T.W., Yoon, K.J., Jun, E.J. and Lee, W.I., Compaction Behavior of Composite Laminates During Cure, SAMPE I., 1988,24 (S), 33-36. Kim, T.W., Jun, E.J. and Lee, W.I., Compaction Behavior of Thick Composite Laminates During Cure, Proc. 34th Inter. SAMPE Symp., 1989, 12 (l),17-19. Kim, Y.R., McCarthy, S.P. and Fanucci, J.P., Compressibility and Relaxation of Fiber Reinforcements During Composite Processing, Polym. Composites, 1991,12 (l),13-19. Lam, R.C. and Kardos, J.L., The Permeability of Aligned and Cross-Plied Fiber Beds During Processing of Continuous Fiber Composites, Proc. Am. SOC.Composites, Third Technical Conf., Seattle, WA, 1988, pp. 3-11. Lam, R.C. and Kardos, J.L., The Permeability and Compressibility of Aligned and Cross-Plied Carbon Fiber Beds During Processing of Composites, Proc. 47th Ann. Tech. Conf. (ANTEC'89), SPE, New York, 1989, pp. 1408-1412. Lee, W.I., Loos, A.C., and Springer, G.S., Heat of Reaction, Degree of Cure, and Viscosity of Hercules 3501-6 Resin, J. Composite Mater., November 1982,16, pp. 510-520. Lee, S.Y. and Springer, G.S., Effects of Cure on the Mechanical Properties of Composites, J. Composite Mater., 1988,22(1), 15-29. Lindt, J.T., Engineering Principles of the Formation of Epoxy Resin Composites, SAMPE Quarterly, October, 1982. Lindt, J.T., Consolidation of Circular Cylinders in a Newtonian Fluid, I. Simple Cubic Configuration, J. Rheology, 1986,30. Loos, A.C. and Freeman, Jr., W.T., Resin Flow During Autoclave Cure of Graphite-Epoxy Composites, High Modulus Fiber Composites in Ground Transportation and High Volume Applications, ASTh4 STP 873, (ed D.W. Wilson), 1985, pp. 119-130. Loos, A.C. and Springer, G.S., Curing of Epoxy Matrix Composites, J. Composite Mater., 1983,17, 135-1 69. Loos, A.C. and Springer, G.S., Calculation of Cure Process Variables During Cure of Graphite/Epoxy Composites, Composite Materials: Quality Assurance and Processing, ASTM STP 797, (Ed. C.E. Browning), 1983, pp.

110-118. Loos, A.C. and Springer, G.S., Curing of Graphite/Epoxy Composites, Air Force Materials Laboratory Report AFWAL-TR-834040, Wright Aeronautical Laboratories, Wright Patterson Air Force Base, Dayton, OH, 1983. Roylance, D., Reaction Kinetics for Thermoset Resins, in The Manufacturing Science of Composites, Proc. Manufacturing International'88, Vol. IV, (ed T.G. Gutowski), 1988, pp. 7-11. Schneider, C.W. and Carroll, H.E., Elastomeric Tooling Application, in Engineered Materials Handbook, Vol. 1: Composites, ASM International, 1987, pp. 595-601. Schwartz, M.M., Composite Materials Handbook, McGraw-Hill, 1983. Sheldon, D.L., Electroformed Nickel Tooling, in Engineered Materials Handbook, Vol. 1: Composites, ASM International, 1987, pp. 582-585. Smith, G.D. and Poursartip, A., Comparison of Two Resin Flow Models for Laminate Processing, J. Composite Mater., 1993,27(17),16951711. Springer, G.S. and Tsai, S.W., Thermal Conductivities of Unidirectional Materials, J. Composite Mater., 1967,1, 166-173. Springer, G.S., Resin Flow during the Cure of Fiber Reinforced Composites, J. Composite Mater., 1982,16,400410. Springer, G.S., Modeling of the Cure Process of Composites, SAMPE J., September/October 1986, pp. 22-27. Tang, J.M., Lee, W.I. and Springer, G.S., Effects of Cure Pressure on Resin Flow, Voids, and Mechanical Properties, J. Composite Mater., 1987, 21,421440. Tsal, S.W. and Hahn, H.T., Introduction to Composite Materials, Technomic Publishing, 1980. Van Den Brekel, L.D., and De Long, E.J., Hydrodynamics in Packed Textile Beds, Textile Research J., August, 1989, pp. 433-440. van Wyk, C.M., Note on the Compressibility of Wool, J. Textile lnst., 1946, 37, T285-T292. Williams, J.G., Morris, C.E.M. and E d s , B.C., Liquid Flow through Aligned Fiber Beds, Polym. Engng Sci., 1974,14 (6), 413-419.

COMPOSITE MACHINING

27

Kent E. Kokkonen and Nitin Potdar

27.1 INTRODUCTION

27.2 CONVENTIONAL MILLING

The processes used to manufacture composite When milling graphite-epoxy with polycrysstructures generally require that trimming and talline diamond (PCD) the chips are formed as other machining operations be performed small particles of powder dust and fumes. The prior to assembly. Machining processes are surface roughness is a function of fiber orienrequired to produce accurate surfaces and tation, cutting direction and the angle between holes to allow precision fitting of components cutting direction and fiber direction. The surinto an assembly. Due to shrinkage during the face may sometimes exhibit many small holes curing stage of the composite structure it is not due to fiber pull out. When taking heavy practicable to place holes in the part during milling cuts there is a greater tendency to the molding stage, therefore milling, cutting, break comers as the tool exits the material so it drilling etc. are considered a post cure opera- is advisable to first machine a step on the edge tion. perpendicular to the final pass. A four fluted Due to the toughness and abrasive nature of end mill will reduce cutting pressure on the modern composites, there is a need for harder laminate and keep it cooler. Climb milling and longer lasting cutting tools. A large data- helps prevent the fibers from separating from base of machining information for various the matrix bond material. high speed steel and carbide cutting tool Advantages of machining composites are: materials exists for machining metal, wood and some thermoplastics. However, much of 0 improved surface finish unless part surface was directly in contact with the mold surthis data cannot be applied to machining modface; ern composites. Modern composites like 0 machined surfaces provide accurate mating graphite-epoxy, aramid-epoxy and carbonsurfaces for parts to be assembled; carbon each have their own machining charac0 eliminates the majority of the problems teristics. Composites are not homogeneous or associated with part shrinkage and insert isotropic, therefore the machining characterismovement during the fabrication processes. tics are dependent on the tool path in relation to the direction of the reinforcing fibers. Tool life factors are: Metals or metal alloys have nearly homogeneous properties throughout the workpiece, 0 PCD end milling cutters will perform sixty to one hundred times longer than carbide; but each material in a composite retains its 0 cutting speed does not have a great effect individual properties. on the flank wear of PCD cutting tools. With increased cutting speeds, the feedrates can be increased and machining time Handbook of Composites. Edited by S.T. Peters. Published decreased; in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

Mechanical drilling of composite materials 597

0

0

0

cutting speeds range from 244 surface m/min (800 surface ft/min) to 762 surface m/min (2500 surface ft/min) with PCD end mills; when cutting parallel to the fiber direction, the wear ratio on the cutting tool increases compared with cutting perpendicularly to the fiber direction; surface finish remains below 20Ra [arithmetical average roughness (see IS0 R488)] when cutting with PCD end mills and the flank wear is approximately 0.127 mm (0.005in); the surface finish deteriorates above 150 Ra when cutting with a carbide end mill and the flank wear has reached 0.127 mm (0.005in); roughing feedrates range from 0.23 mm/rev (0.007 in/rev) to 0.38 mm/rev (0.012 in/rev) and finish feedrates range from 0.076 mm/rev (0.002 in/rev) to 0.13 mm/rev (0.005 in/rev); the depth of cuts should range from one quarter to one half of the diameter of the end mill cutter. Depth of cut will vary depending on the rigidity of machine ways, spindle and workholding devices.

The disadvantages associated with milling of composites include controlling the graphite chips (dust particles), confining them to a small area and having an adequate collection system. A second problem is controlling the outer layers of the composite so that the fibers will shear instead of lifting up under the force of the cutting action and leaving extended fibers beyond the cut surface. Also when cutting perpendicular to the lay of composite fibers, edge break-out can occur. This can be controlled by designing a backup structure in the tooling. 27.3 CONVENTIONAL TURNING

The turning of graphite composite is utilized to produce round surfaces that need to mate with either metal of graphite parts. The cutting speeds can be over 305 m/min (1000 ft/min) if the part can be held securely and PCD tool inserts are utilized.

Depth of cut will vary depending on the thickness of the part and the amount of material to be removed. 27.3.1 ADVANTAGES

Computer numerical controlled lathes (CNC) can be used to machine simple to very complex rotational parts. CNC machining produces accurate parts at a high production rate. 27.3.2 DISADVANTAGES

Delamination can also occur on a lathe (Fig. 27.1), therefore the part may require a finish cut moving from the largest diameter to the smaller diameter. Graphite chips are a serious problem. The spinning chuck creates a fan effect on the graphite particles. The exhaust system must be adequate to control the graphite chips. Also, the machine ways and the ball screws on the machine must have sealed protection to minimize wear. The computer control requires protection from the graphite chip particles. 27.4 MECHANICAL DRILLING OF COMPOSITE MATERIALS

Drilling holes in composites can cause failures that are different from those encountered when drilling metals. Delamination, fracture, break-out and separation are some of the most common failures. Delamination (surface and internal) is the major concern during drilling composite laminates as it reduces the structural integrity, results in poor assembly tolerance, adds a potential for long term performance deterioration and may occur at both the entrance and exit plane. Delamination can be overcome by finding optimal thrust force (minimum force above which delamination is initiated). Figure 27.2 shows push out delamination at exit because at a certain point loading exceeds the interlaminar bond strength and delamination occurs. Figure 27.3 shows peel-up delamination at entrance

598 Composite machining

Fig. 27.2 Drill bit showing push-out delamination at exit.

graphite reinforcement materials. Each of these materials requires individual attention in the selection of cutting tool parameters. The composite materials with metal backup panels require separate drills with different geometries. Cutting speeds and feedrates vary in each of the various combinations of materials. Secondary drilling or reaming operations are required to hold tight tolerances or smooth surface finishes on the holes. Table 27.1 shows Fig. 27.1 Machining direction for turning compos- the drilling results when using four styles of drills.' ite parts on a lathe. PCD tooling offers increased tool life, better hole quality, consistent hole size and higher because the drill first abraded the laminate machining rates. Drilling and countersinking and then pulled the abraded material away along the flute causing the material to spiral up before being machined completely. This type of delamination decreases as drilling proceeds since the thickness resisting the lamina bending becomes greater. Among the variables to be considered for tool selection include the thickness of material, diameter of hole, tolerance requirements, hole finish requirements and the composite materPeeling ial being drilled. Tungsten carbide, micrograin Action tungsten carbide and drill tool materials are used for drilling composite materials. Some commonly used composites are I glass-epoxy, glass-graphite-epoxy, graphiteepoxy, graphite-epoxy with aluminum backup and graphite-epoxy with titanium backup. Other materials include the aramids Fig. 27.3 Drill bit showing peel-up delaminationat (Kevlar@) with combinations of glass or entrance.

I

I

4 I

I I

Mechanical drilling of composite materials 599 Table 27.1 Summary of drill performance: mean hole quality measures as a function of point angle. Maximum recorded values of response parameters are shown in box brackets, [I (Reproduced from Ref 1 by permission of ASM Materials Week)

Criterion/drill

Dagger

8-Facet

4-Facet Master

NAS 907-1HSS

Exit breakout (Rank least = 1)

1

2

3

4

Panel damage, D,

1.96 (3.34)

2.37 (3.18)

2.75 (3.62)

3.63 (5.54)

1

2

3

4

Thrust force, N (1bf)

114 [166] (25.6 [37.4])

201 [378] (45.3 [85.2])

263 [428] (59.3 [96.3])

593 [969] (133.5 [218])

Torque, Nm (ft lbs)

1.29 [2.18] (0.95 [1.61])

1.15 [2.0] (0.85 [1.5])

0.7 [1.64] (0.50 [1.21])

1.53 [2.2] (1.13[1.61])

0.4 [1.6] (26 [641)

0.95 [2.2] (38 [88l)

1.6 [3.0] (64L1.221)

2.4 [4.12] (96 11651)

6.354 [6.379] (0.25016 [0.25115])

6.356 [6.369] (0.25022 [0.25075])

6.367 [6.395] (0.25067 [0.251751)

6.375 [6.397] (0.2510 [0.25185])

Microcrack density (Rank: lowest = 1)

Surface finish, R,, Pm (Pin.) Hole diameter, mm (in)

-

Hole out-of-roundness, (in)

Drill point angle, deg.

0.0061 [0.025] 0.003 [0.005] 0.0043 [0.018] 0.013 [0.03] (0.00024 [0.00101) (0.00012 [0.0002]) (0.00017 [0.0007]) (0.00051 [0.0012]) 30

with a combination tool provides better hole quality. Tool life is normally determined by the extent of delamination and fiber break out. For machining graphite composites with or without aluminum backing, PCD tooling is suggested with the same speeds and feeds used for machining graphite composites without any backing. For machining graphite composites with titanium backing, it is not recommended that the same drill be used for both the titanium and graphite sections. Initially a hole should be drilled up to the titanium layer with a hydraulic depth sensing device at high speeds and feed. A second drill with lower speed and feed for machining titanium should be used. Finally finish reaming operation and countersinking should be performed for assuring hole quality. A study carried out on carbon fiber-epoxy

24,118

140

135

(CFRP) and glass fiber-epoxy (GFRP) laminates using HSS and carbide tipped drills made the following observations. Both chisel edge and flank wear increased on the carbide drill with a higher ratio of wear between 200 and 400 holes (test sample 400 holes). The tool wear was greater in the CFRP laminates due to the abrasive nature of carbon fibers. Flank wear is more pronounced in GFRP when the feed was increased and the same effect is noted when speed is increased. The HSS drills lasted for ten holes in the graphite and twenty holes in the glass. 27.4.1 DFULL GEOMETRY

Drill point geometries influence the torque requirements. Lip relief and rake angles are determined by the application. The dagger drill is ideal to machine graphite composites

600 Composite machining as it eliminates breakout when exiting the workpiece. The dagger drill has 35" included point angle and a 121" chisel edge angle. Twist drills with flute configuration to control metal chips are also used. Fully fluted drills with PCD tips brazed on a solid carbide shaft provide the strength of carbide and hardness of diamond. Drill geometries are continuing to be experimented with to find ways to eliminate the problems associated with the hole making process in composites. Drill cutting parameters are: 0 0

0

feedrates range from 0.025 mm/rev (0.001 in/rev) to 0.063 (0.0025 in/rev); cutting speeds range from 30 surface m/min (100 surface ft/min) to 460 surface m/min (1500 surface ft/min); high cutting speeds can burn the matrix material and reduce bond strength between the composite material and the matrix material.

27.4.2 COOLANTS

A water soluble coolant forced through a cold air blast unit is recommended when machining most composite materials. However if the composite is hydrophilic in nature then a cold air blast unit in combination with dust or vacuum collection system should be used.

be processed. The grinding of polymer matrix composites (PMC) has a number problems. For example in the case of thermoplastic matrix, the surface of grinder becomes covered with melted thermoplastic. In the case of aramid fiber it is hard to get a clean cut surface because the grains cannot abrade the aramid fibers cleanly. Abrasive belts have been used on aramids with some success but dust collection has been a major problem. 27.6 MACHINING O F KEVLAR

Cutting, Trimming, Turning and Milling of Kevlar Because of its inherent toughness, Kevlar is difficult to cut, so sharp, heavy duty upholstery scissors will cut up to 170 g/m2 (5 oz/yd2)fabric of Kevlar. Woven roving and heavier fabrics can be cut using specially designed serrated scissors. An overview of cutting and trimming techniques and applications is shown in Table 27.2. For more information on cutting and machining of Kevlar refer to DuPont's Machining Handbook2. 27.7 ABRASIVE WATER JET MACHINING

Abrasive water jet (AWJ)is used for linear profile cutting, turning, milling and drilling operations in composite materials. 27.5 GRINDING COMPOSITE MATERIALS Conventional tool machining is affected by The grinding process has been used exten- fiber or particle reinforcements rather than the sively for finishing composite golf shafts and matrix material while AWJ machining is not. fishing rods. Five hundred parts per hour can To make a circular hole 6.35mm (0.25in) in be produced on centerless grinders. Silicon diameter in aramid 3.18 mm (0.125in) thick, it carbide wheels are used with an open grain to takes about the same time for both convenreduce wheel galling. Surface speeds between tional as well as AWJ machining. The cutting 1219 surface m/min (4000 surface ft/min) and process parameters for AWJ include water jet 1829 surface m/min (6500 surface ft/min) can pressure, velocity, abrasive grain size, abrasive be achieved. This equipment is specially material, standoff distance and jet impingedesigned for grinding and finishing compos- ment angle. and some additional parameters. ites. Grinding accuraces within 0.0127 mm Water jets without abrasive are also used for (0.0005 in) can be achieved with centerless cutting soft composites. Figure 27.4 shows the grinding. Both straight and tapered shafts can AWJ processes and machining parameters3.

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Abrasive water jet machining 601

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Abrasive water jet machining 603 27.7.1 LINEAR CUTTING WITH AWJ

27.7.2 TURNING

Linear cutting is used to trim composite parts and to cut profile shapes on the inside of a part. The cut surface is normally smoother near the entrance surface then becomes wavy in the lower half of the cut toward the exit surface. In general, the composite material is sheared away by a high velocity abrasive grain. The width of cut (kerf) decreases as the feedrate increases and the waviness increases as the feedrate increases3. Table 27.3 shows some of the observations made by Hashish3.The maximum cutting traverse rate is primarily controlled by the matrix material. Table 27.4 shows results for some composites with different speeds.

In turning with AWJ, a workpiece is usually rotated while the jet is fed along all three axes. The material encountered by the jet is abraded away in the form of a very fine debris. Higher jet pressure produces a smoother surface with a higher material removal rate. Higher traverse rates combined with multiple passes are more efficient than deeper cuts with lower traverse rates. Surface finish is affected by unsteadiness in traverse rate or abrasive flow rate. The repeatability and accuracy of the AWJ turning process depends on control and steadiness. A 10% variation in rotational speed does not affect the surface waviness but a traverse rate variation over 4% will significantly affect the surface waviness. Some methods to improve surface finish are:

Table 27.3 Typical through-cutting traverse rates (in mm/s) with AWJs for different composites3 Material

Thickness (mm) 0.79

1.60

3.18

6.36

12.7

19.1

50.8

15 7.5 17 15 4.7

10 4 12 10 2.5

2.5 0.85 15 4.2 0.63

Organic matrix composites: Plastic and composites Carbon-carbon composites Epoxy-glass composites Graphite-epoxy composites Kevlar (steel reinforced)

53 42 105 74 42

38 32 95 63 25

29 22 76 52 17

21 13 42 40 8.5

Cutting conditions: p = 345 MPa, d, = 0.299 mm, d, = 0.762 mm, garnet mesh 80 Sic abrasives

Table 27.4 Surface waviness and corresponding cutting traverse rates (in mm/s) for some composite materials3 Material A1,OJSiC (20%) 6.35 mm Toughened zirconia (6.36 mm) Mg/B,C (15%) 6.36 I T U ~ Graphite+poxy composites (3.18mm thick) Graphite-epoxy composites (18.5 mm thick)

Rh4S surface waviness (pm) 1.9

2.5

-

0.15

-

0.6

3.8

5

6

8

-

0.29

-

-

-

-

0.2 3 4 0.85

Cutting conditions: p = 370 MPa, dn = 0.299 mm, d, 0.762 mm, garnet mesh 80

8 1.7

0.4 6 12 2.5

20 3.4

0.5 8 30 4.25

604 Composite machining 0

0 0

0

multipassing by traversing the jet without lateral feed; use of finer abrasive and increasing number of passes; to improve surface roughness, use softer abrasives like silica sand, copper slag etc; finishing with slurried abrasive yields improvement in surface roughness.

27.7.3 MILLING

The main objective of AWJ in milling is to produce a cavity with controlled depth. In this method, the jet material interaction is the depth determining factor. The production of kerf irregularity can be reduced by manipulating one of the factors, such as traverse rate, increasing the stand off distance or angling the jets. To mill square pockets the traverse speed can be varied rather than angling the water jet head. In this case the nozzle can be manipulated over the workpiece with an oscillatory drive using a motor and an eccentric. A uniform traverse rate and exposure time can produce a uniform depth cavity. A hard material pattern with the shape to be milled can be used to mask the target surface. This way the mask will allow jetting in the traverse zones where the traverse rate is uniform. Surface finish variations can be achieved by using different abrasive materials or grit sizes. Harder abrasives can be used for higher material removal rates and softer abrasives for finishing operations.

velocity decreases as the depth increases which can be attributed to the effect of return flow which reduces particle velocity and interferes with the impact process. Pressures of 3040 MPa are common for piercing glass. High pressures are necessary to pierce brittle or laminated composites. The higher pressures may cause the following problems: fracture due to shock loading of water; hydrocracking due to hole hydrodynamic pressurization; delamination due to loading. Holes larger than the piercing diameter of the AWJ are first pierced, then profile cut to the finished diameter being offset by the kerf amount. Hole shape variance depends on mixing tube length, target material, standoff distance, depth of hole and dwell time in the cut. Mixing tube length is important when drilling materials with high resistance. Increasing the mixing tube length improves the distribution of the abrasive with the water jet. This produces holes that are straighter and rounded. Advantages offered by AWJ are: 0 0 0 0 0 0

0 0

0

27.7.4 DRILLING

Hole drilling can be performed in any of the following ways depending on the diameter and accuracy of the holes: piercing is suitable for small diameter holes; kerf cutting is suitable for large diameter holes; milling is suitable for blind holes. Techniques of hole piercing vary for each composite material. Piercing glass, acrylic and polycarbonate show that the general geometrical features of pierced holes are similar. Particle

suitable for wide range of composites; can perform many operations like turning, drilling and milling; no thermal stresses; high as well as low material removal rates; no heavy clamping of workpieces required; omnidirectional machining; process can be automated; optimal range of parameters available to prevent delamination, loading and splintering; fine holes of 0.5 mm (0.012 in) can be drilled. Disadvantages:

0 0 0

0

dimensional accuracy is low; temperature rise in cutting region may be observed; limited data is available with respect to applications in metal and ceramic composites; not suitable for materials that are hydrophilic in nature.

Ultrasonic machining 605 Advantages:

27.8 LASER MACHINING OF COMPOSITES

Lasers are used in various industrial applications such as drilling, cutting, welding and heat treatment of metals, etc. In composites, polymer matrix materials are most suited for laser cutting. Laser cutting is a non-contact ablation process in which efficiency is determined by thermal properties of the workpiece material. Two types of laser have been used in industry: Nd-YAG solid state laser and CO, gas laser. The Nd-YAG laser operates in the near infrared (IR) region of the spectrum while CO, gas laser operates in the far infrared region. The Nd-YAG IR region wavelength is not absorbed by glass and many plastics while the CO, far IR region wavelength is. Applications of Nd-YAG solid state lasers extend from drilling fine holes in jet engines to welding implant devices for the medical industry. It has been determined that the NdYAG laser is very effective in cutting graphite-epoxy composite materials. The high power short pulses achieved with this laser vaporizes both the graphite and epoxy matrix before the epoxy resin can be overheated. The CO, gas laser applications extend from drilling holes in baby bottle nipples to welding automotive components in assembly lines. CO, lasers operate in either continuous wave or pulsed mode. Pulsed mode is preferred because of high powers obtained and cool down time. Aramid fiber reinforced plastic (AFRP) has been cut very effectively by the CO, lasers. The general characteristics of a laser cut zone in composite materials are shown in Fig. 27.5. The charred layer which includes a zone with fibers protruding from the matrix and as outer darkened zone in which the matrix has undergone some degradation4t5. Figure 27.6 shows the relationship between kerf width and cutting speed. For three dimensional (3D) machining two laser beams are directed through an optical assembly to intersect in the plane of work piece to cut shoulders and vee grooves.

0 0 0 0

0

superior quality edges due to high temperatures; vaporization of the material in cut zone; extremely localized action; sealing of the edge in the cut zone; pulsed CO, has been demonstrated as the best laser for processing Kevlar composites. Disadvantages:

0

0

0

beam divergence after its focal point; material thickness of about 9.5 mm (36 in) is the maximum thickness that can be cut with 1500 W; heat affected zone of varying dimensions.

27.9 ELECTRIC DISCHARGE MACHINING (EDMI

Advanced composites can be cut by EDM as there is no physical contact between the electrodes or workpiece and the tool. In order to EDM a composite, it should have an electrical resistivity of less than 1-3 ohm/m. Polymer matrix composite manufacturers can add a small amount of copper in the matrix of the product to allow shaping by EDM. EDM can be used with conductive silicides, borides, carbides, etc. The EDM process is more accurate than AWJ machining. Small holes of 0.25 mm (0.01 in) diameter can be drilled in SiC/TiB, composites. The EDM process is found to be slow for many production applications. 27.10 ULTRASONIC MACHINING

Ultrasonic machining (USM) incorporates a tool vibrating at 20 kHz and abrasive in a slurry to perform impact grinding of brittle materials. This technique is particularly useful for machining of ceramic matrix composites that are difficult to process by conventional methods. USM is a mechanical material removal process best suited for machining brittle materials like glass, ceramics, graphite and ceramic matrix composites. The process is limited to workpieces of size below 1OOmm


I

I

\ \

CHARRED LAYER

PROTRUDING FIBRES

\

I i

,i

\

,

'I

ICROSS SECTION

I

1

4 0 L-

beam exit side

Fig. 27.5 Schematic of FRP laser cut. (Reproduced by permission of Marcel Dekker Ltd.) W,: kerf width at the beam entry side; W,: kerf width at the

02 -

-

Fig. 27.6 Kerf width as a function of cutting speed for (0/90), laminates. (Reproduced by permission of Marcel Dekker Ltd.)

'.'a

'

m

u

'

)

Bo

80

Cutting speed (mm/s)

Irn

120

1

Ultrasonic machining 607 (3.94 in) because of the limitation on the size of the tool. Some of the variables that influence USM for close tolerances are as follows:

Abrasive type and size Abrasives contained in the slurry do the actual machining so they must be selected on the basis of the workpiece material and the surface quality needed. As in the case of AWJ, larger abrasive grains give higher material removal rates and rougher surfaces. The grain diameter cannot be larger than amplitude of the sonotrode as this would inhibit the injection of the grains to the machining gap. Common types of abrasive used are A1,0, oxide, Sic, BC and diamond. Table 27.5 shows recommended abrasive for various materials. The grain diameter affects surface roughness, overcut and machining rates. When high removal rates are necessary with no high surface quality required, 180-280 mesh abrasive do the job. For finer surface finish 320-600 mesh abrasive is recommended. Table 27.6 shows surface roughnesses for different workpiece materials.

Sonotrode (tool) material Tools with diamond tips have good material removal characteristics and very low wear but are difficult to machine. Table 27.7 shows accuracy results of using a non-rotating steel sonotrode. Ultrasonic vibrations The ideal condition would be the amplitude of ultrasonic vibration to be equal to the grain mean diameter. If the amplitude is too small the abrasive cannot enter the machining gap, if too large it causes the grains to be incorrectly projected. A mixture of both the types of abrasive may be used and a suitable amplitude selected to determine which size grain enters the machining gap. Surface area This factor influences removal rates and tool wear. With a small diameter, higher feed rate is obtained but also higher tool wear is noticed. This can be overcome by using a diamond tool or with a closed loop force sensitive

Table 27.5 Recommended abrasive for various materials6 Material

Recommended abrasive

Graphite Zirconia

Silicon carbide Silicon carbide or boron carbide Silicon carbide Boron carbide

Ceramic matric composites Metal matrix composites

Table 27.6 Surface roughness for various materials6 Workpiece material

Graphite Zirconia Ceramic matrix composites Metal matrix composites

Surface roughness Ra @ m) 1-2 0.75 0.70 0.90

608 Composite machining servo system maintaining accurate machining on prepreg materials like glass fiber, carbon pressures. Table 27.8 shows typical ultrasonic fiber and Kevlar with reduced fiber damage. Advantages: machining rates for a variety of materials6. USM is used in applications like drilling aerospace cooling holes in ceramic matrix 0 conductive and nonconductive materials can be machined; composite turbine blades, slotting, irregular 0 material hardness is not so important; configurations in ceramics and composites, machining of phased array radar components, 0 there are no chemical or electrical alterations in the workpiece; cutting tool inserts, superconductors, wire 0 3D and complex shapes can be machined draw dies and extrusion dies. A CNC USM can easily and quickly; cut through 6mm (0.24in) thick composite 0 no heat affected zone. layers and produce a controlled depth up to 50mm (1.97in). The latter is important, as Disadvantages: many composites have backing sheets that should not be damaged. The ultrasonic action 0 amplitude of ultrasonic vibrations are very important for proper machining; reduces the amount of force required to sever 0 limited sizes can be machined. the hard materials. This results in a better cut

Table 27.7 Accuracy results with a non-rotating steel sonotrode6

Material

Inlet diameter (mm)

Outlet diameter (mm)

Taper (Yo)

Roundness (mm)

Graphite

10.23-10.25 10.26-10.29

10.07-10.10 10.02-10.05

3.00 2.70

0.03" 0.03b

Metal matrix composite

10.20-10.24 10.09-10.12

8.87-9.92 8.85-9.90

9.00 6.60

0.04b 0.05b

Ceramic matrix composite

10.11-10.15 5.04

10.00 4.99

3.50 1.25

0.04b

5.05

4.85

5.50

-C

Zirconia

-c

Tool 1: Exponential,Diameter = 10 mm Tool 2: Exponential,Tube D = 10 mm, ID = 7 mm ' Tool 3: Exponential,Diameter = 5 mm a

Table 27.8 Typical ultrasonic machining rates for a variety materials7

Drilling diameter = 5 mm

Drilling diameter = 10 mm

Material

Time (min)

Removal rate (mm3/min)

Time (rnin)

Graphite Ceramic matrix composite Metal matrix composite Zirconia

1 3.5 10 210

164 39 7.6 0.65

1.25 5.6 14 90

Removal rate (mm3/min) 224 50 9.3 3.1

References 609 REFERENCES 1. Mehat, M., Reinhart, T.J. and Soni, A-H., Effect of fastener hole drilling anomalies on structural integrity of PMR-l5/GR composite laminates, Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Ill, 1-5 Nov. 1992. 2. Kevlar Cutting and Machining Handbook, E.I. Du Pont de Nemours and Co. 3. Hashish, M. State of the art of abrasive waterjet machining operations for composites. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. 4. Di Ilio, A., Tagliaferri, V. and Veniali, F. Machining parameters and cut quality in laser cutting of aramid fibre reinforced plastics. Materials and Manufacturing Processes, 1990,5(4), 591-608. 5. Lemma, S. and Sheehan, B. Laser Machining of Composite Materials. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. 6. Gilmore, R. Ultrasonic machining of composite materials, Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Ill. 1-5 November, 1992.

FURTHER READING Bhattacharyya, D., Allen, M.N. and Mander, S.J. Cryogenic Machining of Kevlar Composites. Materials and Manufacturing Processes, 1993,8(6), 631,651

Bhatnagar, N., Naik, N.K. and Ramakrishnan, N. Experimental investigations of drilling on CFRP composites. Materials & Manufacturing Processes, 1990, 5(4), 591-608 Geskin, E.S., Tisminetski, L., Verbitsky, D., Ossikou,V., Scotton, T. and Schmitt, T. Investigation of waterjet machining of composites. Proc. Machining of Composite Materials Symy., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. Hochegn, H., Puw, H.Y. and Yao, K.C. Experimental aspects of drilling of some fiber-reinforced plastics. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. Krishnamurthy, R., Santhanakrishnan, G. and Malhotra, S.K. Machining of polymeric composites. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. Lubin, G., ed., Handbook of Composites, 1982, New York: Van Nostrand Reinholt. Ramulu, M., Faridnia, M., Gargini, J. L. and Jorgensen, J. E. Machining of graphite/epoxy composite materials with polycrystalline diamond (PCD) tools. Trans. ASME, J. Engng Mater. and Tech., 1991,113, October . Zaring, K., Erichsen, G. and Burnham, C. Procedure optimization and hardware improvements in abrasive waterjet cutting systems. PYOC. Machining of Comr?osite Materials Svmp., ASM Materials" Week, Chicago, Ill., 1-5 "November 1992.

MECHANICAL FASTENING AND ADHESIVE BONDING

28

D. W. Oplinger

28.1 INTRODUCTION

It would be difficult to conceive of a structure that did not involve some type of joint. Joints often occur at a transition between a major composite part, where most of the structural performance is generated, and a metal feature, which is introduced to allow for very high localized bearing contact for which the composite has inadequate strength or durability. In aircraft such a situation is represented by articulated fittings on control surfaces as well as on wing and tail components which require the ability to pivot the element during various stages of operation. Tubular elements such as power shafting often use metal end fittings for connections to power sources or for articulation at points where changes in direction are needed. In addition, assembly of the structure from its constituent parts will involve either bonded or mechanically fastened joints or both. Joints represent one of the greatest challenges in the design of structures in general and in composite structures in particular. The reason for this is that joints entail interruptions of the geometry of the structure and often material discontinuities, which almost always produce local highly stressed areas, except for certain idealized types of adhesive joint such as scarf joints between similar materials. Stress concentrations in mechanically fastened joints


are particularly severe because the load transfer between elements of the joint have to take place over a fraction of the available area. For mechanically fastened joints in metal structures, local yielding, which has the effect of eliminating stress peaks as the load increases, can usually be depended on; such joints can be designed to some extent by the 'P over A' approach, i.e. by assuming that the load is evenly distributed over load bearing sections so that the total load (the 'I") divided by the available area (the 'A') represents the stress that controls the strength of the joint. In organic matrix composites, such a stress reduction effect is realized only to a minor extent, and stress peaks predicted to occur by elastic stress analysis have to be accounted for, especially for one-time monotonic loading. 28.2 MECHANICALLY FASTENED JOINTS COMPARED WITH ADHESIVE JOINTS

In principle, adhesive joints are structurally more efficient than mechanically fastened joints because they provide better opportunities for eliminating stress concentrations; for example, advantage can be taken of ductile response of the adhesive to reduce stress peaks. Mechanically fastened joints tend to use the available material inefficiently and are characterized by sizeable regions where the material near the fastener is nearly unloaded, which must be compensated for by regions where high stresses occur to achieve a particular

Mechanically fastened joints 611 required average load capacity. As mentioned above, certain types of adhesive joint, namely scarf joints between components of similar stiffness, can achieve a nearly uniform stress state throughout the region of the joint. In many cases, however, mechanically fastened joints cannot be avoided because of requirements for disassembly of the joint, for replacement of damaged structure, or to achieve access to underlying structure. In addition, adhesive joints tend to lack structural redundancy, and are highly sensitive to manufacturing deficiencies, including poor bonding technique, poor fit of mating parts and sensitivity of the adhesive to temperature and environmental effects such as moisture. Assurance of bond quality has been a continuing problem in adhesive joints, primarily because, while ultrasonic and X-ray inspection may reveal gaps in the bond, there is no present technique which can guarantee that a bond which appears to be intact by ultrasonic or Xray inspection does not lack load transfer capability, because of such factors as poor surface preparation. Surface preparation and bonding techniques have been well developed, but the possibility that lack of attention to detail in the bonding operation may lead to such deficiencies needs constant alertness on the part of those responsible for the bonding. Thus mechanical fastening tends to be preferred over bonded construction in highly critical and safety rated applications such as primary aircraft structural components, especially in large commercial transports, since assurance of the required level of structural integrity is easier to guarantee in mechanically fastened assemblies. Bonded construction tends to be more prevalent in smaller aircraft however and for non-aircraft applications as well as in non-flight critical aircraft components.

composites first came into use. It was found early in this period that the behavior of composites in bolted joints differs considerably from that which occurs with metals, primarily because stress concentrations are much more of a factor in joint behavior of composite structures, and stress analysis to quantify the level of various stress peaks is more important. It was fortunate that significant computing power became available in this period to keep up with the need for the intimate details of stress conditions around mechanical fasteners. The current approach to the design of mechanically fastened joints in composite structures evolved mainly out of a number of DoD, NASA and associated university programs aimed at providing a methodology which could be applied routinely to aircraft and other applications. Numerous stress analysis approaches to the mechanics of fastened joints have been conducted over the years since the introduction of 'advanced' composites in the mid-1960s. These have included: the work of Waszczak and Cruse' based on the boundary integral method; the use of two-dimensional complex variable elasticity solutions which treated the problem of variable contact around the fa~tenel-2-~; as well as recent work reported by Madenci and Illeri7; and a number of finite element approaches, especially the work of Crews and Naik8 which featured an inverse method for dealing with the contact problem. Hart-Smith9 developed an analytic approach based on the use of available solutions for isotropic plates with bolt-loaded holes as well as unloaded holes in plates under tension or compression which came out of classical efforts such as those reported in Petersonlo.The latter provided simple functional descriptions of the effect of joint geometry on peak stresses which, with various empirically derived correction factors introduced by Hart-Smith9, provided valuable insight into a number of trends in joint 28.3 MECHANICALLY FASTENED JOINTS behavior. In addition to the analytic efforts, several Mechanically fastened joints for composite structures have been under study since the fairly extensive programs aimed at the develmid 1960s when high modulus, high strength opment of design approaches for structural

612 Mechanical fastening and adhesive bonding systems have been supported by DoD and NASA4r6.Numerous papers have been presented Over the years in a series Of DoD/NASA/FAA 'Onferences On Fibrous Composites in Structural De~ign~,'~-'~. Many of the design principles which have been devel'ped to take into account the characteristics of bolted composite structures have been described4,6, 13. It is not possible within the scope Of this discussion to describe all the details and Processes that are necessary for achieving the design of specific joints. The objective is rather to give the reader Some insight into the factors that control the behavior of mechanifastened joints in structures' The behavior of mechanically fastened types Of information: joints is governed by (l)the features that the behavior Of the joint around individual fasteners and; (2) the behavior of multiple arrays of fasteners. The behavior of individual fasteners can be considered in terms of a generic rectangular element surrounding each fastener (Fig. 28.1), whose length and width are represented as ratios wi& respect to the fastener diameter. The effects of the geometry of this element together with effects of the reinforcement arrangement used in the laminate for the element determine the structural conditions under which the element will fail. Once the characteristics of the rectangular element are selected, its deformation characteristics can be combined with those of other rectangular elements making up the joint to obtain the performance of the joint as a whole. This discussion is organized in terms of these two aspects of joint design. In addition to a preliminary discussion of general features of mechanically fastened joints, the discussion which follows also considers: (1) single fastener joints, including effects of joint geometry together with those of composite material behavior; (2) multi-fastener joints; (3) fastener effects, and (4) a discussion of test methods which provide empirically-based data needed for completing the joint design.

28.3.1 KEY FEATURES OF MECHANICALLY FASTENED JOINTS

~~

Figure 28.1 represents a generic single fastener joint, while Fig. 28.2 depicts a multiple fastener configuration. Many of the most important features of joints are illustrated in the single fastener case s h o r n in Fig. 28.1. Key dimensions are D, the fastener diameter, t, the thickness of the joint structural elements, e, the edge distance (distance from the fastener center to the edge of the upper plate) and W, the width of the part of the upper plate to the left of the fastener. Similar featups apply to the lower plate element. Note that the Same load, p, is passed successively through various sections of the joint, including the bearing section in front of the fastener of area Dt which is in compression, the b o shearout sections of total area 2et which are loaded in primarily in shear, the net section, - D)t which is in tension, the gross section wt which is also in tension. The average stresses associated with these sections are:

(w

ab= P / D t ; average bearing stress, average shearout stress, a,, = P/2et; average net section stress, aN= P/(W - D)t; average gross section stress, aG= P / Wt. (28.1) LOAD PATH

P t

I

'

goss section

\

bearid

sectiqn

net section (W-O)t

wt

Dt

I I I

I I

,/-7-. c

I

I

!

:

I

Fig. 28.1 Single fastener joint.

I

J

section

Mechanically fastened joints 613 Axit!

tit h

w

Lateral Pitch

eff

*

Lateral Pitch

I

Fig. 28.2 Multi-fastener joint.

From the standpoint of the designer, the gross section strength is of primary interest since the objective of good design is to stress the gross section to its highest level. Structural performance of the joint can be rated in terms of joint efficiency, which refers to the ratio of average gross section stress at failure of the joint to the strength of the laminate in the gross section, essentially the strength achieved in coupon tests for tensile or compressive loading of unnotched specimens. For organic matrix composites, single fastener joints achieve joint efficiencies of less than 50%, while for multifastener joints the maximum achievable efficiency is of the order of 60%. In contrast, metallic joints can reach efficiencies of close to 80% because of the opportunity for taking advantage of local yielding around points of high stress, though even in metallic joints, design for avoidance of crack initiation in cases where long life under cyclic loading is required may force the joint efficiency to be lower than for single cycle loading.

-8-1

G

Fig. 28.3 Peak stresses around fastener.

gross section:

KGt = O,&JO~

(28.2)

or

qt= q J m x / o G

where , ) a is the maximum axial stress on the net section. Predictions of the peak stresses in the joint can be made using continuum elasticity analysis1", 7, 12, photoelastic measurements'O or finite element methodsE. 28.3.2 EFFECTS OF JOINT GEOMETRY The behavior of K$ as a function of D / W The peak value of axial stress on the net sec- for isotropic joint elements is given in Fig. 28.4 tion on,),,, (Fig. 28.3) is one of the key which was obtained from photoelastic meacontrolling parameters on joint performance. surements of KNnt against D / W (note that Kgt = It is convenient to express this as a stress con- KNnt/(l - D / W ) ) given by Petersenlo. Similar centration factor, i.e. a ratio with respect to curves for orthotropic plates were obtained by either the mean stress in the net section or the elasticity analysis2,3, 12.

614 Mechanical fastening and adhesive bonding There is obviously a competition for space between the bearing section and the net section. Increasing the fastener diameter, D, to lower abhas to decrease (W - D)t, the net section, thereby increasing uN, and vice versa. Furthermore, if D / W is much less than 1, the case is similar to that of a fastener in an infinitely wide plate; in such a case all the peak stresses become a constant multiplier of the bearing stress ab, which becomes large for D/W>E, or tL>>t,, the resulting P,, will be small compared with P,, so that the second fastener will be nearly unloaded. In such a case With the average gross section stresses and the joint will be equivalent to a single fastener strains for the two plate elements in Fig. 28.20 joint containing an unloaded hole at the locagive by: tion of the second fastener. On the other hand, OGU= ( P - P, )/tuW; the last equation showsthat if the stiffnesses of the upper and lower plates are equal, the two fasteners will be loaded equally. In general, fastener loads will be highly variable in a way that depends on the relative thicknesses and

Fig. 28.18 'OSC' failures for combined bearing

bypass loading17.

Mechanically fastened joints 623 moduli of adjacent plate elements in the regions around each fastener; fastener deflections due to beam-bending of the fasteners as well as clearance effects will add other complications to the situation. In Fig. 28.19, configuration A illustrates the behavior of joints with both elements tapered (i.e. scarf joints); due to load transfer between the elements by fastener load, the net section load P, decreases from the loaded (thick) end of each plate to its unloaded (thin)end, while the corresponding thicknesses decrease keeps the gross section stresses and strains and therefore the stretching deflections uniform along each element, providing for nearly equal load transfer at each fastener. For configuration B of Fig. 28.19, the case of uniform element thickness, the interior fastener loads are smaller than those at the joint ends, which is typical of this situation. Since the configurations shown at the right in Fig. 28.19 are arranged in order of increasing load capability (see the strains

Fig. 28.20 Two fastener joint.

listed under each configuration), it is noted that the scarfed configuration gives the lowest strength of the four, and is about 9% weaker than the uniform thickness configuration, B. It is usually expected that scarfing will lead to a way of keeping the bearing load minimized at the joint ends where the highest P,s are encountered, but other factors having to do with the balance between the effects of local

BOLT LOAD DISTRIBUTIONS 4-ROW BOLTED NINT

-

CONFIGURATION D

E,

O11(#lNIIN.

CONFIGURATIONC

-

= 0A)OMINIIN.

E-

CONFIQURATION A CONFIQURATION B 0- 4 CONFIGURATIONC 1n CONFIGURATION D O----Q

-

CONFIGURATION B 0#)46INIIN. Ob0IN.

1

2 3 BOLT NUMBER

1

1 A 112. 112. 112

3 4 0251N.

!

i I

-

.'I

-

CONFIGURATION A E , O m 1 INAN.

Fig. 28.19 Effectof joint configuration on fastener load di~tribution'~.

-

2

4

I'

624 Mechanical fastening and adhesive bonding element thickness and local bearing load predominates here. Note that configuration D, the strongest, uses a combination of variable fastener diameter and element tapering to achieve a maximum thickness of the outer elements which is greater than that in configuration C, to obtain maximum joint strength. Thus, judicious use of element tapering and variation of fastener diameters as well as other joint parameters can improve joint performance. With untapered joints, the maximum benefit of additional rows of fasteners is not much more than 20% greater than that for single row joints. Joint tapering will provide some improvement over that figure, although the benefit is limited; it should be kept in mind that the net section load at the loaded end of a given element will be the same for single fastener and multi-fastener joints, so that the benefit of adding more than a few rows may not be great. The interaction of the effects encountered in multi-fastener joints is fairly complicated and requires the use of analyses which can take into account the stresses and strains in single fastener configurations with bypass loading present (Fig. 28.13), representing 'unit cells' of the joint configuration, together with finite element calculations which evaluate the interactions among the various unit cells to provide the overall fastener load distribution. Computer codes have been developed under DoD and NASA sponsorship to provide for this type of integrated joint design. For example, Nelson, Bunin and Hart-Smith13 discuss the application of the well-known 'A4EJ' codeI9in conjunction with the code 'BJSFM' (Bolted Joint Stress Field Model15)which were developed by McDonnell Douglas under NASA and Air Force sponsorship for this type of joint design analysis. Northrop similarly developed codes 'SASCJ' (Stress Analysis of Single Fastener Composite Joints) and 'SAMCJ' (Stress Analysis of Multifastener Composite Joints) under Air Force Contract6,20. For information on the theory and application of these codes, the reader is directed to the references.

28.3.5 FASTENER EFFECTS

Joint strengths for local areas (Fig. 28.13) around individual fasteners are affected by a number of parameters associated with the fasteners. Some of these include: whether the joint is in single or double shear (Fig. 28.21); the use of countersunk (flush head) compared with protruding head fasteners (Fig. 28.22); effects of fastener diameter; effects of fastener clamping; fastener clearance effects, and effect of fastener deflections on fastener load distribution. Bearing strengths are significantly affected by the use of single compared with

p7=g=7p2p

P

(A)

Double Shear Configuration

~

P I'2

(B) Single Shear configuration Fig. 28.21 Single shear and double shear configurations.

+fP

3 (A)

Protruding Head Fastener

P P

(B) Countersunk Fastener Fig. 28.22 Countersunk and protruding head fasteners.

Mechanicallyfastened joints 625 double shear configurations,bearing strengths for single shear joints tending to run considerably below those for double shear because of greater through-the-thickness variation of fastener-plate contact pressure. Bearing strength tests referenced in Section 28.3.6 include separate test configurations for the two situations. In addition, as indicated in Fig. 28.22(a),bending moments tend to occur in single shear joints which are not present with double shear arrangements. Fastener head pull through (Fig. 28.10) can be a problem in the presence of such bending effects, and special test methods for fastener pull-through strength are described in Section 28.3.6. Countersunk, or flush head, fasteners (Fig. 28.22(b)) are frequently encountered in exterior surfaces of aircraft components where avoidance of air flow disturbance is required. Countersunk fasteners for composites include (Fig. 28.23) 'tension head' fasteners having the larger head depths and therefore wider heads, and 'shear head' fasteners having smaller head depths, with head angles ranging from 100" to 130". Countersunk fasteners tend to bear against the surrounding element more unevenly through the thickness than protruding head fasteners do. Tension head fasteners are generally preferred over shear head fasteners because of greater strength against head pull-through; however, if the joint element is so thin that the countersunk depth is greater than 70% of the element thickness, the tendency toward uneven bearing pressure in tension head fasteners is too great and shear head fasteners are recommended in this case.

The fastener diameter should be on the order of the thickness of the thicker of the plate elements making up the joint, or greater, ( D / t 2 1) to avoid excessive fastener bending. As in Fig. 28.10 (lower right-hand sketch) excessive bending can lead to failure of the fasteners, which is intolerable. In addition, fastener bending causes uneven distribution of bearing pressure through the plate element thicknesses, so that the full bearing strength is not available in such cases. The effect of large fastener deflections on the clamping pressure provided by the fastener is another adverse effect of fastener bending deflections. Figure 28.2413 illustrates the fact that bending deformations reduce the clamping pressure provided by fastener head, causing a reduction of bearing strength which is in addition to that caused by uneven bearing pressure through the thickness. The beneficial effect of clamping pressure on bearing strength, discussed earlier, has been well established. Required clamping levels are usually described in terms of bolt torques, 'finger tightened' being the lightest level, and installation requirements specify torque levels which supposedly represent particular bolt tensions (and therefore clamping

TENSION JOINT HlQH BEARING LOAD SIDE

LOELAMINATIONS DUE TO BEARING LOAD AND REDUCED HlQH BEARINQ LOAD SIDE

COMPRESSION JOINT

w-

Fig. 28.24 Effects of fastener bending on joint fail-

Fig. 28.23 Tension head and shear head fasteners.

ureI3.

626 Mechanical fastening and adhesive bonding pressures) that can be calculated in terms of the pitch of the fastener threads from machine design formulas. Bolt tensions for a given torque level are notoriously variable because of friction effects in the bolt threads, but specified torque levels which have been determined empirically probably represent minimum clamping levels necessary to insure maximum bearing strengths that can be achieved when variations in service conditions are taken into account. Steps should be taken to avoid loss of clamping pressure due to through-the-thickness viscoelastic deformation of the laminatez1”at elevated temperature and humidity. Fastener clearances are typically on the order of 0.075 mm (0.003 in) or less for typical 0.635 mm (0.25 in) aircraft fasteners. Analytical studies have shown that bearing stresses increase significantly for such relatively small clearances since the angle of contract decreases rapidly as the clearance increases. Clearances also have a significant effect on the fastener load distribution since the last in a series of fasteners cannot take up load until all the clearances have been taken up. In addition to the effect of clearances on fastener load distribution, effects of fastener bending deflections must be taken into account in load distribution calculations such as those provided by the A4EJ, SASCJ and SAMCJ codes described above. In the case of the two-fastener joint shown in Fig. 28.20, bending and rotational deflections of the fasteners will modify the load distribution described in the discussion of that figure for in Fig. zero fastener deflection. For E,t,>>E,t,, 28.20 for example, fastener deflections will allow some load to be transferred to the second fastener, as opposed to the case of no fastener deflections discussed earlier which led to Pf2= 0. Fastener deflection effects can be inferred from bolt bearing tests which provide for deflection measurements. Alternatively, analytical approaches based on beam models for the fastener which include both bending and shear deformations have been used13.

Complications, such as the way in which the fastener head and nut/washer combination bears on the surfaces of the plate element will influence the outcome of such calculations and must be taken into account. In addition, the through-thickness distribution of bearing pressure between the fastener and the surrounding the plate element should be included in the calculations. The method of Harris, qalvo and Hoosonz3which treats the bore of the fastener hole as an elastic foundation for the beam used to model the fastener has been applied6for such calculations. 28.3.6 TEST METHODS

Joint strength tests are needed to establish certain key parameters of the joint as inputs to design analyses. Such data as failure stresses for pure bearing load as obtained from single fastener coupons, and open hole coupon strengths for both tension and compression loading, are needed to establish joint performance for pure bearing and bypass loads. Intermediate combinations of bearing and bypass load must also be considered to provide empirical curves for dealing with the general situation. Because of differences in the way the fastener contacts the surrounding plate materials, bearing tests have to be conducted to treat both single and double lap (single and double shear) configurations. In single lap joints in particular, tests are needed to establish the effect of fastener rotation and bending deflection. Fastener deflections must be determined in tests of the type just described for providing fastener response data in connection with predictions of load distribution in multifastener joints. In addition, fastener head pull-through strength tests have to be performed to allow for joint configurations in which overall bending takes place, in which case out-of-plane forces between the fastener and joint plates tend to be sigruficant. The details of test methods for mechanically fastened joints are described by Shyprykevichz4 and in Mn-HDBK-1725.

Adhesive joints 627 28.4 ADHESIVE JOINTS

28.4.1 INTRODUCTION

Adhesive joints are capable of high structural efficiency and constitute a resource for structural weight saving because of the potential for elimination of stress concentrations which cannot be achieved with mechanically fastened joints. Unfortunately, because of a lack of reliable inspection methods and a requirement for close dimensional tolerances in fabrication, aircraft designers have generally avoided bonded construction in primary structure. Some notable exceptions include: bonded step lap joints used in attachments for the F-14 and F-15 horizontal stabilizers as well as the F-18 wing root fitting, and a majority of the airframe components of the Lear Fan and the Beech Starship. While a number of issues related to adhesive joint design were considered in the earlier literaturezG3, much of the methodology currently used in the design and analysis of adhesive joints in composite structures is based on the approaches evolved by L.J. HartSmith in a series of NASA/Langley-sponsored contracts of the early 1 9 7 0 as~ well ~ ~as~from the Air Force’s Primary Adhesively Bonded Structures Technology (PABST) programw3 of the mid-1970s. The most recent such work developed three computer codes for bonded and bolted joints, designated ‘A4EG’, ’A4EI’ and ’A4EKW under Air Force Contract . The results of these efforts have also appeared in a number of open literature publi~ations~’-~. In addition, such approaches found application in some of the efforts taking place under the NASA Advanced Composite Energy Efficient Aircraft (ACEE) program of the early to mid198Os5O,51. Some of the key principles on which these efforts were based include: (1)the use of simple one-dimensionalstress analyses of generic composite joints wherever possible; ( 2 ) the need to select the joint design so as to ensure failure in the adherend rather than the adhesive, so that

the adhesive is never the weak link;(3) recognition that the ductility of aerospace adhesives is beneficial in reducing stress peaks in the adhesive; (4) careful use of such factors as adherend tapering to reduce or eliminate peel stresses from the joint; (5) recognition of slow cyclic loading, corresponding to such phenomena as cabin pressurization in aircraft, as a major factor controlling durability of adhesive joints, and the need to avoid the worst effects of this type of loading by providing sufficient overlap length to ensure that some of the adhesive is so lightly loaded that creep cannot occur there, under the most severe extremes of humidity and temperature for which the component is to be used. Much of the discussion to follow will retain the analysis philosophy of Hart-Smith, since it is considered to represent a major contribution to practical bonded joint design in both composite and metallic structures. On the other hand, some modifications are introduced here. For example, the revisions of the Goland-Reissner single lap joint analysis36 have been re-revised according to the approach presented in Refs. 53,54. Certain issues which are specific to composite adherends but were not dealt with in the Hart-Smith efforts will be addressed. The most important of these is the effect of transverse shear deformations in organic composite adherends. 28.4.2 SUMMARY OF JOINT DESIGN CONSIDERATIONS

28.4.2.1 Effects of adherend thickness: adherend failures versus bond failures Figure 28.25 shows a series of typical bonded joint configurations. Adhesive joints in general are characterized by high stress concentrations in the adhesive layer. These originate, in the case of shear stresses, because of unequal axial straining of the adherends, and in the case of peel stresses, because of eccentricity in the load path. Considerable ductility is associated

628 Mechanical fastening and adhesive bonding

-

1

- (0) - 4

TAPERED SINGLE-UP JOINT DOUBLE-UPJOINT (F)

1 DUJBLE-STRAP JOINT

(0)

illustrate this point, shows a progression of joint types which represent increasing strength capability from the lowest to the highest in the figure. In each type of joint, the adherend thickness may be increased as an approach to achieving higher load capacity. When the adherends are relatively thin, results of stress analyses show that for all of the joint types in Fig. 28.26, the stresses in the bond will be small enough to guarantee that the adherends will reach their load capacity before failure can occur in the bond. As the adherend thicknesses increase, the bond stresses become relatively larger until a point is reached at which bond failure occurs at a lower load than that for which the adherends fail. This leads to the general principle that for a given joint type, the adherend thicknesses should be restricted to an appropriate range relative to the bond layer thickness. Because of processing considerations and defect sensitivity of the bond material, bond layer thicknesses are generally limited to a range of 0.125-0.39 mm (0.005-0.015 in). As a result, each of the joint

n

4

f

&

TAPERED STRAP JOINT

Fig. 28.25 Adhesive joint types”,

55.

with shear response of typical adhesives, which is beneficial in minimizing the effect of shear stress joint strength. The response of typical adhesives to peel stresses tends to be much more brittle than that to shear stresses, and reduction of peel stresses is desirable for achieving good joint performance. From the standpoint of joint reliability, it is vital to avoid the condition where the adhesive layer is the weak link in the joint, i.e. that the joint be designed to ensure that the adherends fail before the bond layer whenever possible. T ~ isE because failure in the adherends may be controlled, while failure in the adhesive is resin dominated, and thus subject to effects of voids and other defects, thickness variations, environmental effects, processing variations, deficiencies in surface preparation and other factors that are not always adequately controlled. This is a significant challenge, since adhesives are inherently much weaker than the composite or metallic elements being joined. However, the objective can be accomplished by recognizing the limitations of the joint geometry being considered and placing appropriate restrictions on the thicknesses the adherends for any given geometry. Figure 28.26, which 55 to has frequently been used by Hart-Smith39,

ADHEREND THICKNESS

Fig. 28.26 Joint geometry effects39.

Adhesive joints 629 types in Figs. 28.25 and 28.26 corresponds to a specific range of adherend thicknesses and therefore of load capacity, and as the need for greater load capacity arises, it is preferable to change the joint configuration to one of higher efficiency rather than to increasing the adherend thickness indefinitely. 28.4.2.2 Joint geometry effects Single and double lap joints with uniformly thick adherends (Fig. 28.25(b), (e) and ( f ) ) are the least efficient joint type and are suitable primarily for thin structures with low running loads (load per unit width, i.e. stress times element thickness). Of these, single lap joints are the least capable because the eccentricity of this type of geometry generates significant bending of the adherends that magnifies the peel stresses. Peel stresses are also present in the case of symmetric double lap and double strap joints, and become a limiting factor on joint performance when the adherends are relatively thick. Tapering of the adherends (Figs. 28.25(d) and (g)) can be used to eliminate peel stresses in areas of the joint where the peel stresses are tensile, which is the case of primary concern. For joints between adherends of identical stiffness, scarf joints (Fig. 28.25(i))are theoretically the most efficient, having the potential for complete elimination of stress concentrations. (In practice, some minimum thickness corresponding to one or two ply thicknesses must be incorporated at the thin end of the scarfed adherend leading to the occurrence of stress concentrations in these areas.) In theory, any desirable load capability can be achieved in the scarf joint by making the joint long enough and thick enough. However, practical scarf joints may be less durable because of a tendency toward creep failure associated with a uniform distribution of shear stress along the length of the joint unless care is taken to avoid letting the adhesive be stressed into the nonlinear range. As a result, scarf joints tend to be used only for repairs of very thin structures.

Scarfjoints with unbalanced stiffnessesbetween the adherends do not achieve the uniform shear stress condition of those with balanced adherends, and are somewhat less structurally efficient because of rapid buildup of load near the thin end of the thicker adherend. Step lap joints (Fig. 28.25(h)) represent a practical solution to the challenge of bonding thick members. This type of joint provides manufacturing convenience by accommodating the layered structure of composite laminates. In addition, high loads can be transferred if sufficiently many short steps of sufficiently small ’rise’ (i.e. thickness increment) in each step are used, while maintaining sufficient overall length of the joint. 28.4.2.3 Effects of adherend stiffness unbalance All types of joint geometry are adversely affected by unequal adherend stiffnesses, where stiffness is defined as axial or in-plane shear modulus times adherend thickness. Where possible, the stiffnesses should be kept approximately equal. For example, for step lap and scarf joints between quasi-isotropic carbon epoxy (Young’s modulus = 55 GPa = 8 x lo6 lb/in2) and titanium (Young’s modulus = 110 GPa = 16 x lo6 lb/in2) ideally, the ratio of the maximum thickness (the thickness just beyond the end of the joint) of the composite adherend to that of the titanium should be 110/55 = 2.0. 28.4.2.4 Effects of ductile adhesive response Adhesive ductility is an important factor in minimizing the adverse effects of shear and peel stress peaks in the bond layer. If peel stresses can be eliminated from consideration by such approaches as adherend tapering, strain energy to failure of the adhesive in shear has been shown by Ha~?-Smith~~ to be the key parameter controlling joint strength; thus the square root of the adhesive strain energy

630 Mechanical fastening and adhesive bonding density to failure determines the maximum sta- transverse tension, as a result of which the tic load that can be applied to the joint. The limiting element in the joint may be the interwork of Hart-Smith has also shown that for pre- laminar shear and transverse tensile strengths dicting mechanical response of the joint, the of the adherend rather than the bond strength. detailed stress-strain curve of the adhesive can Ductile behavior of the adherend matrix can be replaced by an equivalent curve consisting be expected to have an effect similar to that of of a linear rise followed by a constant stress ductility in the adhesive in terms of response plateau (i.e. elastic-perfectly plastic response) if of the adherends to transverse shear stresses, the latter is adjusted to provide the same strain although the presence of the fibers probably energy density to failure as the actual limits this effect to some extent, particularly in stress-strain curve gives. Test methods for regard to peel stresses. The effect of the stacking sequence of the adhesives should be aimed at providing data on this parameter. Once the equivalent elastic- laminates making up the adherends in composperfectly plastic stress-strain curve has been ite joints is sigruficant. For example, 90” layers identified for the selected adhesive for the most placed adjacent to the bond layer theoretically severe environmental conditions (temperature act largely as additional thicknesses of bond and humidity) of interest, the joint design can material, leading to lower peak stresses, while proceed through the use of relatively simple 0” layers next to the bond layer give stiffer one-dimensional stress analysis, thus avoiding adherend response with higher stress peaks. In the need for elaborate finite element calcula- practice it has been observed that 90” layers tions. Even the most complicated of joints, the next to the bond layer tend to seriously weaken step lap joints designed for root-end wing and the joint because of transverse cracking which tail connections for the F-18 and other aircraft, develops in those layers, and advantage cannot have been successfully d e ~ i g n e d ~ and ~ , ~ ”be ~ taken of the reduced stresses. Large disparity of thermal expansion charexperimentally demonstrated using such approaches. Design procedures for such analy- acteristics between metal and composite ses which were developed on Government adherends can pose severe problems. contract have been incorporated into public Adhesives with high curing temperatures may domain in the form of the ’A4EG’, ‘A4EI’ and be unsuitable for some uses below room tem‘A4EK computer codesmentioned previ- perature because of large thermal stresses ously in Section 28.4.1.Note that the A4EK code which develop as the joint cools below the fabpermits analysis of bonded joints in which local rication temperature. Composite adherends are relatively pervidisbonds are repaired by mechanical fasteners. ous to moisture, which is not true of metal adherends. As a result, moisture is more likely 28.4.2.5 Behavior of composite adherends to be found over wide regions of the adhesive Organic matrix composite adherends are con- layer, as opposed to confinement near the siderably more affected by interlaminar shear exposed edges of the joint in the case of metal and tensile stresses than metals, so that there is adherends, and response of the adhesive to a significant need to account for such effects in moisture may be an even more significant stress analyses of joints. Transverse shear and issue for composite joints than for joints thickness-normal deformations of the between metallic adherends. adherends have an effect analogous to thickening of the bond layer, corresponding to a 28.4.2.6 Effects of bond defects lowering of both shear and peel stress peaks. On the other hand, the adherend matrix is Defects in adhesive joints which are of concern often weaker than the adhesive in shear and include surface preparation deficiencies, voids

Adhesive joints 631 and porosity, and thickness variations in the adherends, porosity may grow catastrophibond layer. cally and lead to non-damage tolerant joint Of the various defects which are of interest, performance. surface preparation deficiencies are probably Bond thickness variations'jl usually take the the greatest concern. These are particularly form of thinning due to excess resin bleed at troublesome because there are no current non- the joint edges, leading to overstressing of the destructive evaluation techniques which can adhesive in the vicinity of the edges. Inside detect low interfacial strength between the tapering of the adherends at the joint edges bond and the adherends. Most joint design will compensate for this condition; other comprinciples are academic if the adhesion pensating techniques are also discussed'jl. between the adherends and bond layer is poor. Bond thicknesses, per se, should be limited to The principles for achieving good adhesion of ranges of 0.12-0.24 mm (0.005-0.01 in) to prethe bond to the adherends (see Chapter 29) are vent significant porosity from developing well established for adherend and adhesive although greater thicknesses may be acceptcombinations of interest. Hart-Smith, Brown able if full periphery damming or high and Won$ give an account of the most crucial minimum viscosity paste adhesives are used. features of the surface preparation process. Common practice involves the use of film Results shown in that reference suggest that adhesives containing scrim cloth, some forms surface preparation which is limited to of which help to maintain bond thicknesses. It removal of the peel ply from the adherends is also common practice to use mat carriers of may be suspect, since some peel plies leave a chopped fibers to prevent a direct path for residue on the bonding surfaces that makes access by moisture to the interior of the bond. adhesion poor. (However, some manufacturers have reported satisfactory results from 28.4.2.7 Durability of adhesive joints surface preparation consisting only of peel ply removal.) Low pressure grit b l a ~ t i n g ~ is ~ ,Hart-Smith45 ~~ discusses differences in durabilpreferable over hand sanding as a means of ity assessment of adhesive joints between eliminating such residues and mechanically concepts related to creep failure under cyclic conditioning the bonding surfaces. loading and those related to crack initiation For joints which are designed to ensure that and propagation which require fracture the adherends rather than the bond layer are mechanics approaches for their interpretation. the critical elements, tolerance to the presence In summary, Hart-Smith suggests that if peel of porosity and other types of defect is consid- stresses are eliminated by adherend tapering erable45.Porosity'jOis usually associated with or other means, and if the principle discussed overthickened areas of the bond, which tend in Section 28.4.2.1 of limiting the adherend to occur away from the edges of the joint thickness to ensure failure of the adherends where most of the load transfer takes place, rather than the adhesive is followed, crackand thus is a relatively benign effect, espe- type failures will not be observed under cially if peel stresses are minimized by time-varying loading, failures being related adherend tapering. In such cases6", porosity primarily to creep fatigue at hot wet condican be represented by a modification of the tions, in joints with short overlaps which are assumed stress-strain properties of the adhe- subject to relatively uniform distributions of sive as determined from thick-adherend tests, shear stress along the joint length. Additional allowing a straightforward analysis of the discussion of viscoelastic response of bonded effect of such porosity on joint strength, as in joints is There is an extensive body of literature6571 the A4EI computer code. If peel stresses are significant, as in the case of over-thick on fracture mechanics approaches to joint

632 Mechanical fastening and adhesive bonding durability, based on measurement of energy release rates for various adhesives together with analytical efforts aimed at applying them to joint configurations of interest. In particular, Johnson and Malln report fatigue crack initiation in bonded specimen configurations with adherend tapering aimed at reduction of peel stresses in varying degrees, in some cases practically eliminating them; data in Ref. 92 indicate that crack initiation will occur even with the adhesive in pure shear, for cycling to lo6cycles above loading levels which are probably considerably below static failure loads. The results given” suggest that for combinations of peel and shear stressing, total (mode 1 + mode 2) cyclic energy release rate can be used to determine whether or not cracking will occur. However, Hart-Smith reportedfi that in ’thick adherend‘ test specimens that provide a relatively uniform shear stress distribution in the adhesive (see MIL-HDBK-17, Vol. 1, Chapter 7, Section 7.3) which were subjected to fatigue tests in the PABST programM,cycling to more than lo7 cycles applied at high cycling rates (30 Hz) were achieved without failure of the adhesive, although in certain cases, namely those involving 6.27 mm (0.25 in) adherend thicknesses, fatigue failures of the metal adherends did result. More study is needed to resolve some of the apparently contradictory results which have come out of various studies. 28.4.3 STRESSES IN ADHESIVE JOINTS

28.4.3.1 General

Stress analyses of adhesive joints have ranged from very simplistic ‘P over A’ formulations in which only average shear stresses in the bond layer are considered, to extremely elegant elasticity approaches that consider fine details, e.g. the calculation of stress singularities for application of fracture mechanics concepts. A compromise between these two extremes is desirable, since the design of structural joints does not usually depend on the fine details of the stress distributions. Since practical consid-

erations force bonded joints to incorporate adherends which are thin relative to their dimensions in the load direction, stress variations through the thickness of the adherend and the adhesive layer tend to be moderate. Such variations do tend to be more sigruficant for organic composite adherends because of their relative softness with respect to transverse shear and thickness normal stresses. However, a considerable body of design procedure has been developed based on ignoring thicknesswise adherend stress variations. Such approaches involve using one-dimensional models in which only variations in the axial direction are accounted for. Accordingly, the bulk of the material to be covered here is based on simplified one-dimensional approaches characterized by the work of Hart-Smith. The Hart-Smith approach makes extensive use of closed form and classical series solutions since these are ideally suited for making parametric studies of joint designs. The most prominent of these have involved modification of Volkersen26 and Goland-ReissnerZ7solutions to deal with ductile response of adhesives in joints with uniform adherend thicknesses along their lengths, together with classical series expressions to deal with variable adherend thicknesses encountered with tapered adherends, and scarf joints. Simple lap joint solutions described below calculate shear stresses in the adhesive for various stiffnesses and applied loadings. For the more practical step lap joints, the described expressions can be adapted to treat the joint as a series of separate joints, each having uniform adherend thickness. 28.4.3.2 Adhesive shear stresses

Figure 28.27 shows a joint with ideally rigid adherends in which neighboring points on the upper and lower adherends slide horizontally with respect to each other when the joint is loaded to cause a displacement difference 6 = uu - uLrelated to the bond layer shear strain by yb = 6 / f b .The corresponding shear stress, zb, is given by zb = Gbyb. The rigid adherend

Adhesive joints 633 one for which E,tL >> E&), stretching elongations in the upper adherend lead to a shear strain increase at the right end of the bond layer. The case in which both adherends are equally deformable, shown in Fig. 28.29(b), indicates a bond shear strain increase at both ends due to the increased axial strain in dTU/dx= zb (28.6) whichever adherend is stressed at the end leads to a linear distribution of Tu and TL under consideration. For both cases, the varia(upper and lower adherend resultants) as well tion of shear strain along the bond results in an as the adherend axial stresses uxuand ax,indi- accompanying increase in shear stress which, cated in Fig. 28.28. These distributions are when inserted into the equilibrium eqn (28.6) leads to a nonlinear variation of stresses. The described by the following expressions: Volkersen shear lag analysisz6provides the simplest calculation of adhesive shear stresses for the case of deformable adherends. This involves the solution of the following differential equation:

assumption implies that 6, y, and t, are uniform along the joint. Furthermore, the equilibrium relationship indicated in Fig. 28.27(c),which requires that the shear stress be related to the resultant distribution in the upper adherend by

where ax = T/t. In actual joints, adherend deformations will cause shear strain variations in the bond layer which are illustrated in Fig. 28.29. For the case of a deformable upper adherend in combination with a rigid lower adherend shown in Fig. 28.29(a) (in practice,

B,

E&;

B, = E,t,

(28.8)

which applies to the geometry of Fig. 28.30

I [a RIGID

=

KkU

f-

Fig. 28.27 Elementary joint analysis (rigid adherend model).

I

634 Mechanical fastening and adhesive bonding (A)

(a) AXIAL STRESS DISTRIBUTXOH

AXIAL RESULTANT DISTRIBUTION

Fig. 28.28 Axial stresses in joint with rigid adherends.

[A) RIGID UIIW

AaEml

--f

TFig. 28.29 Adherend deformations in idealized joints.

below. The solution for this equation which provides zero traction conditions at the left end of the upper adherend and the right end of the lower adherend, together with the applied load T at the loaded ends gives the resultants as:

-

TL = T - T U

where

-

t=-

+-

1 +PB SinhPZ/t

(28.9)

tu + tL ; PB = 2

BL/B"

Using eqn (28.6) to obtain an expression for the shear stress distribution leads to:

Adhesive joints 635 I 0I

X

+- PB

1 + pB tanhpZ/T

)

(28.11)

where Gx= Fig. 28.30 Geometry for Volkersen solution.

Also of interest in the discussion which follows is the minimum shear stress in the joint. This occurs approximately at x = 1/2, leading to:

B, 2 B,;

I

0

44

&4

a6

Q1

T/t

to be discussed subskquently. Figure 28.31

t

1.2

1.4

1s

1 1

2

IS

1

I

.-. .

4

U a.4

.

.

.

:

:

.

.

OS

a.0

1

IS!

1.4

IJ

1.8

#-

b

2

. . . . . . . . . . o)

a4 w os

1

IZ 11) i g t i

2

I---

I

636 Mechanicalfastening and adhesive bonding

Fig. 28.32 Comparison of average and maximum shear stress vs. l / t .

shows the distribution of axial adherend stresses and bond layer shear stress for two cases corresponding to E, = E , and E, = 10Eu with tu = t,, p = 0.387 and l / t = 20 for both cases (giving p l / t = 7.74) and a nominal adherend stress 0, = 10. As in the approximate analysis given earlier, the shear stresses given by eqn (28.10) are maximum at both ends for equally deformable adherends (B, = B,); for dissimilar adherends with the lower adherend more rigid (B, > E$,), the maximum shear stress obtained from eqn (28.10) occurs at the right end of the joint where x = I , again as it did for the approximate analysis. Figure 28.32 compares the behavior of the maximum shear stress with the average shear stress as a function of the dimensionless joint length, l / t , for equal adherend stiffnesses. The point illustrated here is the fact that although the average shear stress continuously decreases as the joint length increases, for the maximum shear stress which controls the load that can be applied without failure of the adhesive, there is a diminishing effect of increased joint length when q = p l / t is much greater than about 2.

An additional point of interest is a typical feature of bonded joints illustrated in Fig. 28.31(d) which gives the shear stress distribution for equal adherend stiffness, namely, the fact that high adhesive shear stresses are concentrated near the ends of the joint. Much of the joint length is subjected to relatively low levels of shear stress, which implies in a sense that that region of the joint is structurally inefficient since it does not provide much load transfer. However, the region of low stress helps to improve damage tolerance of the joint since defects such as voids and weak bond strength may be tolerated in regions where the shear stresses are low, and in joints with long overlaps this may include most of the joint. In addition, Hart-Smith has suggested51 that when ductility and creep are taken into account, it is a good idea to have a minimum shear stress level no more than 10% of the yield strength of the adhesive, which requires some minimum value of overlap length. Equation (28.12) can be used to satisfy this requirement for the case of equal stiffness adherends. The two special cases of interest again are for equal adherend stiffness and a

Adhesive joints 637 rigid lower adherend, since these bound the range of behavior of the shear stresses. As a practical consideration, we will be interested primarily in long joints for which pZ/t >> 1. For these cases eqn (28.11) reduces to:

relatively obvious due to the offset of the two adherends which leads to bending deflection as in Fig. 28.33@).In the case of double lap joints, as exemplified by the configuration shown in Fig. 28.34, the load path eccentricity is not as obvious, and there may be a tendency p1/t >> 1; to assume that peel stresses are not present for this type of joint because, as a result of the lateral symmetry, there is no overall bending deflection. However, a little reflection brings to mind the fact that while the load in the sym1 B, = B,; zJrnaX=-pax (28.13) metric lap joint flows axially through the 2 central adherend prior to reaching the overlap Thus, for long overlaps, the maximum shear region, there it splits in two directions, flowing stress for the rigid adherend case tends to be laterally through the action of bond shear twice as great as that for the case of equally stresses to the two outer adherends. Thus deformable adherends, again illustrating the eccentricity of the load path is also present in adverse effect of adherend unbalance on shear this type of joint. As seen in Fig. 28.34(c), the shear force, designated as F,, which represtress peaks. sents the accumulated effect of zb for one end of the joint, produces a component of the total 28.4.3.3 Peel stresses moment about the neutral axis of the upper Peel stresses, i.e. through-the-thickness exten- adherend equal to FsHz/2. (Note that F , is sional stresses in the bond, are present because equivalent to T/2, since the shear stresses react the load path in most adhesive joint geome- this amount of load at each end.) The peel tries is eccentric. It is useful to compare the stresses, which are equivalent to the forces in effect of peel stresses in single and double lap the restraining springs shown in Fig. 28.34(b) joints with uniform adherend thickness, since peel stresses are most severe for joints with uniform adherend thickness. The load path eccentricity in the single lap joint (Fig. 28.33) is

, \

i Fig. 28.33 Peel stress development in single lap ioints.

Fig. 28.34 Peel stress development in double lap ioints.

J

638 Mechanical fastening and adhesive bonding It is important to understand that peel and (c) have to be present to react the moment stresses are unavoidable in most bonded joint produced by the offset of FsHabout the neutral configurations However, they can often be axis of the outer adhered. Peel stresses are reduced to acceptable levels by selecting the highly objectionable. Later discussion will adherend geometry appropriately. indicate that effects of ductility significantly reduce the tendency for failure associated with shear stresses in the adhesive. On the other 28.4.3.4 Effects of joint geometry hand, the adherends tend to prevent lateral contraction in the in-plane direction when the In this section the behavior of joints is considbond is strained in the thickness direction, ered with linear response of the adhesive in which minimizes the availability of ductility shear assumed. Effects of ductility will be coneffects that could provide the same reduction sidered later. of adverse effects for the peel stresses. This is illustrated by the butt tensile test shown in Fig. 28.4.3.4.1 Single and double lap joints with 28.35 in which the two adherend surfaces adjauniform adherend thickness cent to the bond are pulled away from each other uniformly. Here the shear stresses asso- Double lap joints will be considered first since ciated with yielding are restricted to a small they are somewhat simpler to discuss than sinregion whose width is about equal to the gle lap joints because of deflection effects in the thickness of the bond layer, near the outer latter. Shear and peel stresses in double lap edges of the system; in most of the bond, rela- joints with uniform adherend thickness were tively little yielding can take place. For organic treated by Hart-Smith%.For the shear stresses, matrix composite adherends, the adherends the type of analysis discussed in Section 28.4.3.2 may fail at a lower peel stress level than that at can be applied with suitable changes in notawhich the bond fails, which makes the peel tion, i.e. the expressions for the shear stresses given in eqns (28.11)and (28.12)can be applied stresses even more undesirable. with subscripts 'i' and '0' ('inner' and 'outer') substituted here for 'L' and ' U ('lower' and 'upper') used in eqns (28.6-28.11); in addition, the outer adherend thickness in the earlier equations is now equivalent to half the thickness of inner adherend because of vertical symmetry of the double lap joint. However, we will also introduce the effects of thermal mismatch effects in the following expressions for later reference. The notation used here is: Bond

P t

P

Bo = toEo; Bi = tiEi; E@e Region (Distornal Strains)

pB = Bi/Bo;

B B. Tth = (ao- a,)AT; Bo + Bi L A

Fig. 28.35 Shear stresses near outer edges of butt tensile test.

ax = T / t ; &* =T,/t

-

(28.14)

Adhesive joints 639 where a,, a, are thermal expansion coefficients and AT is the temperature change. Note that is related to the resultants (axial adherend stress times thickness) at the ends of the joint as shown in Fig. 28.36. The shear stresses are then given by: Zb

[

=/35

coshp(x - I)/: +pB sinhBl/t 1

~

assuming that Bi 1 Bo, the maximum value of the shear stresses occurs at the right end of the joint as noted earlier (Fig. 28.31). With thermal effects present, the situation is complicated by the sign of &* which is positive if (a, - a,)and AT have the same sign and negative otherwise. The peel stresses in the double lap joint are described by a beam-on-elastic foundation type differential equation of the form:

**--

-

coshp(Z - x ) / t

]

(A) DOUBLE STRAP JOINT

2t

1*’ 1

d40 ?d b + 4 - 0 dP t4

=-f-

2

yd

=

O 114

(28.18b)

The solution to eqn (28.18) depends on whether a strap joint or a lap joint is considered. The exact form of the solution contains products of hyperbolic and trigonometric functions but for the practical situation of joints longer than one-or-two adherend thicknesses and B 1 failure has not occurred and R represents a factor of safety (e.g. if R = 2, then the applied stress can be safely doubled

[ F I , ~ l ~ ~+][F,u,] R‘ R -1 = 0

i, j

(30.29)

= x, y, s

The positive and negative roots of the quadratic equation can be found and represent failure of the laminate in tension and compression (where the absolute value of the negative root is used), respectively. Failure envelopes can be plotted to show laminate strength for any combination of loads. Instead of the stress space representation, however, the examination of failure envelopes in strain space is a useful alternative. The representation of failure envelopes in strain space is preferred because strain is usually specified in laminated plate theory. Strain, unlike stress, is at most a linear function of the thickness. Furthermore, failure envelopes are fixed in strain space, and are independent of other plies with different angles which may exist in a laminate. Thus, they can be regarded as material properties. Another additional advantage of strain space is that the axes are dimensionless. 30.4.2 STRENGTH OF LAMINATES

Traditional failure criteria based on strength of materials are limited to the prediction of the FPF, the point beyond which the continuous and homogeneous material assumptions are no longer valid. The use of a simple method for modeling of degraded plies is recommended, from which the FPF can be estimated. The load-carrying capability of a laminate

Laminate design 697 beyond the FPF can be formulated using a ply degradation model. Two possible methods are recommended: first, the simplified micromechanics model based on the modified rule-of-mixtures relations can be used. Plies with transverse cracks are replaced by plies with reduced matrix modulus, Em. Micromechanics translates the effect of the altered constituent material properties to the ply level, e.g. how a change in the matrix modulus affects the shear and transverse modulus of the unidirectional ply. Degraded plies are modeled by quasi-homogeneous plies so that laminated plate theory can be reapplied to determine the ply stresses and ply strains. Another approach for the prediction of post-FPF strength can be based on macromechanics, without resorting to micromechanics. The degradation factor (DF) is applied directly to the transverse and shear modulus, as well as the major Poisson's ratio. The exact value for the degradation factor must be determined empirically.A value between 0.1 and 0.3 is recommended. If the degradation factor is given a value close to zero, the quadratic failure criterion can be made to resemble the maximum strain criterion and results in a generally conservative estimation of laminate strength.

of the preselected orientations results in a quasi-isotropic laminate. This is the performance baseline, because load-carrying fiber is in effectively all directions. Laminate performance can only be improved beyond that of a quasi-isotropic laminate as fiber is biased into load directions, since, of course, fiber would never be put in unnecessary directions. Heretofore, quasi-isotropic laminates have been used because they give properties like those of metals, and predictable responses that are familiar, although they are not optimal in strength-to-weight or stiffness-to-weight ratios. Many laminates used today on aircraft structures tend to be of this type. In general, however, the more directional the loading, the bigger the payoff possible with anisotropic tailoring. To improve on the performance obtained with a quasi-isotropic laminate, the cost to design and analyze the anisotropic part (using the tools like those discussed in this chapter) is unfortunately often thought not to be worth the additional weight savings. This attitude is commonly rationalized by worry about holes, increase in work associated with more complicated fiber placement (preform assembly), etc. In practice, laminate designs, if not quasiisotropic, are certainly still symmetric about the midplane, balanced (equal quantity of -8 30.5 LAMINATE DESIGN and +8 plies), and orthotropic. Capitalizing on To simplify the analysis, it is commonly initially the benefits of anisotropy will probably occur specified that a laminate will be constructed of in other industries first before being adopted plies oriented with fibers in a few preselected by the more conservative aircraft industry. directions, where only the percentage distribuAn exception to traditional aircraft laminate tion in each orientation must then be design is the X-29 experimental aircraft, which determined. Laminates with plies distributed demonstrated a unique attribute of anisotropy every 45" are called n/4 laminates (plies can be (Fig. 30.8). The basis for this design lies in the in the 0, 45, 90 or 4 5 directions. Ply orienta- important assumption that the 1,2,6 axes are tions are usually specified as a value between usually the primary load directions for the -90 and 90". For example, instead of identifying laminate. With the coordinate system for loadthe orientation as 135, the laminate orientation ing changed to be 20" off a designated is more commonly called 45", although they laminate system, it can be shown that the lamare the same). Another class of laminates are inate behavior in flexure and torsion is called n/3, where plies are placed every 60" coupled. In fact, twisting will result with flex(plies can be in the 0, 60 or -60 directions). In ural loading, even though the material would both cases, an equal percentage of plies in each normally behave as most metals. This is the

698 Laminate design

principle used on the X-296. The normal ten- the laminate. Composite materials are not dency for forward swept wings to diverge at merely a light-weight substitute for heavyincreasing speeds was counteracted by this weight metals. Structural performances which laminate design: the increase in lift creates a are not possible with metals are easily achievdecrease in angle of attack, as the laminate able. Examples of such unique properties twists in the direction opposing the forces. include Poisson’s ratios greater than unity or It is conceivable that in the future the even negative, bending-twisting coupling, and graphite golf shafts currently gaining in popu- zero or negative coefficients of thermal expanlarity could be tailored to the individual golfer. sion (CTE). The problems and examples below The same coupling principle could be applied. illustrate the engineering constants of angleA golfer’s tendency to consistently slice the ply and related laminates. Examples of large ball might allow the designer to customize a and negative Poisson’s ratios and examples of golf shaft which not only bends, but also bend-twist coupling are also given. twists slightly under the bending load of the bad swing. 30.5.2 UNUSUAL POISSON’S RATIOS

30.5.1 UMQUE BEHAVIOR

The most unique features of composite materials are the highly direction-dependent properties. Highly coupled deformation and load-carrying capability can be designed into

Personal computer software based on a computer spreadsheet allows rapid sensitivity studies and parametric analysis of the behavior of laminates. Laminated plate theory with micromechanics is programmed into ’MicMac/In-Plane’2. A companion charting tool,

Fig. 30.8 Top view of the Grumman X-29 aircraft with wings that twist under flexure to counteract the detrimental aerodynamic effects.@ NASA)

Laminate design 699 'Chart-quick', can be used to plot variation of CTE as a function of independent variables (0, E,,, E,, vf, etc.). For the following problems and examples, the carbon fiber reinforced polymer material data used are shown in Table 30.1. Figure 30.9(a) shows the engineering constants for a unidirectional laminate as it is rotated from the on-axis. The Poisson's ratio, vx, of a 0" laminate is approximately 0.3. With increasing angle of the off-axis laminate, the Poisson's ratio decreases. The Poisson's ratio of a 90" laminate is effectively zero, because contraction in the transverse direction is constrained by the fibers. Figure 30.9(b) shows the engineering constants for an angle-ply laminate. It is interesting to observe the very large Poisson's ratio of 1.32 for a [ S O ] laminate. A value of greater than one implies that the transverse dimensional change is more than in the dimensional change in the longitudinal direction of loading. When the ply angle is either 0 or 90", the laminates (and consequently the values for the engineering constants) in Figs. 30.9(a) and 30.9(b) are the same. In both Figs. 30.9(a) and 30.9(b), the transverse modulus, E,, is a 'mirror

Table 30.1 Material property data for three different carbon fiber systems: IM6/Epoxy, T300/5208 and M40J/F584 lM6/ Epoxy

T300/

5208

M40J/ F854

Longitudinal tensile modulus, E x (Msi)

29.44

26.27

32.8

Transverse modulus, EY(Msi)

1.62

1.49

1.2

Poisson's ratio

0.32

0.28

0.26

Shear modulus, Es (Msi)

1.22

1.04

0.66

Longitudinal CTE, a1 Transverse CTE, a2

15

Volume fraction V,(%)

66

70

22.50

1.5 22.50

15.00

1 0.5

7.50

0.00

0

0.00

f

Poisson's Ratio 2 1.5

15.00

7.50

Ply Angle, 8 (degrees)

62

image' of the longitudinal modulus, Ex. Figure 30.10(a) shows the engineering constants for cross-ply laminates. For any given laminate, the longitudinal modulus, Ex, and the transverse modulus, E , are equal. The Poisson's ratio, vx, of a [d/90] laminate is approximately zero, because of the presence of fibers in the transverse direction. The largest

Modulus Poisson's (Msi) Ratio 30.00 2

Modulus (Msi) 30 00

(4

-0.14

-0.5

(b)

1

'

0.5 0 15

30

45

60

75


Fig. 30.9 Engineering constants of IM6/epoxy laminates as a function of 6 for (a) off-axis unidirectional and (b) mgle-ply [+el,.

[e],;

700 Laminate design

LO7 (6 + 90>1,,

LO,

f

6,Is

Modulus

Poisson's

(Msi) 30.00 22.50 15.00 7.50

0.00 15

30

45

60

15

75


(4

30

45

60

75


(b)

Fig. 30.10 Engineering constants of IMG/epoxy laminates as a function of 19for (a) cross-ply [I9,(0+ 90)],,;

and @) LO,,

* qs.

Poisson's ratio is 0.55 for a [*45] laminate. The shear modulus, E , is a maximum, of course, for the [*45] laminate. Figure 30.10(b) shows the engineering constants for laminates with 50% 0" plies and 50% angle-plies. With the exception of the transverse modulus, the results are similar to those for the angle-ply laminate shown in Fig. 30.9(b).When the ply angle is 90", the values for the engineering constants in Figs. 30.10(a) and 30.10(b)are the same. Figures 30.11(a) and 30.11(b)show the engineering constants for some unusual laminates. When the ply angle, 8, is 15", Fig. 30.11(a) shows an off-axisunidirectional laminate and Fig. 30.11(b)shows an angle-ply. For all other ply angles, the laminates are unbalanced. From Fig. 30.11(a),it can be observed that the [15/60Is laminate exhibits an extremely large negative Poisson's ratio of -0.32, meaning the laminate will expand in the transverse direction under longitudinal tension loading and compress in the transverse direction under longitudinal compressive loading. From Fig. 30.11(b), it can be observed that the [-15/30Is

laminate exhibits a very large Poisson's ratio of 1.32, when compared with that of an isotropic material (0.3). Besides the unique Poisson's ratio behavior, it is also important to examine the values of the other coupling coefficients.

EXAMPLE Table 30.2 considers the resulting deformations on coupon specimens under load, and Fig. 30.12 indicates the relative magnitude of deformation due to large and negative Poisson's ratios. 30.5.3 STIFFNESS AND COUPLING

It is useful to look at the A, B, D stiffness matrices of some simple laminates. For ease of comparison, the stiffness matrices can be normalized to have units of [force/length2]by defining

[A*]= [ A ] / h , [B*] = 2[B]/h2,

Laminate design 701 Poisson‘s Ratio T 2

,r”

30.00 T ’ 22.50

[-I 5/9 14s

Modulus (Msi) 30’00

Poisson‘s Ratio

T

22.50

1.5

15.00

1

1

15.00 7.50

0.5

7.50

0.5

0.00

0

0.00

0

1


1 -0.5

(4


-0.5

(b)

Fig. 30.11 Engineering constants of IM6/epoxy laminates as a function of 8 for (a) [15/8],s; and (b) [-w~I,.

Table 30.2 Strains, deformations and strength ratio (based on first-ply-failure) of 10 in x 1 in x 0.1 in specimens under 1000 lb longitudinal load, N, Longitudinal strain Material

Transverse strain E2

El

(1C3in/in)

in/in)

Longitudinal Transverse displacement displacement A1

A2

(1C3in)

(10-3 in)

3.4 3.4 6.5 9.5 10.2 10.9 24.0 32.0 62.5

-1.0 -1.1 -0.2 -12.9 -3.0 3.5 -17.7 -20.7 -1.1

Strength ratio R

.-

40ksi Steel IM6/Epoxy IM6/Ep IM6/Ep 30ksi Aluminum IM6/Ep IM6/Ep E-glass/Ep IM6/Ep

0.34 0.34 0.65 0.95 1.02 1.09 2.40 3.20 6.25

-0.10 -0.11 -0.02 -1.29 -0.30 0.35 -1.77 -2.07 -0.11

3.9 50.0 7.9 6.8 2.9 3.1 2.7 4.5 0.8

Fig. 30.12 Relative deformation of 10 x 1 x 0.1 in specimens under 1000 lb load along the centerline (laminates are IM6/epoxy, unless otherwise indicated).

702 Laminate design

B* matrix with nonzero terms. The first and fourth laminates are balanced and so the A*16 and A*26coefficients are zero. For the second and third laminates which differ by the sign of the off-axis plies, the stiffness behavior differs only in that the A*16,A*26,D*,6 and D*26coefficients are of opposite signs. Table 30.5 displays different quasi-isotropic laminates. Note that the normalized A* matrix

A four-ply laminate consisting of two 0" and two 90" plies can be combined into four different laminates. From Table 30.3 it can be observed that while the A* matrix remains unchanged through varied ply stacking sequences, large differences arise in the B* and D*matrices. From Table 30.4 it can be observed that only the fourth laminate is unsymmetric and has a

Table 30.3 Normalized stiffness coefficients for four IM6/epoxy laminates, in units of Msi 190/0/0/901

10/90/90/01

Layup

' 15.624

[A*]

.15.624 0.525

0 ' 0.525 15.634 0 0 0 1.220,

0.525

0.525 35.634 0 0 0 1.220

~

0

0

0

0

0 ' 0

, o

0

o , ,

'

[B*]

26.129 0.525 0.525 5.138 0 0

0 0

.

1.220.

'

0

0

0

0

0

0

' 5.138

.

0.525 0.525 26.129

0

0

f 0/0/90/901

~0/90/0/901 '

15.624 0.525

15.624 0.525 0.525 15.634

0

,

:]I0

0

3.499 0 0 -3.499

0

0 ' 0 1.220,

0

' 15.634 0.525 0.525 15.634

0

,

0.525 15.634

0 1.220

0

0

0 -

- 6.997 0 0

0

o * , 0 0

0

0

1.220

0

0 ' 0 0 .

-6.997

0

. .15.634 0.525

1.220,

0 ' 0.525 15.634 0 0 0 1.220,

Table 30.4 Normalized stiffness coefficients for four IM6/epoxy laminates, in units of Msi Layup

[0/0/+45/-451$

10/0/+45/+451+

' 19.463 3.692 [A*]

,

0 ' 3.692 5.469 0 0 4.387. o

* o IB"1

0

0 0

, o

0

:I 0

27.087 1.316 0.656 1.316 2.597 0.656 0.656 0.656 2.012

' 19.463 3.692 3.499 3.692 5.469 3.499 , 3.499 3.499 4.387

0 0

J

o

0

0 0

[0/0/-45/-45Is

I

19.463 3.692 -3.4993.692 5.469 -3.499 .-3.499 -3.499 4.387

:ll! 0

27.087 1.316 0.875 1.316 2.597 0.875 0.875 0.875 2.012

f+45,/45,1

0 0 0

.

0 6.859 9.297 0 , 0 0 7.555 * 9.297 6.859

: [:

0 3.499 0 3.499 3.499 3.499 0

0

27.087 1.316 -0.875 1.316 2.597 -0.875 -0.875 -0.875 2.012

II

9.297 6.859 6.859 9.297 0

0

0 0

7.555

I

Laminate design 703 Table 30.5 Normalized stiffness coefficients for four IM6/epoxy laminates, in units of Msi

0 12.466 3.692 3.692 12.466 0 0 0 4.387

20.932 3.098 1.312 3.098 5.188 1.312 1.312 1.312 3.793

I I

12.466 3.692 0 3.692 12.466 0 0 4.387 0

1[

20.2115 2.504 1.968 2.504 7.094 1.968 1.968 1.968 3.200

12.466 3.692 0 3.692 12.466 0 0 0 4.387

12.466 3.692 0 3.692 12.466 0 I O 0 4.387

18.194 2.782 3.735 2.782 8.558 2.208 3.735 2.208 3.477

22.001 1.932 0.737 1.932 6.451 1.956 0.737 1.956 2.628

terms are equivalent for all quasi-isotropic laminates. This means all have the same stiffness to weight ratio. The differences between these laminates thus manifests themselves only in how they behave in bending.

(iii) Evaluate the displacements by integrating the strains. (Note that for the tube,

PROBLEM

1

Find the amount that an anisotropic 20-layer [0/30], T300/5208, 3 in diameter, 12 in tube tube will extend, change in circumference, and twist under an in-plane load, N, = 100 lb/in.

Multiplying out the strains, a12N2/

= a16N1

‘6

E

I

= a , , ~ , , E2 =

dd) dx

=r-.)

Eldx = l:$dx

= l:llNldx

E2dy=$$dC

=[a12NldC

-

u = allNIL

---t

z,

= ul2N12nR

SOLUTION (i) Compute the laminate stiffness matrix and invert to get the compliance matrix:

I

21.159 2.567 3.935 2.567 2.476 1.458 , 3.935 1.458 3.195

I

62.563 -26.647 -64.893 [a*] = -26.647 563.836 -224.518 -64.893 -224.518 495.374

L

1 R

-

+@ = a , 6 ~ l R

(iv) Evaluate the displacements numerically: =-a*ll N L = - 62.563 x 100 x 12 h 0.1

I

= 0.75 x 10-3in

v = -Nl2nR a*,, h

= -___ 26*M7 x 100 x 2 x 3.14 x 1.5

0.1 = -0.25 x 10-3 in

(ii) Evaluate the strains: @

= -U*16 Nh

L = ‘R

= -0.52 x

64 893L x 100 x 12/1.5

0.1 radian = -0.03 degree

704 Laminate design 30.5.4 CTE BEHAVIOR

the a1 is -2 to -2.5 (versus -0.14 for a1 of the The following four figures show the coefficient unidirectional tape, as shown in Fig. 30.13(a)). As the number of ply angles increases, the of thermal expansion (CTE) in two principal directions (referred to as a , and a,) for CTE behavior becomes less intuitive. Figure M40J/F584 carbon fiber laminates. Figure 30.14(a) shows the CTE of a laminate with 50% 30.13(a)shows the CTE of an off-axis unidirec- 0 plies and 50% angle-ply; Fig. 30.1303) shows tional ply; Fig. 30.1303) shows the CTE of an the CTE of a laminate with 25% 0 plies, 25% 90 angle-ply laminate. From Fig. 30.13@), it can plies and 50% angle-ply. Note that when & be observed that, due to the Poisson coupling equals 45", the resulting laminate is quasieffect, laminate CTE values less than that of a isotropic [0, 90, &45] as confirmed by a , unidirectional material are possible for specific equaling a*. Examination of the fundamental ply angles. In Fig. 30.13(b), for 0 of GO to do", trends in Figs. 30.13 and 30.14 indicates

16.00

16.00 14.00

14.00

12.00

1200

10.00

10.00

8.00

am

-

6.00

600

4.00

4.00

c

200

200

0 C

a, ._ 0

0

0.00 -200

-r

0.00

p

10

20

30

40

SO

60

70

80

90 -200

0


-4.00

(b)

Fig. 30.13 Coefficient of thermal expansion of M40J/F584 laminates as a function of 8 for (a) off-axis undirectional [e,],; and (b) angle-ply [&J,.

4.00 0 c

c

200

--Ply Angle, B (degrees)


(4

(b)

Fig. 30.14 Coefficient of thermal expansion M40J/F584 laminates as a function of 0 in the following laminates (a>[o,, and (b) IO, ,90,, 4 1 , .

Laminate design 705 potential near-zero CTE laminates, particularly useful in spacecraft applications to minimize deformation due to the large cyclic thermal loading.

EXAMPLE To remove a composite shaft from a metal mandrel after elevated temperature cure, the laminate CTE in the hoop direction of the cylindrical section has to be less than that of the mandrel material to prevent lock-on. The composite is considered to be stress-free at cure-temperature, and thus the temperature loading is associated with the temperature decrease to room-temperature.

dominantly in the longitudinal direction to accommodate flexural loading (like a mast or golf club), Fig. 30.14(a) indicates that a [O,, G0JS with a steel mandrel would be a problematic choice, resulting in a composite shaft locked on to the mandrel as shown in Fig. 30.15@).There is a need for sufficient fibers in the hoop direction (90") to result in a laminate CTE less than that of the mandrel material. The CTE of the metal materials given in Table 30.6 indicates that it is easier to remove a composite shaft from an aluminum mandrel than from a steel mandrel.

Table 30.6 Average coefficients of linear thermal expansion of selected materials

Figure 30.15(a) illustrates that it is preferable Aluminum alloy to have the metal mandrel contract more than Concrete the composite during cool-down, which Invar Steel means that the metal CTE must be more than Titanium alloy the composite CTE. For a shaft with fibers pre-

a composite ' aofmandrel (significantamount fibers in hoop direction)

12.8 6.7 0.39 6.5 4.9

a composite ' a mandrel (predominantly longitudinalfibers)

comDosite

AT =

wp [1 -&)lP,(y)

(32.16)

(b) Plate whose longitudinal edges (y = 0 and

y = b) are simply supported and transverse edges (x = * a/2) are clamped W(X1Y) =

wp 11-f,(X)lP,(y)

c parameters in these equations have the form b

c1 = p d Y

(32.17)

(c) Plate with clamped longitudinal and simply supported transverse edges

b

c2 = l0(PWY b

wp [1 -f,(41P2(y) (32.18) (d) Plate with clamped longitudinal and transverse edges W(XlY> =

w(x,y) = wp [1 -f2(X)IP2(y) (32.19) The following are used:

c3

= l0(Pf7ZdY b

c = ITdY For solutions in eqns (32.16) and (32.17) it should be taken c, = 0.04921 b9, c, = 0.48571 b7, c3 = 4.8b5, c = 0.2 b5. For solution in eqns (32.18) and (32.19) c, = 0.001587 b9, c2= 0.01905 b7, c3= 0.8 b5, c = 0.03333 b5. 32.5.2 BUCKLING AND POST-BUCKLING BEHAVIOR OF SYMMETRICALLY LAMINATED PLATES

In-plane compression or shear (Fig. 32.17) (where the forces Tx, T,, Tx, are uniformly

Rectangular plates 751 (b) if the edge y = 0 is simply supported and the edge y = b is free, then

Under pure shear (Tx= Ty = 0), the critical load for an orthotropic symmetrically laminated plate is expressed as 7e

Tx; = k--.\j(D,,D,) ab Fig. 32.17 In-plane loading of a rectangular plate.

(32.22)

Coefficient k is given in Table 32.1 for typical values of the following parameters:

distributed along the plate edges) can result in plate buckling. For a simply supported rectangular (a 2 b) plate under uniaxial compression (T, = Txy= 0), the critical load is expressed as Since the value of the critical load for an (32.20) orthotropic plate does not depend on the direction of shear forces, the parameter p in where Table 32.1 can be replaced by 1/ p , so that Table 32.1 presents coefficients k as p varies from k = 2 1 + D12 + 2033 (32.21) 0.02 to 50. Critical combination of compressive and Note that eqn (32.21) is valid if the longitudi- shear forces can be determined using the folnal edges of the plate (y = 0 and y = b in Fig. lowing equation: 32.17) can experience displacement along the 2 T X =I y-axis. If these edges are fixed in this direction (which is often the case), then the compressive forces Tx give rise to transverse compressive where T,' and Tx; are specified by eqns (32.20) forces Ty = vq T, due to Poisson's effect. Then and (32.22). Used as the skin elements of stringer panels [DllA + 2(D1, + 2D,,) +-k = or shear webs, composite plates (just as metal A 1 + (vx,/4 ones) can sustain high compressive or shearshould be substituted into eqn (32.20) and ing loads after buckling. However, in contrast minimized with respect to 1 = (rnl/a)2 where rn to metal panels whose ultimate loads are usuis the number of half waves in the x-direction. ally determined by rib fracture, buckling The following approximate expressions are failure of composite panels (particularly made derived for the k coefficients with different from carbon-epoxy composites) is often plate edge supports: caused by skin fracture due to bending. (a) if the edges y = 0 and y = b are clamped, Therefore, traditional engineering methods of strength analysis, such as the method of then reduced width for compressed panels and the Dl2 + 2033 concept of diagonal stress field for shear webs, can hardly be used for composite panels; more

T,' = k--./(D,,D,) 7c2 b2

) TvJ

q+[+]

752 Analysis methods Table 32.1 Buckling coefficient k for pure shear

P q 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.2 3.4 3.6

0.02 0.04 25.5 28.9 32.2 35.3 38.4 41.3 44.1 46.8 49.4 51.8 54.2 56.4 59.3 62.1 67.8 70.6 73.4

18.4 20.8 23.1 25.4 27.5 29.6 31.5 33.4 35.2 36.9 38.5 40.1 42.1 44.0 48.0 50.0 52.0

0.06 15.3 17.3 19.2 21.0 22.8 24.4 26.0 27.6 29.0 30.3 31.6 32.8 34.4 36.7 39.3 40.9 42.5

0.08 13.5 15.2 16.9 18.8 20.0 21.4 22.8 24.1 25.4 26.6 27.7 28.8 30.8 31.4 34.2 35.6 37.1

0.10 12.4 13.9 15.4 16.8 18.1 19.4 20.7 21.8 23.0 24.0 25.0 26.0 27.1 28.3 30.8 32.1 33.4

0.14 10.9 12.2 13.4 14.6 15.8 16.9 18.0 19.0 20.1 21.1 22.0 23.0 24.1 25.3 27.4 28.5 29.5

0.18

10.4 11.5 12.4 13.6 14.7 15.7 16.7 17.7 18.6 19.5 20.4 23.1 22.3 23.3 25.1 26.0 26.8

rigorous solutions of the corresponding nonlinear problems are required. 32.5.3 NONSYMMETRICALLY LAMINATED PLATES

In contrast to symmetrically laminated plates, bending of plates with an arbitrary stacking sequence of the layers is accompanied by stretching of the basic surface. Also, the plate deflections depend on boundary conditions imposed on the inplane displacements2. For in-plane compression, nonsymmetrically laminated plates experience bending which should, in general, be described by nonlinear equations. This longitudinal bending can be unstable and can be usually accompanied by the so-called mode-jumping. 32.5.4 AXISYMMETRIC DEFORMATION OF

CIRCULAR PLATES AND DISKS

The problem of axisymmetric bending and inplane deformation of orthotropic composite plates and disks (in cylindrical coordinates Y,

0.24

0.32

0.40

0.50

0.60

0.80

1.00

9.61 10.4 11.5 12.4 13.3 14.2 15.2 16.0 16.8 17.8 18.4 19.3 20.3 21.0 22.8 23.7 24.5

8.40 9.27 10.2 11.0 11.9 12.8 13.6 14.5 15.4 16.4 17.1 17.9 18.8 19.5 21.3 22.1 22.9

7.50 8.36 9.24 10.1 11.0 11.8 12.7 13.7 14.4 15.2 16.1 16.7 17.9 18.7 20.4 21.2 22.0

6.77 7.66 8.52 9.39 10.4 11.2 12.1 12.9 13.7 14.6 15.4 16.2 17.1 17.9 19.7 20.4 21.2

6.32 7.20 8.08 9.05 9.93 10.8 11.6 12.4 13.3 14.1 14.9 15.8 16.7 17.4 19.1 19.9 20.7

5.87 6.73 7.68 8.55 9.45 10.3 11.1 12.0 12.9 13.8 14.6 15.4 16.3 17.1 18.7 19.6 20.3

5.76 6.66 7.56 8.47 9.34 10.1 11.0 11.9 12.8 13.7 14.5 15.3 16.2 17.0 18.7 19.5 20.2

p, z) is reduced to the following set of

equations in terms of radial displacement of the basic surface ut, deflection w,and rotation of the normal to the basic surface 6 , i.e. ?ut’”’

+ ~ Y ~ U +~ (7 “ ’ - np2 - n2)Yur” + (1 - np2

- n:)ur’

-

I

+ (n?:

- n 4)2 =‘ [ -(C ‘ 1 C

Y

Bll Dll

p’rdr) - Y ( Y F ~ )-” ( Y F ~ )+’ n;F,

1

(32.23)

1 In r Kll Here, ()’ = d()/dr, j7 = p - q (see Fig. 32.3), Fr(r) is a radial body force (e.g. a centrifugal force for a spinning disk), C,, C, are constants of integration, and

w

= -(Cll

Cylindrical shells 753 nb2

=

32.6 CYLINDRICAL SHELLS

D22 ~

Dll

n: =

___ c2;

BllDIl Coordinates of the basic surface e = I $ ) / I z ) provide, in accordance with eqn (32.12), the zero value of the radial coupling stiffness (Cll = 0). Radial and circumferential strains at an arbitrary point of the plate are expressed as

er = u,’ 1 ep = -(ur r

’,

a j

h

+ zOrf + z0,)

The general solution for eqn (32.23) has the form

1c p + 6

ur =

Filament wound composite cylindrical shells are used as pressure vessels, reservoirs, pipes, aircraft and ship elements. The governing equations for a cylindrical shell (Fig. 32.18) can be obtained from eqns (32.1)-(32.3) if we take = = l / R i = or ‘ 2 = and rep1ace 7 with yTz.

Fig. 32.18 Cylindrical shell.

up

1=3

where up is a particular solution and si are the roots of the equation

s4- (nP

+ n,’)s2 + np2n:

- n:

= 0

Six constants of integration can be found from the corresponding boundary conditions according to which ur = w = Or = 0 for a clamped edge, w = Nr= M , = 0 for a simply supported edge, w = ur = Mr = 0 for a hinged edge fixed in the radial direction, and Nr= Mr = Q, = 0 for a free edge. To write the force boundary conditions in terms of displacements, the following expressions can be used: Ur

O

Nr= Bllu,’ + B 1 2 T + C122 r

32.6.1 AXISYMMETRIC DEFORMATION

One of the most important loading cases for cylindrical shells is the axisymmetric loading with pressures p , q and axial forces N (Fig. 32.18). In this case, the equations account for the first-order nonlinear effects of the axial forces on the curvature of deformed shell meridian. These equations have the following form:

M,’- Q, = 0 N Q,‘+ Nw” - 2 + j7 = 0 R N,= N W N, = B,,u‘ + B l2R

U

0

M, = C12& +DllBr’+ D 12 r

_f_

W N, = B,,u‘ + B 22 + C,,Ox’ R

M, = C

W

-

l2

R

+ D,,OX’

754 Andysis methods

Q,

= K,,(ex + w')

(32.24)

where ( )' = d( )/dx, j? = p - q. Stiffness coefficients B, C, D, K are specified by eqn (32.6) in which e = I (,)/I(0) 11

Boundary conditions should be written in terms of w and force Sx = Nw'.

11

The foregoing set of equations, (32.24), can be reduced to the ordinary differential equation

32.6.2 NONSYMMETRIC DEFORMATION

In the general case of loading, composite cylinw""- 2s2w"+ PW = kp (32.25) drical shells can be usually described rather adequately by the so-called semimembrane where theory that, in addition to membrane theory, takes into account the circumferential bending s2 = C,, (1 + C ) + RN + ___ moments. The model of a semimembraneshell 2RCD can be represented by a system of rings with inextensible axes that take only circumferenB tial bending moments and by a system of t4 = R2B,,CD absolutely flexible beams that connect the rings and sustain axial and shear forces only. The semimembrane theory assumptions lead to the following equations: ~

=

'11['

+

C,, + RN RCK,,

1

aNx -+ax

a%, = o ay

a%, %+ay ax

+ - Q, +q R

Y

=o

B = B,,B, - B,:

Constants of integration entering the solution of eqn (32.25) can be found from the corresponding boundary conditions according to which w = Ox = 0 for a clamped edge, w = M x = 0 for a simply supported edge, and M x = S , = 0, where Sx= Q, + Nw' for a free edge. Thin-walled composite pressure vessels can be described by nonlinear membrane equations using the assumption that D,, = C, = 0. These equations can be reduced to W" - k2W = kp

Q,

where E

=

K,VY

a u w =-+-=(I

' a Y R

Cylindrical shells 755 8 Y

= q + - -v - aw R ay

(32.26)

where j7 = p - 9, 9Y is the circumferential surface traction, and stiffnesses B, D are specified by eqn (32.6) in which e = I$)/I$). Decomposition into Fourier series, i.e.

= Nx; = 0 for a free edge. The following expressions can be used for the boundary conditions Un =

BllR

,,,

‘11 wnr------w,

~

A,2B3:

N,”= BllR wn - R4 E1 , ~

m

B33

n=l

Nx; =

...! Ny”(x),My”@),p,(x)lcos Any

Wnr

-__

mnz

m t F 3 3

BllR

5 1 ~

m,3

Wnlfr

-

~

An’s3

I,

Z~~ w n’

where

m

It must be noted that the semimembrane theory is not valid for the case of axisymmetricloading whereAn= n/R, allows reduction of eqn (32.26) (n = 0) and reduces to the membrane theory for to the following governing equation: n = 1. Proper combination of solutions of eqn (32.25) for n = 0 and of eqn (32.27) for n 2 1 w ~ ” ”- 2s’~~’’ + t4wn= kp (32’27) allows consideration a wide range of practical where problems for composite cylindrical shells.

..., Qy”(x),qy”(x)lsinAny

52

=

A;(n2 - l)D, 2B33

32.6.3 BUCKLING

Under axial compression by forces N (here, in contrast to Fig. 32.18, N are compressive forces), cylindrical shells can experience three modes of buckling: column-type, axisymmetric, and nonsymmetric buckling. The actual critical load is the smallest of the three values. For a column-type buckling

Nc =

n2rn2R2B 2L2B2,(1 +

n2rn2R2B L2B22B33

)

where B = Bl1B2, - B1: and rn depends on the character of end fixity. If the end cross sections n are fixed in such a way that they can freely Four constants of integration entering the rotate (hinged column), then rn = 1.If the ends solution of eqn (32.27) can be found from the are clamped, then rn = 2. corresponding boundary conditions according Axisymmetric mode of buckling is typical to which u, = vn= 0 for a fixed edge (it is essen- for thick and sandwich shells. The corretial that inextensibility condition E = 0 yields sponding critical load is specified by the w,, = -RAnvn,so wn = 0 for a fixed edge) and N,” following equation that allows for transverse

Fn = P, -

9;

756 Analysis methods shear and radius variation through the shell

where

17 I=[."." BllD,

+=)I

+{(l

BP22

where

=(y) 2

1 ;

For a homogenous shell

Stiffnesses B, C, D are specified by eqns (32.6) and (32.7) in which e = I$)/Iio); Am,,should be where changed for Am,, where Ex= Ex/(l - vX,,vy,). (i) - A (1) A (11 = A (i), 421 = A ~ ( ~ ) / S ~ - i 11 12 12 The critical value of the lateral external presand sure can be approximated by 1 si = 1+ - (ti + ti-,-2e) 2R

41

Transverse shear stiffness is

Note that the shell is assumed to be simply supported at x = 0 and x = L (see Fig. 32.18). If transverse shear deformation is not taken into account, then

The critical load, corresponding to the general mode of buckling of a thin simply supported orthotropic shell, can be found as

N, = DllA:,,

+ D, 2 II", + R2Ai($+%]

Here, B,, and D, are specified by eqn (32.6) in which e = I$)/I$) and parameter c depends on the boundary conditions. For a simply supported shell c = 1,for a hinged shell whose end cross sections cannot move in the axial direction c = 1.5, and for a shell with one end hinged and the other end free c = 0.6. Buckling pressure for an infinitely long shell is

9, =

3D2,

x3

Finally, note that the derivation for equations presented in this chapter can be found elsewhere3.

References 757 REFERENCES 1. Tarnopolskii, Yu.M. and Kincis, T.Ya. 1985. Static Test Methods for New Van Nostrand Reinhold.

2. Whitney, J.M. 1987. Structural Analysis of Laminated Anisotropic Plates. Lancaster, Pensylvania: Techomic Publishing co,,Inc, 3. Vasiliev, V.V. 1993. Mechanics of Composite Structures. Washington: Taylor & Francis.

DESIGN ALLOWABLES SUBSTANTIATION

33

Christy Kirchner Lapp

33.1 INTRODUCTION

Designing with composite materials requires knowledge of a significantly greater number of properties than for conventional isotropic metals. The selection of lamina and laminate allowables can be critical in the analysis of a composite structure. However, composite design allowables may not always be obtained from a single source of data. Several references must often be consulted to determine all the properties in the necessary directions, especially if several fibers or matrices are being considered in the design. This can be a timeconsuming effort, especially during the initial design phase. In addition, some organizations may not have easy access to all the necessary references required to collect the data. This chapter assembles lamina data from numerous sources so that the engineer may have a single reference point for initial design and analysis of composite structures. A broad range of fibers has been included since composites are increasingly being applied outside the aerospace community. Design allowables in this section are for both elastic and strength properties. Elastic properties are necessary for laminate design or the analysis of composite structures. These properties include elastic moduli and Poisson’s ratios. Strength properties are required to predict laminate strengths or perform a failure analysis of the structure.


Strength properties for compression, tension and shear must be determined. Both elastic and strength properties can be influenced by numerous variables such as the fiber, matrix, fiber volume and processing method. The test method used to determine the allowable can also affect the property. In addition, some allowables cannot be readily tested, especially properties through-the-thickness of the composite. Often, complete characterization of a fiber/matrix system may not be available and the engineer must estimate or assume properties. It can be expensive and time-consuming to completely characterize a fiber/matrix system during the initial design phase, so it is important for the engineer to have allowables based on reliable data and to understand the limitations. This chapter concentrates on providing the material database and techniques for assembling the necessary composite allowables for preliminary design. Final design allowables may require additional testing. The allowables provided in this chapter should be considered preliminary design values. Allowables for both two dimensional (2-D) and three dimensional (3-D) properties are included. As a starting point, lamina allowables for commonly used fiber/matrix systems are described and listed. The references for these allowables are included. Several references may be listed for a single material. Methods for estimating properties when data is not available or testing cannot be readily performed are also defined. The effects of processing methods on allowables and methods

Lamina allowables 759 for adjusting lamina properties for these vari- 33.3 LAMINA ALLOWABLES ables is described. The intent is to provide a Lamina allowables may be used in a laminated baseline for design allowables which can then plate code to predict laminate elastic and be expanded with additional testing data as strength properties, or they may be used required or modified for a specific application. directly in a finite element analysis code. Table 33.1 defines the lamina properties and the basis for each property. Most of the properties 33.2 NOMENCLATURE FOR DESIGN are based on test data for the 2-D properties. ALLOWABLES The 3-D properties typically represent calcuThe nomenclature used to describe composite lated values based on the equations shown in lamina and laminate properties is not consis- Table 33.1. tent within the industry. For this chapter, the Properties for commonly used fiber and lamina properties parallel to the fiber are spec- epoxy resin systems are included and the conified with a 1 and the lamina properties struction is assumed to be continuous fiber transverse to the fiber are designated with a 2. reinforcement. A wide variety of reinforcing Figure 33.1 illustrates this nomenclature for fibers is included since composites may be composite lamina and laminate properties. used in variety of applications, each with very The properties in the 1-2 plane are often different design requirements. Glass, polyethreferred to as the in-plane lamina properties ylene, aramid and graphite fibers are covered. whereas the properties through-the-thickness Graphite fibers include standard, intermediate (out-of-plane)of the composite are designated and high modulus types. The fiber/resin syswith a 3. The 1, 2 and 3 directions can be tems have very different properties and the referred to as the longitudinal, transverse, and actual composite application will dictate the through-the-thickness properties respectively. selection. For example, glass fibers are less expensive and more impact resistant than Laminate property definition

X

2

amina property definition

Fig. 33.1 Nomenclature for lamina and laminate properties.

760 Design allowables substantiation Table 33.1 Definition of lamina properties and equations used to calculate material properties

Lamina material properties

Equation used to calculate material property

Definition

Elastic El

E2 E3

GI2 G23

G13 v12

Elastic modulus in the fiber direction Elastic modulus transverse to the fiber direction Elastic modulus through-the-thickness Shear modulus in the 1-2 plane

Property based on test data Property based on test data Transverse isotropy: E, = E, Property based on test data E3

Shear modulus in the 2-3 plane

G23

2(1

=

+ YB)

Shear modulus in the 1-3 plane Poisson's ratio in 1-3 plane

Transverse isotropy: G,, = G,, Property based on test data

Poisson's ratio in 1-3 plane

Transverse isotropy: v13= vI2

Tensile strength in the fiber direction Compressive strength in the fiber direction Tensile strength transverse to the fiber Compressive strength transverse to the fiber Tensile strength through-the-thickness Compressive strength through-the-thickness Shear strength in 1-2 plane (in-plane) Shear strength in the 1-3 plane (interlaminar) Shear strength in the 2-3 plane (interlaminar)

Property based on test data Property based on test data Property based on test data Property based on test data u3 = u2 a3= a, Property based on test data Property based on test data

'23

'12 t13

'23

graphite fibers, but have a higher density and lower modulus. Since the lamina properties are assumed to be used as preliminary design and analysis parameters, the effects of temperature, environment and fatigue are not considered. However these conditions must be considered during the final design development. 33.3.1 TWO-DIMENSIONAL (2-D) LAMINA PROPERTIES

The elastic lamina properties required for a composite 2-D analysis are typically E,, E,, G,,, and vI2. The specific properties may depend

'23

=

'13

upon the type of analyses being performed or the analysis code being used. If a strength or failure analysis is performed, then the following strength allowables may be required; q, o,,-q,-0,and T,,. Tables 33.2 and 33.3 include these elastic and strength properties, which were compiled from various sources, such as military standards, material supplier data and published literature. References are included for each property so that the engineer may directly consult a particular reference if further information is required. These references are listed in Tables 33.4 and 33.5. The lamina properties for glass, polyethylene and aramid fibers are listed in Table 33.2.

Lamina allowables 761 Polyethylene and aramid fibers are more commonly known by their trade names as Spectra and Kevlar. Kevlar 29 and 49 are included for aramid properties. The polyethylene fibers include Spectra 900 and 1000. E-glass and Sglass (trade names) are included for glass fibers. Table 33.3 lists lamina properties for standard, intermediate and high modulus graphite fibers. Data for non-USA produced fibers is not included. The type of material system is also indicated in Tables 33.2 and 33.3. Design allowables for a prepreg material system versus a wet filament wound system may result in different properties. It is important to consider the processing method when selecting the properties to be used in a design. One processing method may result in a lower moduli or strength than another. The effects of different processing methods on design allowables is discussed in further detail at the end of this chapter. The majority of the 2-D properties is based on test data, not micromechanics equations. The test data is typically based on 'thin specimens' (typically less than 2.54 mm (0.10 in) thick). Properties based on thin specimens may not represent those for thick composites (typically greater than 6.35 mm (0.25 in) thick). The exact definition of a thick composite is not consistent within the composites community. Thick composite structures may have properties lower than those of thin composites, so the engineer may need to perform additional testing for certain applications.

these properties are difficult to test and data is not always readily available. The 3-D lamina properties listed in Tables 33.2 and 33.3 are estimated based on 2-D properties. The following section describes the methods for calculating lamina properties in the 3-direction when data is not available. 33.3.3 ESTIMATING LAMINA PROPERTIES WHEN DATA IS NOT AVAILABLE

Lamina properties through-the-thickness (3direction) are often not readily available, although they may be required to perform an analysis. When data is not available, these properties can be determined by assuming that the lamina is transversely isotropic. For a transversely isotropic lamina, the properties in the 2- and 3-directions are assumed to be the same. Thus the following equations may be used to determine elastic properties in the 3direction':

E, = E,

G*, = G12 r L3 G23

= 2(1 + vz3) '13

= '12

The transverse Poisson's ratio, v23, can be determined from the following relationship2:

33.3.2 THREE-DIMENSIONAL (3-D) LAMINA PROPERTIES

Irutially the majority of analyses performed on composite structures were two dimensional. However with advancement of finite element analysis programs and as new applications for composites arise, more analyses are being performed for the 3-D case. Thus through-the -thickness lamina properties have become necessary in performing certain analyses. Often

where vf is the fiber Poisson's ratio, V , is the fiber volume fraction, vm is the matrix Poisson's ratio, and Em is the matrix elastic modulus. Therefore to calculate certain properties in the 3-direction, the engineer needs to know certain fiber and matrix properties, and these have also been listed in Tables 33.2 and 33.3. For initial design purposes, it is simplest

762 Design allowables substantiation

to assume that the fiber and matrix are isotropic, although certain fibers are considered to be anisotropic. The fiber or resin shear modulus can be calculated by: E G=----'-2(1 + v) The following equations may be used to estimate strength properties in the 3-direction. u3= u2

-a3= -u2 The equations listed above should be considered a starting point for estimating 3-D lamina properties when actual test data is not available. The values may need to be verified by testing as the design progresses. 33.4 LAMINATE ALLOWABLES

Laminate elastic properties and strengths can be determined by testing or by using a laminated plate code. During the initial design phase, laminate allowables are typically determined by using a laminated plate code. This is especially true if the composite lay-up deviates from a 'standard' lay-up, such as a quasi-isotropic laminate ([90, +45, 4 5 , OIJ. There is typically more test data available for a quasi-isotropic lay-up than any other lay-ups. Laminate testing is often performed after completion of the initial design, material selection and composite lay-up has been decided. Laminate testing would be performed to confirm predictions and processing effects. 33.5 EFFECTS OF PROCESSING VARIABILITIES ON DESIGN ALLOWABLES

The actual fabrication method used to build a composite structure can have an impact on the design allowables. Some processing methods can result in a higher fiber volume fraction and lower void content than others. For example, an autoclave cured part using prepreg tape will typically have a higher fiber volume

and lower void content than a wet filament wound part. This can affect properties such as the elastic modulus (E,) or tensile strength (uJ. It is important that the engineer understand the limitations of the selected processing method and adjust the design allowables accordingly. Processing parameters which can affect the lamina allowables are fiber volume fraction and void content. Design allowables should be modified if the process used to determine the lamina properties deviates from the intended process for the final composite part; this ensures that unrealistic properties are not being used to design the part. There are numerous fabrication methods applicable for composites. The methods which will be discussed are those that apply to continuous fiber reinforcement. These include filament winding, hand lay-up and resin transfer molding (RTh4). The method of cure can also affect the lamina properties. For example, one part can be filament wound with prepreg tow and another can be hand layed up with prepreg tape. These parts have different fabrication methods, but they may be cured in the same manner; in an autoclave with vacuum and pressure. It is very likely that these two parts would have similar properties and require no adjustment of lamina properties. However, if the filament wound part was wet wound and cured in an oven without vacuum or pressure, then the lamina properties would need to be adjusted if they were based on properties derived from testing using prepreg tape. A wet filament wound part typically has a lower fiber volume fraction and higher void content. Thus in determining if lamina properties need to be modified, the engineer must consider the complete method of processing, including the raw material and cure method, not just the automated or manual process which is being used to fabricate the part. In general, if a part is wet filament wound, it will possess a lower fiber volume fraction and a higher void content than a hand-layed up part using prepreg tape and cured in an autoclave. Wet filament wound parts cured in

References 763 an oven without vacuum and pressure typically have a fiber volume fraction between 0.55 and 0.60 with a void content between 1 and 5%. Parts fabricated from prepreg tow or tape, which are cured in an autoclave with vacuum and pressure will typically contain fiber volume fractions between 0.60 and 0.65. Parts which are fabricated using the RTM process will typically possess fiber volume fractions of approximately 0.50. Determining lamina properties for RTM parts is particularly difficult since the preform is usually woven and properties are not readily available. Also weaving in some conditions may slightly degrade the properties. The adjustment of lamina properties should focus on the 2-D elastic properties; E,, E,, G,,, v,, and the 2-D strength properties; u,, -ol, u,, -pz and rl,. The simplest method is to adjust the desired property by multiplying the value by the ratio of the fiber volume for the selected processing method to the fiber volume listed in Tables 33.2 and 33.3. For example, if one has properties based on prepreg tape which has been autoclaved cured and wants to adjust these properties for a wet filament wound part the following calculation would be used to adjust the longitudinal elastic modulus (El):

E , (wet filament wound)

V,(wet filament wound) Adjusting properties by the ratio of fiber volume fraction is applicable for modulus, tensile and compressive strength, but does not serve well for Poisson’s ratios which would require micromechanics. The Poisson’s ratio can be calculated based on the following equation,: Y1,

=

v,v,+ vm(l - V,)

where vf is the fiber Poisson’s ratio, Vf is the fiber volume fraction and vm is the matrix’s Poisson’s ratio. REFERENCES Whitney, J., Daniel, I. and Pipes, B., 1984.

Experimental Mechanics of Fiber Reinforced Composite Materials. Brookfield Center, Connecticut: The Society for Experimental Mechanics. Vinson, J. and Sierakowski, R. 1987. The Behavior

of Structures Composed of Composite Materials. Boston: Kluwer Academic Publishers.

764 Design allowables substantiation Table 33.2 Lamina properties for glass, aramid and polyethylene fibers in epoxy matrices

Properties

E-glass/ DER 332

S2-glass/ DER 332

S2-glass/ xP251 s

Kevlar 29/934

Material system Fiber type Supplier Resin type Supplier Fiber volume Composite density, g/cm3 (ib/in3)

Wet wound E-Glass Owens Epoxy Dow 60% 2.05 (0.074)

Wet wound S-Glass Owens EPOXY Dow 60% 1.98 (0.072)

Prepreg S2-Glass Owens EPOXY 3M 60% 1.98 (0.072)

Prepreg Aramid DuPont EPOXY Fiberite 58% 1.38 (0.050)

51 (7.5) 17 (2.5) 17 (2.5) 7 (0.98) 7 (0.95) 7 (0.98) 0.25 0.32 0.25

54 (7.9) 5 (0.7) 5 (0.7) 2 (0.24) 2 (0.24) 2 (0.24) 0.40 0.47 0.40

Lamina elastic properties E,, GPa (psi x lo6) E,, GPa (psi x lo6) E,, GPa (psi x lo6)

G,,,GPa (psi x lo6) G,,,GPa (psi x

lo6)

G,,,GPa (psi x IOh) '12 '23 VI i

48 (7.0) 12 (1.8) 12 (1.8) 6 (0.84) 5 (0.70) 6 (0.84) 0.19 0.26 0.19

54 (7.9) 16 (2.3) 16 (2.3) 7 (0.98) 6 (0.89) 7 (0.98) 0.25 0.32 0.25

Lamina strength properties 01

(72

ff3

212

Tension, MPa (psi x lo3) Compression, MPa (psi x lo3)

1613 (234) 462 467)

1779 (258) -641 493)

1069 (155) -272 439)


39 (54 -103 -05)

58 (8.4) -186 -(27)

9 (1.3) -130 -09)


39 (5.6) -103 415)

58 (8.4)

9 (1.3) -130 -09)

MPa (psi x

lo3)

(3.3)

28 (4.0)

-186

427) 75 (10.9)

37 (5.3)

Tables 765

Kevlar 49/934

Kevlar 49 /DER332

Kevlar 149/934

Spectra 9OO/EPON 826

Spectra 1000/EPON 826

Pprepreg Aramid DuPont EPOXY Fiberite 58% 1.38 (0.050)

Wet Wound Aramid DuPont EPOXY Texaco 60% 1.35 (0.049)

Prepreg Aramid DuPont EPOXY Fiberlite 58% 1.38 (0.050)

Wet Wound Polyethylene Allied EPOXY Shell 55% 1.12 (0.040)

Wet Wound Polyethylene Allied EPOXY Shell 55% 1.12 (0.040)

72 (10.5) 5 (0.7)

5 (0.7) 2 (0.24) 2 (0.24) 2 (0.24) 0.41 0.48 0.41

82 (11.9) 5 (0.7) 5 (0.7) 2 (0.26) 2 (0.27) 2 (0.26) 0.31 0.38 0.31

106 (15.4)

6 (0.9) 6 (0.9) 2 (0.24) 2 (0.32) 2 (0.24) 0.34 0.42 0.34

31 (4.5) 4 (0.5) 4 (0.5) 1 (0.21) 1 (0.21) 1 (0.21) 0.32 0.40 0.32

50 (7.3) 1 (0.1) 1

(0.1) 1 (0.10) 0 (0.05) 1 (0.10) 0.28 0.36 0.28

1151 (167) -281 441)

12 (1.7) -134 419) 43 (6.3)

24 (3.5)

49 (7.1)

24 (3.5)

17 (2.5)


766 Design allowables substantiation Table 33.2 continued Lamina properties for glass, aramid and polyethylene fibers in epoxy matrices E-glass/ DER 332

S2-glassl DER 332

S2-glass/ XP251S

Kevlar 29/934

MPa (psi x lo3)

66 (9.5)

66 (9.5)

77 (11.1)

34 (5.0)

MPa (psi x

66 (9.5)

66 (9.5)

77 (11.1)

34 (5.0)

72 (10.50) 0.09 33 (4.8) 3103 (450) 2.60 (0.094)

87 (12.60) 0.18 37 (5.3) 3792 (550) 2.49 (0.090)

87 (12.60) 0.18 37 (5.3) 3792 (550) 2.49 (0.090)

83 (12.00) 0.44 29 (4.2) 3620 (525) 1.44 (0.052)

3.4 (0.49) 0.35 1.2 (0.18) 64 (9.3) 1.22 (0.044)

4.1 (0.60) 0.35 1.5 (0.22) 83 (12) 1.30 (0.047)

Properties 213

r23

lo3)

Constituent properties Fiber E,, GPa (psi x lo6)

vf G, GPa (psi x IO6) Tensile strength, MPa (psi x lo3) Density, g/cm3 (lb/in3) Resin E , GPa (psi x lo6) 213

G , GPa (psi x lo6) Tensile strength, MPa (psi x lo3) Density, g/cm3 (ib/in31

3.4 (0.49) 0.35 1.2 (0.18) 64 (9.3) 1.22 (0.044)

3.4 (0.49) 0.35 1.2 (0.18)

64 (9.3) 1.22 (0.044)

Tables 767

Kevlar 49/934

Kevlar 49 DER332

Kevlar 149/934

Spectra 9OO/EPON 826

Spectra 1OOO/EPON 826

50 (7.2)

50 (7.2)

38 (5.5)

23 (3.4)

23 (3.4)

50 (7.2)

50 (7.2)

38 (5.5)

23 (3.4)

23 (3.4)

124 (18.00) 0.45 43 (6.2) 3620 (525) 1.44 (0.052)

124 (18.00) 0.28 48 (7.0) 3620 (525) 1.44 (0.052)

172 (25.00) 0.33 65 (9.4) 3448 (500) 1.47 (0.053)

117 (17.00) 0.30 45 (6.6) 2586 (375) 0.97 (0.035)

172 (25.00) 0.22 70 (10.2) 2992 (434) 0.97 (0.035)

4.1 (0.60) 0.35 1.5 (0.22) 83 (12) 1.30 (0.047)

3.4 (0.49) 0.35 1.2 (0.18) 64 (9.3) 1.22 (0.044)

4.1 (0.60) 0.35 1.5 (0.22) 83 (12) 1.30 (0.047)

2.8 (0.40) 0.35 1.0 (0.15) 83 (12) 1.30 (0.047)

2.8 (0.40) 0.35 1.o (0.15) 83 (12) 1.30 (0.047)

768 Design allowables substantiation Table 33.3 Lamina properties for graphite fibers in epoxy matrices

Properties

AS4/3501-6

Material system Fiber type Supplier Resin type Supplier Fiber volume Composite density, g/cm3 (ib/in3)

Prepreg Graphite Hercules EPOXY Hercules 60% 1.58 (0.057)

lM6/3501-6 Prepreg Graphite Hercules EPOXY Hercules 60% 1.55 (0.056)

IM7/3501-6

lM8/3501-6

Prepreg Graphite Hercules EPOXY Hercules 60%

Prepreg Graphite Hercules

1.57 (0.057)

EPOXY

Hercules 60% 1.58 (0.057)

Lamina elastic properties E,, GPa (psi x IO6)

143 (20.7)

159 (23.0)

159 (23.0)

186 (27.0)

E,, GPa (psi x lo6)

10 (1.4)

10 (1.4)

10 (1.4)

10 (1.4)

E,, GPa (psi x lo6)

10 (1.4)

10 (1.4)

10 (1.4)

10 (1.4)

GI,, GPa (psi x

lo6)

6 (0.85)

5 (0.71)

5 (0.72)

6 (0.80)

G,,GPa (psi x lo6)

3 (0.41)

3 (0.41)

3 (0.41)

3 (0.41)

G,,,GPa (psi x lo6)

5 (0.68)

5 (0.68)

5 (0.68)

5 (0.68)

0.30 0.52 0.30

0.30 0.52 0.30

0.30 0.52 0.30

0.30 0.52 0.30

'12

'23 '13

Lamina strength properties 01


2172 (315) -1558 -(226)

02

Tension, MPa (psi x lo3) Compression, MPa (psi x lo3) 03


59 (8.5) -186 427)

2413 (350) -1655 -(240)

2620 (380) -1862 -(270)

2689 (390) -1931 -(280)

Tables 769

T300/5208

T40/1962

T50/1962

P55/1962

P75/1962

Prepreg Graphite Amoco EPOXY Fiberite

Prepreg Graphite Amoco EPOXY Amoco




62%

62%

62%

62%

62%

1.60 (0.057)

1.60 (0.058)

1.72 (0.058)

1.72 (0.062)

(0.062)

141 (20.5)

172 (25.0)

241 (35.0)

241 (35.0)

338 (49.0)

9 (1.2)

10 (1.5)

7 (1.1)

8 (1.1)

7 (1.0)

9 (1.2)

10

7

(1.5)

(1.1)

8 (1.1)

7 (1.0)

6 (0.92)

7 (1.00)

6 (0.84)

5 (0.79)

6 (0.85)

3 (0.41)

4 (0.54)

3 (0.39)

3 (0.39)

3 (0.36)

5 (0.68)

7 (1.00)

6 (0.84)

5 (0.79)

6 (0.85)

0.30 0.52 0.30

0.33 0.40 0.33

0.28 0.35 0.28

0.34 0.41 0.34

0.30 0.37 0.30

1524 (221) -1482 -(215)

3241 (470) -1724 -( 250)

1413 (205) -965 -(140)

931 (135) -510 474)

965 (140) 441 464)

36 (5.2) -159 -P3)

69 (10.0) -159 423)

37 (5.3) -159 423)

33 (4.8) -159 423)

33 (4.8) -159 423)

36 (5.2) -159 423)

69 (10.0) -159 423)

37 (5.3) -159 423)

33 (4.8) -159 423)

33 (4.8) -159 423)


770 Design allowables subsfantiation Table 33.3 continued Lamina properties for graphite fibers in epoxy matrices

Properties 212

13

‘23

AS4/3501-6

lM6/3501-6

MPa (psi x lo3)

87 (12.6)

85 (12.3)

96 (13.9)

80 (11.6)

MPa (psi x lo3)

124 (18.0)

121 (17.5)

121 (17.5)

131 (19.0)

MPa (psi x IO3)

94 (13.6)

94 (13.6)

94 (13.6)

94 (13.6)

234 (34.0)

276 (40.0)

276 (40.0)

303 (44.0)

0.26 93 (13.5)

0.26 109 (15.9)

0.26 109 (15.9)

0.26 120 (17.5)

IM7/3501-6

lM8/3501-6

Constituent properties Fiber E , GPa (psi x lo6) Vf

G,, GPa (psi x

lo6)

Tensile strength, MPa (psi x lo3)

3930 (570)

5102 (740)

5309 (770)

5447 (790)

1.80 (0.065)

1.74 (0.063)

1.77 (0.064)

1.80 (0.065)

vr

4.4 (0.64) 0.36

4.4 (0.64) 0.36

4.4 (0.64) 0.36

4.4 (0.64) 0.36

G , GPa (psi x lo6)

1.6 (0.24)

1.6 (0.24)

1.6 (0.24)

1.6 (0.24)

1.26 (0.046)

1.26 (0.046)

1.26 (0.046)

1.26 (0.046)

Density, g/cm3 (ib/ i n 3 ) Resin Er, GPa (psi x IO6)

Tensile strength, MPa (psi x lo3) Density, g/cm3 (lb/ in3)

Tables 771

~

T300/5208

T40/1962

T50/1962

P55/1962

P75/1962

77 (11.2)

97 (14.0)

63 (9.2)

115 (16.7)

97 (14.0)

69 (10.0)

94 (13.6)

97 (14.0)

69 (10.0)

231 (33.5)

283 (41.0)

393 (57.0)

379 (55.0)

517 (75.0)

0.27 91 (13.2)

0.32 107 (15.6)

0.24 159 (23.0)

0.33 142 (20.6)

0.27 204 (29.5)

3241 (470)

5654 (820)

2413 (350)

1724 (250)

2069 (300)

1.77 (0.064)

1.80 (0.065)

1.80 (0.065)

1.99 (0.072)

1.99 (0.072)

3.9 (0.56) 0.35

3.7 (0.54) 0.35

3.7 (0.54) 0.35

3.7 (0.54) 0.35

3.7 (0.54) 0.35

1.4 (0.21)

1.4 (0.20)

1.4 (0.20)

1.4 (0.20)

1.4 (0.20)

66 (9.6)

66 (9.6)

66 (9.6)

1.27 (0.046)

1.27 (0.046)

1.27 (0.046)

50 (7.3) 1.27 (0.046)

66 (79.6) 1.27 (0.046)

772 Design allowables substantiation LAMINA PROPERTY REFERENCES

A

Weight density was calculated based on the following relationship:

D

Y~~was calculated based on the following equa-

tion:

pf = fiber density

Vf = fiber volume fraction p, = resin density

AAAS4 and IM6 fiber properties are based on DD Hercules data for IM7/3501-6 from the Graphite Fiber Products Handbook, based on mechanical Hexcel supplied data which was based on tow test data. test data for the fiber modulus and strength. The fiber Poisson's ratio was 'back' calculated E Assumed: based on the composite Poisson's ratio ( Y ~ , ) ,the o3= o2and *3 = +T~ matrix Poisson's ratio and the fiber volume EE IM7 and IM8 fiber properties are based on (see reference H). B L.L. Clements and R.L. Moore, Composite propHercules supplied data (Graphite Fiber Products Handbook) which was based on tow test data erties for E-glass fibres in a room temperature for the fiber modulus and strength. The fiber curable epoxy matrix, Composites, 1978, 9(2), 93-99. Properties for tI3were set equal to valPoisson's ratio was 'back' calculated based on ues for S2-glass/DER 332 since no data was the composite Poisson's ratio (Y~,), the matrix reported. Poisson's ratio and the fiber volume (see referBB Hercules supplied data for IM6/3501-6 from the ence H). Graphite Fiber Products Handbook based on F The following was assumed: mechanical test data. '23 = 31' C Composite is assumed to be transversely isotropic. A transversely isotropic composite is FF Hercules data for IM8/3501-6 from the Graphite a material which exhibits a special case of Fiber Products Handbook, based on mechanical orthotropy, whereby the properties are identitest data., cal in two orthotropic dimensions, but not the G Owens Corning fiber data and short beam third. The properties are the same in both shear data. transverse directions, but not in the longitudi- GG Properties based on test data listed in BASF nal direction. The following equations apply Hexcel Technical Information handbook. for transversely isotropic materials: H Equation for calculating the fiber Poisson's ratio (vf): E, = E, Y12- Y,(1

Vf

GI3 = GI2

1 '3

= '12

= 2(1

+

- VJ

Vf

E3 G23

=

Y2J

CC Properties were set equal to those for AS4 / 3501-6.

HH Amoco data for T300 fibers. Actual test method for fiber modulus and strength was not defined. The fiber Poisson's ratio was 'back calculated based on the composite Poisson's ratio using the equation in reference H. I Fiber and resin are assumed to be isotropic, therefore G, or GI is calculated as follows:

E* Gf = 2(1+

VJ

Lamina property references 773

R

I1

Properties based on test data supplied by Amoco for the T40/1962 system. Amoco supplied data for the 1962 resin system. Resin properties from publication by Texaco, J Huntsman Chemical Co. ’Jeffamine’, resin properties are based on 100 parts of epoxy resin and 45 parts of Jeffamine T-403. The DER 332 epoxy resin was cured with Jeffamine T-403. JJ Properties were set equal to T300/5208 value. KJK Amoco data for T40 and T50 from technical information sheets. Test method for fiber modulus and strength was not specified. The fiber Poisson’s ratio was ‘back‘ calculated based on the composite Poisson’s ratio (vlJ , the resin Poisson’s ratio and the fiber volume (see reference H). L H. Hahn, D. Hwaug, H. Chang, S. Lo, Flywheel Materials Technology: Design Data Manual for Composite Materials, UCRL-15365 Volume 1, P.O. 6641009, Lawrence Livermore Laboratory, July, 1981. LL Amoco test data for T50/1962 M Assumed

S

T

U V

u3= u2and -u3 = +2 MMAmoco test data for P55/1962. N Mil Handbook 17. NN Amoco data for P55 from technical information sheets. Test method for fiber modulus and strength was not specified. The fiber Poisson’s ratio was ‘back’ calculated based on the composite Poisson’s ratio (vlJ , the resin Poisson’s ratio and the fiber volume (see reference H). 0 Properties were set equal to S2-glass/DER332. 00 Amoco data for P75/1962. P Properties for XP251S epoxy were set equal to DER 332 since no data was available. PP Amoco data for P75 which was based on tow test data for fiber modulus and strength. The fiber Poisson’s ratio was ’back’ calculated based on the composite Poisson’s ratio (vI2), the resin Poisson’s ratio and the fiber volume (see equation D). Q DuPont supplied data for Kevlar 29/934, Kevlar 49/934, and Kevlar 149/934. Laminates were fabricated and tested by Boeing

W

X Y

Z

Technology Services, Boeing Commercial Airplane Co. No data was listed in DuPont literature for G12, therefore values for Kevlar 29/934, Kevlar 49/934, and Kevlar 149/934 were set equal to those for Kevlar 49. Hexcel rubber-toughened epoxy system based on a paper by S.R. Swanson, G.R. Toombes, and S.W. Beckwith, In-Plane Shear Properties of Composites Using Torsion Tests of Thin-Wall Tubes, 29th National SAMPE Symposium, April 3-5,1984. DuPont supplied data for Kevlar 29, Kevlar 49, and Kevlar 149 based on tow tests (ASTM D2343).The fiber Poisson’s ratio was ’back‘ calculated based on the composite Poisson‘s ratio (vIz), the resin Poisson’s ratio and the fiber volume (see reference D). Fiberite data for 934 resin system from Fiberite Material Handbook. Values for Kevlar 49/DER 332 set equal to those for Kevlar 49/Epoxy XD7575.03-XD7114Tonox 60-40 from Reference 6 of this list. D.F. Adams, R.S. Zimmerman and H.W. Chang, Properties of Polymer-Matrix Composites Incorporating Allied A-900 Polyethylene Fiber, SAMPE J., September/ October, 1985, pp. 44-48. Note: The modulus of Spectra composites is much lower than expected from the rule-of-mixtures relationship. A possible explanation is that the Spectra fiber modulus is a function of strain rate. For example a single fiber tested at 100%/min strain rate exhibited a modulus of 17 msi versus 11msi for a 8’%/min strain rate. H.W. Chang, L.C. Lin, A. Bhatnagar, Properties and applications of composites made of polyethylene fibers, 31st Intern. SAMPE Symp., April 7-10, 1986. t I 3for Spectra 1000 set equal to the value for Spectra 900. Hercules supplied test data for AS4/3501-6 determined by independent firms; Delsen Labs and McDonnell Aircraft Company. Hercules supplied data for 3501-6 resin. All data is listed in the Hercules Graphite Fiber Products Handbook. R.Y. Kim, E Abrams and M. Knight, Mechanical characterization of a thick composite laminate, Proc. Amer. SOC.Composites, 3rd Technical Conference, 1988, pp. 711-718.

774 Design allowables substantiation Table 33.4 References for lamina properties of glass, aramid and polyethylene fibers in epoxy matrices

Properties

E-glass DER 332

S2-glass DER 332

S2-glass XP251 S

Kevlar 29/934

A

A

A

Q

L L C

N N C 0

Material system Fiber type Resin type Fiber volume Composite density Lamina elastic properties

C C 0 D C

G23 G13 '12

'23 1'

?

Lamina strength properties (psi) ff1

02

Tension Compression

B B

L L

N N

Q Q

Tension Compresson

B B

L L

N N

Q Q

Tension Compression

E E

M M

M M

M M

B B F

L G F

N N F

Q Q

G H I G G

G H I G G

G H I G G

S H I S

J

J

J

J

P K I P K

T K I T T

0 3

212 '13

'23

F

Constituent properties Fiber E, f' Gf

Tensile strength Density Resin Er r'

GI Tensile strength Density

K I K

K

r

K

S

Tables 775

Kevlar 49/934

Kevlar DER332

Kevlar 149/934

Spectra 900/€ PO N 82 6

Q

L

Q

A

Spectra 1OOO/EPON 826

A

W W C

W C C W

D C

Q Q

L U

Q Q

V V

W W

Q Q

L U

Q Q

V

K

W K

M M

M M

M M

M M

M M

Q Q

Q Q

F

L L F

V V F

W X F

S H I

S

S

H I

H

V H

W H I W W

F

S S

S

I S

I V

S

S

V

T K I T T

J

T K I T

K K I K

T

K

K I

J K

K K I K K

776 Design allowables substantiation Table 33.5 References for lamina properties of graphite fibers in epoxy matrices

Properties

ASA /35016

lM6/ 3501-6

lM7/35016

lM8/ 3501-6

A

A

A

A

BB BB C BB

DD

FF

C DD

C

cc cc cc cc cc

C FF

Material system Fiber type Resin type Fiber volume Composite density Lamina elastic properties El E2

E3

G12

cc cc cc cc

G23

G,, v12 v2.3 31'

cc

cc

cc cc cc cc cc

Lamina strength properties (psi) 0,

Tension Compression

Y Y

BB BB

DD DD

FF FF

Tension Compression

Y Z

BB

cc cc

cc cc

Tension Compression

Z Z

E E FF FF

02

03

21' 31' 32'

cc cc cc

Y Y Z

BB BB

cc

E E DD DD

AA H I AA AA

AA H I AA AA

EE H I EE EE

EE H I EE EE

Y Y I Y Y

Y Y I Y Y

Y Y I Y Y

Y Y I Y Y

cc

cc

Constituent properties

Fiber Ef Yf

G*

Tensile strength Density Resin Er vr

Gr

Tensile strength Density

Tables 777

T300 5208

T40/ 1962

T50/ 1962

A

A

A

A

A

GG GG

I1 I1

LL LL

MM MM

00 00

C

C

C

C

L

cc cc

I1

LL

MM

C C

C C

C

L

cc

I1 D

LL D

C h4M D

C 00 C C

C

C

C

C

C

GG GG

I1 I1

LL LL

Mh4 MM

00 00

L L

I1

LL

MM

00

JJ

JJ

JJ

JJ

E E LL LL F

E E MM Mh4 F

E E

E E GG GG

cc

P55/ 1962

P75/1962

00

D

00 00

F

HH H I HH HH

KK H I KK KK

KK H

NN H

PP H

I KK KK

I NN

NN

I PP PP

GG K I GG GG

I1 K I I1 I1

11 K I I1 I1

I1 K I I1 11

I1 K I1 I1

II

MECHANICAL TESTS

34

Yu.M. Tarnopol’skii and VI L. Kulakov

34.1 STRUCTURAL HIERARCHY OF FIBROUS COMPOSITES

Fibrous composites are inhomogeneous materials with multiple levels of structural scale. The three levels of structural scale can be arranged in a hierarchy. The characteristic dimensions for the three levels are: fiber diameter, lamina thickness and plate thickness. The most appropriate test methods and structural analysis techniques are different for each level in the hierarchy. Test objectives and associated problems are also different for each level. The smallest scale is the diameter of the reinforcing fiber. The properties of the reinforcing fiber and polymer matrix and their interaction are studied in the field of micromechanics. The second level scale is the thickness of the unidirectional lamina. Macromechanics describes the properties of a monolayer under loading at an angle to the fiber direction. A monolayer is defined as a flat or curved element of material composed of a polymeric matrix and reinforcement of the same type and orientation throughout the layer. It is the basic structural element of laminated and fibrous composites. The characterization of monolayers by mechanical test methods is given particular emphasis in this chapter since testing of anisotropic materials is a relatively novel and seldom studied field of mechanics.


The test results are used to calculate the properties of materials with more complex configurations of fiber arrangement and of hybrids, i.e. materials with different reinforcements in the same lay-up. For multilayered composites, the largest structural scale is the thickness of the laminated plate that is equal to the sum of stacked laminae and interleaves. The theory of laminated plates allows one to determine the properties of the plate using the properties of the monolayers and their stacking sequence. If the monolayers are part of a structural element, then the highest level of scale is the characteristic size of the object. The dimensions of structural elements typically exceed the thickness of the plate by several times. The properties of components are measured by traditional mechanical and physical test methods’. The results of the analyses or tests at the first level are used as input for the analyses at the second level. The same relationship holds for the second and third levels. Upon converting each scale to a continuum at the next higher scale, it is important that for each level under consideration the number of elements be sufficient, e.g. through the width and thickness of the lamina or through the thickness of the plate, so the transition from a discrete medium to a continuum is performed without great error. This progression up the hierarchy eventually leads to solutions of real life problems involving structural elements or prototypes.

Problems of composite testing 779 34.2 PROBLEMS OF COMPOSITE TESTING

systems of coordinates are introduced: the axes of elastic symmetry in the material (1, 2, 34.2.1 HISTORY 3) and axes of loading (x, y, z for flat specimens; 8, z and Y for ring and tubular The first reference to mechanical testing of a specimens). It is preferable to use methods in structural material for an engineering applicawhich the x, y, z axes (or 8, z, Y ) coincide with tion is dated July 4,1662. The objective of the 1 , 2 , 3 axes. the tests was to compare the tensile strengths of The majority of laminated and fibrous comcords made of Riga and Dutch yarns. The posites exhibit low interlaminar shear and stronger in this contest was the thinner cord transverse tension strengths. Shear strength is manufactured in Riga. In the years that folcharacterized by the relations between E x / G x Z lowed, equipment and methods for testing (shear stiffness) and a;/zx; (shear strength). engineering materials, particularly metals, Transverse tension and compression strengths have reached a high degree of perfection and perpendicular to the fibers are determined by consistency. The appearance of composites E x / E Z , ax’u/az’u, axt”/a;, where Ex the relations and their ever-expanding use has once again and EZ are the moduli of elasticity in the x and made it necessary to improve mechanical test z directions; GxZ is interlaminar shear modulus; methods. Although significant progress has a,“ and a; are strengths in the x and z direcbeen made, there are vast differences in their tions; zxzuis shear strength in xz plane. The x maturity. The methods differ primarily in the and y axes are located in the fiber lay-up (reindegree to which they minimize extraneous forcement) plane, the z-axis is perpendicular stresses and strains. Although test methods to this plane; the (t) and (c) designate tension tend to become more complex as their accuand compression, respectively. racy increases, economics must be considered in their selection. Factors such as complexity of specimen preparation, amount of material 34.2.3 UNIQUE REQUIREMENTS OF required, and the requirements for specially COMPOSITE TESTING designed equipment must be considered. The anisotropy and unique structural properties of composite materials cause serious difficulties. For example, a large number of 34.2.2 DETERMINABLE VALUES strength and elastic properties must be deterThe purpose of mechanical tests is to determined for complete characterization of the mine the strength and elastic properties of a material. Since the number of determinable material. However, only loads, displacements characteristics depends on the state of stress and strains can be measured in a mechanical and the degree of anisotropy23, one should test. The theory of elasticity for an anisotropic select the loading methods for which the body is used to determine the desired proper- experimentally determinable values are most ties of composites from these measurable simply related to the material characteristics. quantities. It should be remembered that The selection of techniques for analyzing the advanced fibrous composites with unidirec- data is critical as well as the determination of tional, laminated or spatial fiber lay-ups are their range of validity. Since the composite inhomogeneous, essentially anisotropic mate- analysis techniques are based on the theory of rials. The customary terms, i.e. tension, elasticity for an anisotropic body, it is necescompression, shear and bending, are meaning- sary to consider the error in treating an less without specification of the direction of inhomogeneous anisotropic medium as a conthe load and its relationship to the axes of elas- tinuous anisotropic medium. For example, the tic symmetry of the material. Therefore, two number of structural elements (fibers, lamina,

780 Mechanical tests etc.) must be sufficiently large to support this approximation43r6. Once the general test method has been selected, the details of the loading and the sample geometry must be selected. For fibrous composites, the principal difficulties lie in the generation of a uniform stress field in a representative volume of material, i.e. the elimination of end and edge effects. This is difficult even for the most simple types of tests. The difficulties increase with increasing degree of anisotropy, i.e. materials reinforced by high-modulus or high-strength fibers (boron, carbon and organic fibers). End effects are primarily influenced by the method of fastening and loading of the specimen, the length of the grip section, and the fiber orientation. The region involved in end effects extends in the direction of the greatest stiffness of the material and increases with the anisotropy of the material. Edge effects are primarily influenced by the size and shape of the specimen, the fiber orientation and the angle of specimen cutting. If strength anisotropy is present, improper loading and fastening can lead to changes in the failure mode and the resulting strength value. A most important considerationis the selection of the specimen width. The width must be large enough to avoid the effect of cut fibers at free edges which is important for specimens of off -angle, angle-ply and cross-ply materials. Edge effects are manifested as interlaminar stresses at the free edges of the specimen, the direction and magnitudes of which depend on the fiber lay-up. Material quality also causes unique requirements for testing composites. Quality cannot be ignored during testing because the material and structure are formed simultaneously. In addition, composites are extremely sensitive to mechanical and thermal history. Structural imperfections, in particular porosity, waviness and misalignment of fibers, require special attention. The presence of porosity affects the measurement of polymer matrix dominated properties, e.g. shear strength. Small amounts

of fiber waviness can cause the measured values of longitudinal modulus of elasticity and strength to be considerably lower than those of materials with ideally straight fibers. Fiber waviness also influences on the coefficient of thermal expansion in the fiber direction. The modulus of elasticity perpendicular to the fiber direction and the in-plane shear modulus are not significantly affected by fiber waviness. All of the aforementioned unique testing requirements apply to composites of a fibrous and laminated structure. Additional difficulties arise when spatially reinforced composites are tested because the transverse strength and stiffness are derived from a rigid framework rather than from a compliant matrix. 34.2.4 SUMMARY TABLES

The most common methods of testing composites in tension, compression, torsion and bending are described in Tables 34.1-34.5 . The high performance test fixtures designed specifically for composite testing, their description and recommended applications are given in Reference 7. 34.3 TEST SPECIMENS

The important relationships between fabrication methods, test methods and required specimen shapes are shown in Fig. 34.1. Specimens for mechanical testing are classified as flat specimens (bars and plates), rings (complete and segments) and tube@. The specimens and test methods in Fig. 34.1 are used to characterize the monolayer. Flat monolayers can be characterized with specimens that have a different fiber lay-ups but the same general, flat, long, narrow shape. To adequately characterize wound monolayers, it is necessary to use both rings and tubes. Ring specimens of a unidirectional fiber lay-up are used to assess characteristics in the fiber direction. Tubular specimens with a 90" wind angle are used to measure properties perpendicular

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methods, loading types, and failure modes are all different for the two test methods (Table 34.1). A uniaxial tension or compression test specimen has several functionally different parts: two loading sections, a gage section, and two transition sections. The loading sections provide a means of fastening the specimen in the testing machine. They receive and transmit the external loads to the gage section of the specimen. In the gage section, deformations are measured and stresses are calculated according to the geometrical dimensions and external load. The transition 34.4 TENSION AND COMPRESSION sections serve to attenuate stress-strain perturbations in the loading section to isolate 34.4.1 TENSION TESTING OF FLAT SPECIMENS them from the gage section. The specimen In spite of any analogy that may be drawn dimensions that are recommended in the between loading of flat composite specimens existing standards meet these requirements. in tension and compression, only the calcula- The specimen dimensions (length, width and tion relationships (taking into account the sign thickness) specified in standards as a function of the stresses and strains) are common to both of the type of fiber lay-up are shown in Table cases. The specimen shape and size, gripping 34.1 (Methods (a) and (b)). to the fiber direction. Tubular specimens with different balanced fiber lay-ups (fiber angles are symmetric with respect to the longitudinal specimen axis) are employed to assess shear characteristics and to study complex states of stress. The specimen shape, to a great extent, depends on the objective of the test: verification of scientific hypotheses, engineering specification of the material, or quality control of the materials. The most rigorous requirements are imposed on specimens of the second group.

Tension and compression 787 The greatest technical challenge in tension testing of composites, especially unidirectional composites, is the reliable transmission of tensile forces from the grips to the specimen. This is generally performed through the use of friction forces. Tabs bonded to the specimen improve the efficiency of load transmission considerably. The tabs should be made of a material that has a much lower modulus of elasticity and a higher total elongation than the respective characteristics of the specimen material. Tabs have been made of fiberglass reinforced composites, aluminum and wood veneers. The thickness of tabs should be between 1.5t and 4t, where t is the specimen thickness. The tabs must have a large enough area that the ultimate shear load capacity of the bond between the tabs and the specimen is greater than the breaking load of the specimen gage section. The mode of failure in tension depends on the relationship between the external load and the reinforcing fibers and on the type of reinforcement lay-up. When unidirectional composites are loaded in the reinforcement direction, they fail by breakage of the reinforcing fibers. This is accompanied by transverse cracks, longitudinal shear cracks and delamination of the polymer matrix. Increasing the angle between the load and the reinforcing fibers causes the mode of failure to change gradually from shear and splitting of the polymer matrix parallel to the fiber direction to pure transverse tensile cracking of the polymer matrix. The failure mode of composites with a balanced angle-ply reinforcement depends on the angle of the fiber lay-up. 34.4.2 COMPRESSION TESTING OF FLAT SPECIMENS

The main problem in compression testing of flat specimens is the selection of a loading method that ensures compressive failure. When the loading is achieved by normal

forces acting on the specimen ends (Table 34.1, Method (c)),it is impossible to achieve a sufficiently uniform stress distribution over the faces of the specimen. As a result, premature local failure of the specimen occurs. When the loading is achieved by shear forces acting on the sides of the specimen as specified by ASTM D3410 (Method (d)), the stress distribution in the specimen is also non-uniform, especially when flat wedge grips are used. The best method is a combination of the two methods in which normal forces are applied at the ends and shear forces are applied along the faces of the specimen grip section (Method (e)). In compression testing of unidirectional composites in the fiber direction, three basic modes of failure are observed: buckling of the reinforcing fibers, transverse cracking of the matrix, and shearing of reinforcing fibers at a 45" angle without local buckling of the reinforcement. Transverse cracking is caused by differences in the Poisson's ratios of the material components and by a non-uniform transverse strain distribution along the specimen length. Materials reinforced at an angle to the specimen's longitudinal axis fail in shear without crumpling at the end faces because all of the shear load is taken up by the matrix. The aforementioned basic modes of failure can be accompanied by a series of other phenomena: inelastic and non-linear deformation of the reinforcing fibers and matrix, delamination, surface peeling, overall buckling and crushing of the end faces. Failures with different combinations of these phenomena can make the determination of the failure mode very difficult. In compression testing, great care must be taken to ensure stability of the specimen, especially in the gage section. Buckling of the specimen side face is not always detectable and will cause erroneous strain measurements. Special test fixtures are used to prevent overall buckling of the specimen.

788 Mechanical tests 34.4.3 TENSION TESTING OF RINGS

The stress concentration problem is eliminated in tests that use uniform internal The most popular means of tension testing pressure generated by the use of a compliant rings uses a half-disk loading device (Table (d)) or by a hydraulic system ring (Method 34.2, Methods (a) and (b)). This is because the (Method (e)). The disadvantages of the comtest is easy to perform and the data is easy to pliant ring test method are the need for analyze. However, it has several significant periodic calibration of the loading element disadvantages: the strain distribution over the and the need for very careful preparation of specimen circumference is non-uniform, fricthe specimen surface. The disadvantage of the tion between the specimen and the half-disks hydraulic test technique is the need for expenhas a strong effect on the results, and there is a stress concentration in the specimen at the gap sive hydraulic equipment. between the two half-disks because of a variation in the radius of curvature of the specimen. In thin-walled rings, the stress concentration 34.4.4 COMPRESSION TESTING OF RINGS takes the form of a slight increase in radial ten- Radial compression of rings is accomplished by sile stresses that causes a minor effect on the external pressure (Table 34.2).The analog of the test results. However, in thicker specimens, half-disk tension test is the simplest compresthe resultant interlaminar shear stresses can be sion test method (Method (c)). The primary high enough to cause failure at lower loads difference is that in the compression test, it is than failure due to the circumferential stresses possible to reduce the stress concentration in alone. The error increases with increases in the the specimen at the split line. The best results relative specimen thickness t / R , the degree of are obtained with a semi-circular housing that anisotropy, and the ultimate strain of the mate- has a locking arrangement that prevents radial rial. Since this test method yields erroneously growth at the split-line (Fig. 34.2). There are low strength values, it can only be used for also compression analogs for the compliant qualitative comparison of composites. ring (Method (f)) and the hydraulic system Corrections have been suggested but have not (Method (g)).In the compliant ring method, the been used in practice. compliant ring is the elastic foundation of the

Fig. 34.2 Typical interlocking features for ring compression fixture.

Shear 789 specimen and to a certain extent it prevents buckling of the specimen. The external pressure may also be applied by mechanical devices such as multiple cam. The primary difficulty in compression testing of rings by external pressure is the selection of a relative thickness, t / R , to reduce secondary loading effects. Depending on the relative thickness of specimen, t / R , and the degree of anisotropy of the material E,/G,, three different failure modes have been observed. Thin-walled rings fail by buckling, thick-walled rings fail by biaxial compression, and optimum thickness rings fail by circumferential compression. Analysis of the test data must consider the radial as well as circumferential stresses. Delamination of the inner layers of the ring makes it difficult to correctly calculate the compressive strength. This delamination is often noisy. The occurrence of this failure mechanism also depends on the relative thickness of the ring. Delamination of helical windings can lead to unwinding of the specimen. Hoop wound rings can fail by layer-by-layer delamination. 34.5 SHEAR

34.5.1 IN-PLANE SHEAR

Shear properties, especially shear strengths, are difficult to measure. The simple and economical rail shear test is often used for this purpose (Table 34.3, Methods (a) and (b)).The extent of edge effects and the uniformity of the shear stress distribution over the specimen width depends on the length-to-width ratio of the specimen gage section L/w and on the relation of elastic constants G,/E of the material. Edge effects are negligible tor L/w > 10. Edge effects cannot be eliminated for materials with v, = vyxG -1. Thus, this method cannot be used for such materials. The elastic constants obtained by the rail shear test are less sensitive to the relative ratio, L/w, since the measurements are taken in the center of the specimen

gage section where the state of stress is the most uniform. However, edge effects have considerable influence on the shear strength. Therefore, it is better to bond the specimen to the rail links than to use mechanical fasteners. The stress distribution is not affected by the loading direction, i.e. along the diagonal or parallel to the sides of specimen gage section. Measuring the shear properties by tensile loading of an anisotropic strip is distinguished by its apparent simplicity (Table 34.1, Method (f)). The strip can have one of several different fiber lay-ups. This method is not used to determine in-plane shear strength because it yields low values. A state of pure shear is not assured even with a +45" lay-up. A similar test method involves tensile loading of a strip of a unidirectional material cut at an angle, 8, to the reinforcing fibers. The optimum angle is the one for which the relative shear strain y 1 2 / ~is x maximized and the shear stress rI2reaches its critical value. This angle depends on the anisotropy of the elastic and strength properties of the material tested. For advanced composites, the optimum angle is 10 to 15". Because the stress ratios are very sensitive to changes in the angle, rigid tolerances, 4 degree, are set on the specimen cut angle, the strain gage angle, and the direction of loading. In order to ensure that the stress state is uniform, relatively narrow strips, L/w = 14 to 16, are used. The in-plane shear modulus is often measured by twisting a square plate with four point loading (Method (c)). The wide acceptance of this method may be attributed to the simplicity of its calculations. However, the experiments should be performed with utmost care. This method is only applicable for small deflections, wp < O.lt, on plates made of materials which are uniform in thickness and orthotropic along the specimen axes. Test results for several different materials have shown that the P-wp relationship remains linear up to wp/t z 1. However, in practice, the deflection, w should be limited to 0.5t to preP' vent instability. Only the initial linear section

790 Mechanical tests of the P-wp curve should be used to determine the shear modulus. The optimum range of relative plate thickness, L / t , is determined by two conditions: the contribution of transverse shear to the deflection at small values of L / t and the possible loss of stability at large values of L / t . The limits of L / t are given for BFRP in Table 34.3. However, tests run on GFRP (glass fiber reinforced plastic), CFRP (carbon fiber reinforced plastic) and BFRP (boron fiber reinforced plastic) with different fiber lay-ups have shown that reliable data can be obtained at L / t > 15. The specimen must be flat and of constant thickness because the calculated shear modulus is related to t3. The distance from the point of support or load application to the corners of the plate should not exceed 2t. Experimental evaluation of these three shear test methods has shown that they all yield comparable values of in-plane shear modulus9. The successful application of the double Vnotch or Iosipescu shear test (Method (d)) to all types of fiber lay-ups is well known'O. It is invaluable for testing spatially reinforced composites since these materials do not possess planes of low shear strength. All other shear strength test methods use this characteristic of laminated composites to induce shear failure first and therefore are useless for spatially reinforced materials. In-depth investigations have shown that the Iosipescu method and its modifications yield good results in shear tests of carbon-carbon composites reinforced along three mutually perpendicular directions (3-D) and four principal diagonals of cube of 4-D". Spatially reinforced composites are less sensitive to the dimensions of the notches and gage length than laminated composites. The distribution of shear stresses is essentially uniform throughout the gage section of 3-D and 4-D materials cut at a 90" angle with a total notch depth equal to a half of the specimen thickness. Moreover, it is possible to prevent stress concentrations at the notch tips. Specimens with extra side notches yield the best shear strength data.

34.6 TORSION

Torsional loading of thin-walled tubes is a standard test for measuring in-plane shear modulus and strength (Table 34.4, Method (c)). In this test, the stresses are distributed uniformly around the circumference and along the length of the specimen. The shear strains are practically constant through the thickness of the specimen wall. In torsion, the definition of 'a thin-walled tube' is a function of the degree of material anisotropy EJE,, which can vary over a wide range. The disadvantages of this method are that it requires relatively large specimens, special test fixtures, inserts to prevent buckling of some specimens and wound specimens or specimens of special configurations, e.g. materials in which the fiber lay-up is parallel to the specimen axis. The results obtained by torsional shear tests compare favorably with results obtained by test methods using flat specimens. Torsional loading of split rings is also used to measure shear moduli (Method (b)). If the sample size limitations indicated in Table 34.4 are followed, bending effects are negligible. 34.6.1 INTERLAMINAR SHEAR

Good estimates of interlaminar shear properties, especially for spatially reinforced materials, have been obtained by torsion testing of rods with a circumferential notch (Method (a)).The specimens can be tested with or without a central bore. The important geometric parameters of the notch are the relative width, Lp/d, diameter, d, and wall thickness, t. It has been shown that within a range of L / d = 0.2-1.0, the length of the notch does not afPect the measured shear strength, tnu.The gage section diameter can be increased from 5 mm to 15 mm (0.6 in) without affecting rnu. However, increasing the diameter beyond 15 mm (0.6 in) causes a sharp drop in the measured strength.

Bending 791 34.7 BENDING

for measuring interlaminar shear strength. However, refined analysis has shown that the 34.7.1 THREE-POINT BENDING state of stress in a short bar of anisotropic The most popular type of bend test is the material is significantly different from the three-point bend test (Table 34.5, Methods (a) state of stress predicted by isotropic theory2J2. and (b)).The four-point and five-point bend The shear stresses through the thickness of a tests are less popular in spite of their consider- relatively short anisotropic bar have a paraable technical advantages over the three-point bolic distribution only in the middle of the span. At the loading points, the distribution of bend test. Theoretically, the moduli of elasticity of shear stresses through the thickness of the homogeneous materials in tension, compres- specimen has peaks near the surface directly sion, and bending are the same, i.e. Exf = E; = beneath the loading points. In relatively short E:. However, due to imperfections, the state of anisotropic bars, there are no planar regions of stress in bending and differences in the fiber constant maximum shear stress. Moreover, on lay-up through the thickness of the material, relatively short bars (L/w 4 5), the compresthe bending modulus of elasticity E: can differ sive transverse stresses from the load somewhat from Exf or E;. This difference is application points can extend over the entire length of the specimen and can exceed mean emphasized by the superscript ’b’. The formulae used to determine the elastic shear stress by up to a factor of 15. These combending constants, E: and GxZb, from the deflec- pressive stresses constrain crack opening at tion of a bar at its midspan must take into sample delamination and result in an apparent consideration the effect of interlaminar shear. increase of interlaminar shear strength. As a The effect of interlaminar shear can be result of these deviations from the ideal paraneglected for large values of relative span bolic stress distribution, the experimentally length L / t . For highly anisotropic materials, the determined interlaminar strength appears to decrease with increasing relative span width. relative span length must be greater than 40. Therefore, shear testing of relatively short bars When determining bending strengths, failcan provide only a qualitative comparison of and ””, ures limited by the normal strength, ., different composites. failures limited by the shear strength, zxF, Interlaminar shear strength can also be must be separated. Unlike those of isotropic materials, the two strengths for composite measured by three-point bending of curved materials can differ by an order of magnitude. segments. The shear strength is calculated The shear stresses can have a considerable with the same equations used for bending of prismatic bars. However, the additional intereffect on the failure. laminar normal stresses unique to curved Failure due to normal stresses occurs by beams must be taken into consideration. The fracture of the extreme outer layers in comnormal stresses act over the entire length of pression or tension. Failure due to shear the specimen. The sign of the stress depends stresses occurs by delamination approximately on the orientation of the specimen. In the case at the midplane of the specimen. Laminated of segments loaded with their convexity materials can fail by a violent debonding of the upwards (center load applied to the outer compressed outer layer. Very short bars experidiameter of the specimen) the stresses are tenence a third failure mode. They fail by 0 : . When the convexity is downwards, the sile, crumpling and shearing which is accompanied radial stresses are compressive, a;. In the forby an apparent increase in shear strength. Three-point bend testing of short bars or mer case, shear and tensile radial stresses ring segments is the most widely used method combine to decrease the apparent shear strength. In the latter case, the compressive


and 11),and the Arcan Test (mixed Modes I and 11). The geometry of the specimens, preparation methods, and analysis procedures have been described in detai113J4. Advanced composites, especially those reinforced with carbon and aramid fibers, have highly anisotropic thermophysical properties. This is reflected in the thermomechanical behavior of the structures fabricated with these materials. The combination of a polymer 34.7.2 BENDING OF RINGS matrix having a high coefficient of thermal Bending of complete rings by diametrically expansion and fibers having a negative coeffiopposed loads (Method (c)) is used to deter- cient of thermal expansion allows the mine elastic and strength properties of fabrication of composites with extremely low composites. Reliable results are obtained if the thermal expansion. This property of composrelative specimen thickness, t l r , is properly ites is used in the fabrication of structures selected. The acceptable range of relative which are stable over wide temperature thickness for determining the shear modulus, ranges. Low thermal expansion is commonly Go,: is based on the material anisotropy, achieved in two directions and the process can E,b/G,b. The shear modulus is calculated from be extended to materials which are spatially the load-diametral deflection data using the reinforced in three directions (3-D) or along same equations for three point bending of four diagonals of a cube (4-D)15.These comprismatic bars with a correction factor for the posites possess a thermal expansion coefficient fraction of deflection induced by shear which is both isotropic and very low. Measurement of the thermal expansion stresses. When the test is used to determine the interlaminar shear strength, rOpU,the rela- coefficients of carbon and aramid composites, tive specimen thickness must be chosen to especially in the reinforcement direction, is not ensure failure by shear delamination at the a trivial experiment. It must employ modem specimen mid-radius rather than failure by dilatometer and interferometer methodsI4. normal circumferential stresses, a?.

radial stresses impede the growth of the shear delamination crack and raise the apparent interlaminar shear strength. For accurate determination of the interlaminar shear strength, r&bu,segment dimensions must be selected so that the normal circumferential stresses, as, and normal radial stresses, or,are negligible compared to the shear stresses.

34.9 STRUCTURAL TESTING

34.8 SPECIAL TESTS

Laminates can fail by interlaminar delamination along specific planes. Toughness data is as important for characterization and failure prediction of composites as strength and stiffness data. Cracks in composites can propagate by Mode I, (crack faces opening normal to the crack plane), by Mode I1 (crack faces sliding in their planes), or by Mixed-Mode (combination of Mode I and Mode 11).The interlaminar fracture tests include the Double Cantilever Beam Test (Mode I), the Edge Delamination Test (Mode I), the End Notched Flexure (Mode 11), the Notched Three-Rail Shear Test (Mode 11), the Cracked Lap Shear Test (mixed Modes I

The unique challenges of designing and testing composite parts are due to the fact that the material and its micro- and macro-structures are created at the same time as the part. The design of critical structures must include the design of the material and must consider the unique behavior of composite materials which is influenced by processing techniques, actual service loading, and environmental conditions. Structural testing should start with tests of small-scale models fabricated by the same manufacturing process as the full-scale structure, followed by tests of prototype parts, specimens cut from structural elements, and finally, full-scale tests9J9.

References REFERENCES

1. Lubin, G. (ed.). Handbook of Composites, New York: Van Nostrand Reinhold, 1982. 2. Tamopol‘skii, Yu.M. and Kincis, T.Ya, Static Test Methods for Composites. New York: Van Nostrand Reinhold, 1985. 3. Witney, J.M., Daniel, I.M. and Pipes, R.B.,

793

11. Greszczuk L.B., Shear Modulus Determination of Isotropic and Composite Materials. ASTM

Special Technical Publication, 1969,460: 140-9. 12. Uemura M., Problems in Mechanical Testing Methods of Advanced Composite Materials. In Proc. 10th Tsukuba General Symp., 1990, pp. 43-54. 13. Pagano, N.J. (ed.), Interlaminar Response of Experimental Mechanics of Fiber Reinforced Composite Materials, (Composites Materials Composite Materials, Rev. Ed., Society for Series; 5) Amsterdam: Elsevier, 1989. Experimental Mechanics. Englewood Cliffs, NJ: 14. Carlsson, L.A. and Pipes, R.B., Experimental Prentice-Hall, 1984. Characterization of Advanced Composite Materials. 4. Kelly. A. (ed.). Concise Encyclopedia of Composite Englewood Cliffs, NJ: Prentice-Hall, 1987. Materials, Oxford: Pergamon Press, 1989. 15. Tarnopol’skii, Yu.M., Zhigun, I.G., and 5. Tarnopol’skii, Yu.M. and Vasiliev. V.V. (eds.). Polyakov, V.A., Spatially Reinforced Composites. Structural Composites. A Handbook, Moscow: Lancaster: Technomic, 1992. Mashinostroyenie, 1990. 16. Mechanical Testing of Advanced Fibre Composites, 6. Tsai, S.W., Theory of Composites Design, Dayton, Imperial College of Science and Technology, Paris and Tokyo: Think Composites, 1992. London: University of London, 1992. 7. W y o m i n g Test Fixtures. High Performance Test 17. Sims G.D., Nimmo, W., Johnson, A.F. and Fixtures. Product Catalog, Laramie: Wyoming Ferriss, D.H., Analysis of Plate-Twist Test for InTest Fixtures Inc, 1993. Plane Shear Modulus of Composite Materials. 8. Peters, S.T., Humphrey, W.D. and Foral, R.F., Teddington: National Physical Laboratory, 1992. Filament Winding Composite Structure Fabrication. 18. Lee, S. and Munro, M., Evaluation of in-plane Covina: SAMPE, 1991. shear test methods for advanced composite 9. Nikolaev, V.P., Panfilov, N.A., Popov, V.D., and materials by the decision analysis technique. Sinitsyn, E.N., Analysis of the Failure Composites, 1986,17(1), 13-22. Mechanism of Large-Scale Structures. Mech. 19. Nikolayev, V.P., Popov, V.D. and Sborovskii, Composite Mater., 1993,29(2):203-11. A.K., Strength and Reliability of Wound 10. Pinderra, M.J., Gurdal, Z.C., Hidde, J.S. and Fiberglass Reinforced Plastics. Leningrad: Herakovich, C.T., Mechanical and Thermal Mashinostroyeni ye, 1983. Characterization of Unidirectional Aramid/ Epoxy. Report CCMS-86-08, VPI-E-86-29, Virginia Polytechnic Institute and State University, Blacksburg, VA, 1987.

DURABILITY AND DAMAGE TOLERANCE OF FIBROUS COMPOSITE SYSTEMS

35

Ken Reifsnider

35.1 DEFINITIONS AND ISSUES

Durability and damage tolerance are critical to the design of composite structures. Damage tolerance is the approach often required for the certification of safety-rated structures such as aircraft components; durability has been identified as one of the most important technical drivers for the design of major composite structures such as the High Speed Civil Transport. Recent reports from the National Materials Advisory Board and a great volume of other literature focus on these Of course, there are many nuances in the definitions of durability and damage tolerance. However, the basic concepts are quite

simple, and are illustrated in Fig. 35.1. Damage tolerance is the remaining strength after some period of service, and durability, in general, has to do with how long the component will last, i.e. with the life of the structure. In this context, durability is often discussed in terms of the resistance or susceptibility to damage initiation. Both concepts imply that the subject component is being exposed to applied conditions such as mechanical loading and environments such as temperature and chemical agents over long periods of time that are typical of the projected service life of the component. There are several technical concepts that form a foundation for our discussion of these closely related topics. The first of these is the

Damage Tolerance (Remaining strength) 1

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Time / Cycles Handbook of Composites. Edited by S.T.Peters. Published in 1998 by Chapman & Hall,London. ISBN 0 412 54.020 7

Fig. 35.1 Basic definitions of ’durability’ and ‘damage tolerance’.

Definitions and issues 795 question of the relationship of material strength to structural strength. In general, the strength of (fiber reinforced) composite materials is represented by an array of values that reflect the anisotropic nature of the materials (Fig. 35.2). For planar materials, at least the tensile strength and compressive strength in the fiber direction and transverse to the fibers and the in-plane shear strength are required for a complete answer to the question of ’how strong is this material’. However, as an array, those values do not directly show ’how strong is a composite structure’. Several possible answers to that question are typically given. One may use a ’failure criterion’ that compares all of the point stress components with all of the material strength components (such as the Tsai-Hill or Tsai-Wu riter ria)^ in some collective form based on concepts such as critical energy, critical shear resistance, etc. The salient point to be made is that the complexity of (inhomogeneous) composite materials and their array of anisotropic material strengths give rise to the development of a corresponding array of damage and failure modes in these materials that must be understood and correctly modeled to answer the question of

’how strong is this composite structure’, even if the array of material strengths are known (shown in Fig. 35.2). Hence, there is a need to develop understandings and representations of the critical damage and failure modes that control the performance of engineering components. This technology is currently incomplete, but discussions of those topics will follow. A second fundamental concept is microstructural architecture. As shown in Fig. 35.3(a), many fibrous composite components are made in layered or laminated form, with the fibers in different layers having different directions; in some cases the plies are made from different materials to form a ’hybrid’ composite. In addition, the fibers may not be straight, but may be woven, braided, or arranged in mats of various types (Fig. 35.3(b),(c)). These details have a major influence on the durability and damage tolerance of the materials. In fact, most composite material systems are ‘designed’ to be ’fiber dominated’, to take advantage of light, strong and stiff (but brittle) fiber materials that are available. Typically, the fibers, their geometry and their arrangement are important parts of the question.

Five in-plane strength values for fiberous composites: Tension and compression strength in the fiber direction Xt or X,

Tension and compression strength in the direction transverse to the fibers Yt orY,

in-plane shear strength

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Fig. 35.2 Schematic illustration of ’principal strength’ directions in a unidirectional continuous fiber composite laminate.

796 Durability and damage tolerance offibrous composite systems

Fig. 35.3 Typical engineering composite reinforcement types: (a) fibrous, unidirectional pile; (b) fibrous woven; (c) fibrous, braided.

A third technical issue has to do with the degradation of intrinsic strength and stiffness. For metals, the material stiffness and strength are generally constant during the life of the engineering component. This may not be true for composites. Stiffness changes of the order of 10-20% may be caused by micro-cracking, for example. Since many structures are stiffness designs, this mode of degradation must be considered. In addition, the intrinsic material strengths (indicated in Fig. 35.2) may also be degraded, especially by such things as physical or chemical aging. This behavior must also be part of the supporting predictive technology developed for these materials. Nondestructive methods of tracking such degradation are under development, but this remains as a challenge currently. Methodologies for the assessment and prediction of durability and damage tolerance of composite materials typically involve the following features:

0

0

0

0

Remaining strength and life models are developed and predictions are made for each independent failure mode (such as fiber failure in tension or micro-buckling in compression, etc.). Mechanics representations of the state of stress and state of material are constructed on the basis of a 'representative volume' of the material that is typical of the distributed damage state that controls the remaining stiffness and strength of the composite. A typical representative volume of material is a controlling ply in a laminate, but may be a micro-buckling ligament, a small group of fibers, etc. Various methods are used to characterize and monitor the rate of strength degradation in composites. A typical parameter which is useful for that purpose is stiffness change; however, that parameter is not appropriate in some cases. Micromechanics (mechanics analysis at the

Damage modes and failure modes 797 fiber/matrix level of representation) is increasingly used for remaining-strength modeling, for the calculation of stiffness change (which leads to internal stress redistribution), and for the estimation of remaining strength for a given failure mode. Statistical considerations are essential for the correct representation of the long-term behavior of composites. Composites typically fail because of the statistical accumulation of defects, which eventually interact to create a critical condition. This is in contrast to self-similar single crack propagation that is the typical mechanism of failure for common metals. Time-dependent behavior such as viscoelastic creep, creep rupture (driven by such things as internal stress redistribution or oxidation), and aging are typically important in the consideration of the longterm durability and damage tolerance of polymer composites, particularly for components that serve at elevated temperatures. This chapter will discuss the range of physical and engineering details that define and control this subject. Of course, a complete discussion would fill several volumes, so the reader should regard this discussion as only a starting point for further study.

35.2 DAMAGE MODES AND FAILURE MODES

The failure of 'typical' (homogeneous isotropic) engineering materials is a familiar topic. The subjects of ductile rupture and brittle fracture are widely discussed and taught in undergraduate and graduate courses. However, composite materials generally do not behave in a manner easily described by either plasticity (or yield) theory or by selfsimilar crack growth concepts. The reason for this different behavior is the fundamental difference in the micro-structure of composite materials, i.e. in the way they are

made. Composites consist of mechanical 'mixtures' of distinct phases (such as fibers or particles) in a matrix material. The geometry and arrangement of the reinforcement phase is carefully chosen to achieve the desired composite properties. As a result, such material systems are always inhomogeneous, often anisotropic, and often brittle. These three basic characteristics control the nature of damage development and failure in composite materials. The most salient single feature of damage in composites is the process of damage accumulation. Damage development usually involves many damage modes which create a widely distributed damage state, and failure is usually the result of a statistical accumulation of damage (rather than the statistical occurrence of damage). As discussed below, these multiple damage accumulations on failure modes are often closely related to the manner in which the composite is made, especially to the basic nature of the inhomogeneity and anisotropy of the material. This damage development process ultimately controls durability and damage tolerance, so we will discuss some typical major features of that phenomenon. The most pervasive damage mode in composite materials is microcracking, most often in the matrix material. Figure 35.4 shows two embodiments of this mode. Figure 35.4(a) shows an X-ray radiograph of a cross-ply laminate with cracks in both ply types, and Fig. 35.4(b) shows matrix cracking parallel to the fibers in the off-axis plies of a laminate, as seen from a tracing of those cracks as they appear on the edge of this [0,45,45,90Is laminate. A typical scenario for the development of such cracks is the formation of matrix cracks as a function of increasing applied load or increasing cycles of loading. These cracks typically extend through the thickness of a ply and generally extend quickly in the fiber direction if the local stress is uniform. Several other important features of matrix cracking are suggested by Fig. 35.4. As shown by Fig. 35.5, matrix crack formation releases


"'r!

'r!-.eI

rt .1!

.I:

I

Strain Fig. 35.5 Change in slope of the elastic stress-strain curve induced by microcracking. APPLIED STRESS (MPo) 200 300 400 500

100

I

I

b

600 I

I

m

Fig. 35.4 Microcracking in the matrix, parallel to the fibers; a radiograph of a cross-ply laminate with (a) inter-ply delamination at crack intersections (arrow) and (b) a tracing of matrix cracks on the edge of a [0,+45,-45,90] laminate.

stored energy in the cracked ply or material, and changes the stiffness of material proportionately, a matter of concern to engineering applications, as noted earlier. However, the density of cracks in the ply of a laminate reaches a stable saturation value, as first observed by Reifsnider et a1.5,8,called the characteristic damage state (or CDS) of the ply. That CDS is a function only of the properties of the plies, their thickness, and their stacking sequence. Figure 35.6 shows that the same CDS is formed by static or cyclic loading. This CDS can be readily predicted since the crack spacing is determined by the rate at which the surrounding material can transfer stress back into the broken ply. Moreover, the stiffness change caused by this cracking can also be predicted as well7r9-*l. A second important damage mode is delamination, as shown in Fig. 35.7. Delamination is driven by the fact that local regions of the composite would deform differently in response to

01 0

I

I

I

I

0.2

0.4

0.6

0.8

NO.

I .o

OF CYCLES (MILLIONS)

Fig. 35.6 Data showing identity of the equilibrium crack spacing ('characteristic damage state' or CDS) for quasi-static and cyclic loading of a laminate.

the local loads if they were not bonded together in the composite. Hence, stored energy is released if they separate, and that energy drives the separation process. The most common example of this damage mode is the separation of the plies of a laminate near a free edge, as shown in Fig. 35.7. This process has been widely studied and is well described. More will be said of this driving mechanism below. It should be noted that delamination is usually nucleated by other damage modes (such as matrix and although it is a common damage mode, it is not usually a failure mode, per se. Delamination usually begins

Damage modes and failure modes 799 fW"82tr'tiKT W 7 K " V q T 4 ~ K ~ W m matrix ~ Tcracks in one ply may cause fiber fracture in an adjacent ply due to the local stress concentrations21,2z. Figure 35.8@) shows a second feature of importance. When the matrix and fibers have comparable stiffness and strength, the fibers may break many times along their length before the composite fractures. In this situation, fiber fracture can cause a significant I stiffness loss as well as a strength reductionu. Another generic damage mode is micro' buckling, induced by local or global compressive loads, as shown in Fig. 35.9.

-7

1

Fig. 35.7 Edge micrograph of delamination (arrow) showing (a) relationship to matrix cracking; (b) plan view radiograph of edge delamination in a crossply laminate (shaded regions).

at an edge, such as a cutout, bolt hole, rivet hole, etc. If it is in a region of nonuniform stress, it may stop growing when it reaches the boundary of that region. Even if it grows to large dimensions, it usually does not cause significant loss of strength in engineering sized structures. Still, the loss of integrity can lead to other damage and failure modes, so it should be avoided. A third generic damage mode is fiber fracture. Many composites are 'fiber dominated', i.e. they depend on the fibers for their stiffness and strength. Hence, fracture of the fibers is both an important damage mode and failure mode. However, fiber fracture is difficult to detect and has been studied less completely than many other damage modes. However, considerable data have been ~ollected'~'~. Figure 35.8 shows two examples of such data, driven by two important mechanisms. Figure 35.8(a) shows fibers broken beside one another, a typical situation. In many composites, the fibers are coated with a material that decreases the tendency for the fracture of one fiber to cause the fracture of neighboring fibers by forming an 'interphase region' around the fibers that tends to 'isolate' the fracture effectslS2O.It is also important to note that the

800 Durability and damage tolerance offibrous composite systems There are several aspects of this phenomenon that are of importance to durability and damage tolerance. For example, the compression strength (or remaining strength) of the composite may be controlled by the local stress required to initiate the local instability, in which case one wants a large diameter, stiff fibers in a stiff matrix. Or, the strength may be controlled by local resistance to shear deformation after buckling begins, in which case one would choose a tough matrix or interphase region between the fibers and the matrix. This is another case in which a damage mode may or may not be a failure mode, an important distinction.

Fig. 35.9 Localized microbuckling in a polymer matrix composite. Printed with permission, I.M. Daniel and 0. Ishai, Engineering Mechanics of Composite Materials, Oxford University Press, Oxford, 1994.

Another 'damage mode' of considerable importance to polymer composites and all composites used at high temperatures is the phenomenon of creep, i.e. time-dependent deformation at constant applied stress. Figure 35.10 shows a typical form of that behavior, with an initial transient region, a steady state region (in which most engineeringdesign is done), and a tertiary (usually unstable) region. This phenomenon is usually represented by introducing viscoelastic

Time

Fig. 35.10 Schematic of typical creep deformation at constant load and temperature.

or rheological models that represent the behavior in terms of a change in the stiffness of the material with time, as a function of temperature. Quite often, the reinforcing fibers do not show creep behavior at low temperatures, but at high temperatures, essentially all constituents may creep. The changes of stiffness with time can be characterized in the laboratory, and must be modeled carefully, based on those data. In fact, this part of the behavior is critically important to the correct calculation of internal stress states, since the creep of the constituents changes the internal stress distribution greatly in some cases. For example, if the matrix creeps more than the reinforcing fibers (a typical situation), that creep 'relaxes' the stress in the matrix, and increases the load carried by the fibers. If we wish to calculate a fiber-controlled strength, for example, a correct representation of this behavior must be included in our model. Finally, another failure mode is creep-rupture. This is a fairly general terminology used to refer to a variety of physical phenomena that produce time-dependent failure. This can be due to, say, oxidation of the fibers, or to other physical degradation processes which eventually cause rupture. It is clear that these phenomena must also be modeled correctly if we are to discuss durability and damage tolerance of material systems.

Damage drivers and damage 'resistance' 801 35.3 DAMAGE DRIVERS AND DAMAGE

'RESISTANCE'

In the previous section, a number of damage and failure modes that occur in composite materials, and ultimately control durability and damage tolerance, were identified. Many of these modes are related to the manner in which the composites were put together. This raises the basic question of 'can one design composite materials to be durable and damage tolerant?' Most of the rest of this discussion will address this question. Some general concepts will be followed by some micro-mechanics methods of quantifymg answers. Microcracking is likely to be the most pervasive damage mode in typical composites, especially under long-term loading, and most especially under cyclic loading. Even though most matrix materials are chosen because they offer some level of ductility, in most composite systems the matrix is highly constrained so that cracks develop due to local constraint, local stress concentrations, and local defects that grow rapidly under what is generally a 'plane strain' condition. Hence, matrix toughness, in the general sense, is the key to the reduction of matrix cracking. Increasing the strain to failure of the matrix material is a primary objective, and increasing the plane-strain fracture toughness of the matrix is a companion objective. There is a richly developed science and technology associated with matrix toughening; some starting points are listed in Wilkinson et al.24and Hedrick et al?5. A second damage mode identified earlier is delamination. This problem is a strong combination of structural and material concerns. The material concerns are essentially the same as those discussed for matrix cracking, with one important exception. Matrix toughness does not translate directly into interlaminar toughness. Hence, resistance to delamination cannot be controlled entirely with material property increases. The structural part of the problem does, however, present opportunities. It was mentioned before that delamination is driven

by local discontinuitiesin stress state, typically caused by neighboring plies or ply groups (bonded together) that would have very different strain states if they were not bonded. Hence, the orientation of the plies in a laminate and the stacking sequence of those plies are controlling players in the development of the interlaminar stresses that drive delamination. This problem has been exhaustively studied, and methods of reducing interlaminar stresses have been widely d i s c ~ s s e d ' ~ - ~ ~ , but because of the inhomogeneous and often anisotropic nature of composites, interlaminar stresses generally cannot be eliminated in laminated systems, so mechkical methods are widely used to control that tendency. The most successful of these is weaving, i.e. to use woven fiber architectures to reduce the anisotropy of a given ply, and therefore, to reduce the 'disagreement' between the response of any two or more plies. Woven materials are now widely used, especially for this reason. A second approach is to 'stitch' the composite in the region of non-uniform stress, typically near an edge of the laminate. Stitching simply 'clamps' the edge of the material to prevent it from separating; the internal stresses are still present. Stitching has a somewhat smaller number of proponents, but is a successful method as well. Finally, threedimensional reinforcement, such as mats or braids, also serve the purpose of providing constraint to the delamination drivers. These methods are not as widely used at this time, largely because of the difficulty associated with manufacturing. A less obvious influence on durability and damage tolerance is the bonding between the fiber and matrix. The nature of this influence has only come to light in recent years. Some of the mechanics models needed for this discussion will be developed in the next section; only a few general points will be made here. First, the properties of composite materials are determined not only by the properties of the constituents, but they are also greatly influenced by the manner in which the constituents

802 Durability and damage tolerance of fibrous composite systems interact. This critical interaction is, of course, controlled by the bonding between the constituents, between the fibers and matrix in our case. Typically, this bonding is ‘controlled’ by a fiber coating or ’sizing’. However, it is now known that such things as notched fatigue behavior can be improved by as much as two orders of magnitude by carefully ’designed’ ’interphase’ regions between the fibers and the matrixz6.There are at least two basic concepts operating in these effects. First, if one can toughen the composite by toughening the interphase between the fibers and matrix, the composite is likely to be more durable, as discussed above. Second, the interphase region can greatly influence the local stress state, and reduce the driving force for fiber-matrix debonding. An illustration of that is shown in Fig. 35.11. If one considers the strength of a composite under loads applied perpendicular to the fiber direction, then it is clear that the fiber causes a local stress concentration,in proportion to the difference between its properties and those of the matrix. However, if a coating around the fiber is introduced, this local concentration can be greatly reduced. In fact, for a ’rigid’ fiber, compared to the matrix, it is not surprising that a compliant coating on the fiber will increase the transverse composite strength by as much as a factor of two, and the

strain to failure by as much as a factor of In general, although design rules are not yet fixed, design of the interphase region is a new and important opportunity for the enhancement of the durability and damage tolerance of composite system^^^^*. The final subject in this section is ’failure criteria’; which are used to describe remaining strength. In general, failure criteria are chosen on the basis of the known failure mode. If fiber fracture controls strength, then a suitable criterion may be just the stress in the fiber direction divided by the strength in that direction. If matrix behavior is controlling, a shear stress or combined stress criterion may be appropriate. Figure 35.12 shows a comparison between strength ‘envelopes’predicted by two popular criteria. It is important to note that the inputs to the failure function will, in general, change as a function of time and loading history. The general form of any failure criterion will usually be some function of the ratios of stress in principal material directions to strength in those directions, as mentioned earlier. Under long-term conditions which induce damage, the local stress changes as damage causes redistribution, and the principal values of material strength change, due to such things as constituent degradation or micro-damage. Hence, to calculate damage tolerance by using

Interphase region Criterion:

I Maximum stress Applied

Stress (ksi)

0

90

Angle of Loading (deg)

Composite Fig. 35.11 Schematic diagram of the geometry of the interphase region in a fibrous composite, subiected to loading. ” transverse to the fibers.

Fig. 35.12 Allowable uniaxial loading as a function of angle of loading relative to the fiber direction in a unidirectional lamina, estimated from a maximum stress and a popular effective stress criterion.

Composite micro-strengthand remaining strength models 803 failure functions (or criteria) to calculate remaining strength, one must be careful to use the correct local stress state and material state in those expressions, especially when degradation has changed those states from their initial values.

utility of such models. The example is a recent model of tensile strength. (Figure 35.13). The stress in the broken fibers builds back up to the undisturbed level by shear transfer from the surrounding matrix, composite, and interphase region. That rate of buildup is directly proportional to the stress concentration in the next nearest fibers; if the buildup occurs over a 35.4 COMPOSITE MICRO-STRENGTH AND short distance (a short 'ineffective length), the REMAINING STRENGTH MODELS stress concentration in the neighboring fibers The importance of material principal strengths is great, and they tend to break causing very was noted, and the importance of composite brittle composite behavior. However, if the microstructure in the determination of those buildup occurs over a large distance (i.e. if the strengths has been emphasized. The proper- material around the fiber is very compliant or ties, geometry, arrangement, and bonding of breaks easily ), the strength of each fiber is lost the constituents determine the resulting val- completely when the first local fiber break ues of composite principal strengths. So, if occurs. A model has been developed that those factors are understood, strong, durable, describes the physics and mechanics of this damage tolerant composites can be designed. behavior, which estimates the fiber strength as: That understanding is currently incomplete, 2z0L l / m + l 2 l / m + l m + 1 but some models are available. Such models 4= ( K T ) m +2 are very valuable since they can tell us the preferable way to make composite materials, (1 + m)l/" (35.1) in contrast to how they can be made (the job of (C," + q m - 1 + ... + ly" the materials science community). In this limited space, one example will suf- where a, is the Weibull characteristic strength fice to demonstrate the general nature and of the fibers, z, is the shear stress between the

...'-+1(7r)

Composite

FlbersC

Fiber breaks

4 zt:

EE t

Normal stress In:

:

.

broken fiber

P

nelghboring

I

Average global values away from fiber fracture

Fig. 35.13 Schematic diagram of the local stress distribution around broken fibers in a unidirectional composite.

804 Durability and damage tolerance offibrous composite systems fibers and the matrix (usually taken as the continuous fiber reinforced composites is interphase strength), m is the Weibull shape outlined. A great many details will have to be parameter for the fiber strength distribution, omitted due to space limitations; the interand Cnis the local stress concentration when n ested reader can find them in other fibers are broken together in a local region. publication^"^^. Hence, the tensile strength in the fiber direcA start is to identdy a well-defined failure tion can be estimated on the basis of the mode, as defined earlier. Since damage is disproperties of the constituents and the inter- tributed, this damage mode will be 'typical' of phase region between the fibers and the any 'representative volume' of material; a matrix. If any of those constituent characteris- mechanics boundary value problem on such a tics change, the model can show how the representative volume (RV), as suggested in strength of the composite changes, i.e. the Fig. 35.14 can be 'set'. This RV may be disconmodel can be used to calculate the damage tol- tinuous; i.e. it may have cracks, delamination, erance of the composite if the failure mode is debonds, etc. But some part of it will remain controlled by fiber strength in tension. intact until fracture of the composite, and this Comparable models can be constructed for part of the RV that defines the fracture event is compression failure, and for other failure a 'critical element'. Therefore the objective is mode^^*^^. the calculation of the state of stress and state of material in the 'critical element.' One can write all failure functions, Fa, in that critical element, 35.5 ESTIMATION OF REMAINING and claim that when these failure functions STRENGTH AND LIFE (for each distinct failure mode) predict failure, As indicated earlier, damage tolerance is the composite will fail. defined by remaining strength, and durabilInvoking kinetic theory we can derive an ity is usually discussed in terms of life. In this equation that relates changes in stress state and final section, one approach to the estimation material state with time and loading history to of the durability and damage tolerance of remaining strength, i.e. allow the incorporation

failure modes

Fig. 35.14 Diagram illustrating how experimental observation of failure modes define the representative volume (used to set the boundary value problem) and the critical elements in which all continuum states are defined.

Estimation of remaining strength and life 805 of the explicit time, cycles, and environmental dependence that leads to phenomenological behavior such as creep, creep rupture, fatigue, and aging into the calculation of remaining strength. From thermodynamic principles, the following expression can be derived:

F, = 1 -

lyl(l

- Fa),(

N) n d( $) 1-1

sile fiber failure, and it is assumed that some fatigue behavior of unidirectional material under uniaxial stress in the fiber direction has been measured, a 1-D SN relationship can be derived, of the form:

s, = A + B (log S"

(35.2)

N)p

(35.3)

where, for our example, A = 1, B = -0.1, Su = 100 ksi (the initial ultimate strength), p = 1,and Sa is the applied stress amplitude. Equation (35.3) provides an input, N, into (35.2) since

where in the critical element, F, is the normalized remaining strength, n is cycles, and N is the life of the critical element under the current state of stress and state of material. The methodology of this calculation is shown in Fig. 35.15. Remaining composite strength, F , is calculated directly; life is calculated by the coincidence of Fa and Fr. Numerous comparisons of such calculations with experimental data have been made over the last 10 years or so, and there are a few examples at the end of this chapter. The immediate purposes are served by using eqn(35.2) to examine the effects of some hypothetical change in material state and stress state on remaining strength (damage tolerance). If the failure mode is ten-

State of

where u,, is now the current local ply stress in the fiber direction, and X , is the current local principal material strength in tension, given by eqn (35.1). Substituting eqn (35.4) in eqn (35.2) and assuming that no other phenomena are present (and using j = 1.2, a known typical value), the curve (a) in Fig. 35.16 results. Now, suppose the ply is the critical element in a multiaxial laminate having off-axis plies that crack and 'dump' stress into the critical element as a function of

State of

I

life

Subcritical

Critical

N,

I

N2

reipresentative volume

Fig. 35.15 Schematic flow diagram of the manner in which the MRLife simulation scheme calculates

remaining strength and life.


I

Remaining Strength

-

I

Cycles

Cycles

Fig. 35.16 Calculated remaining strength predictions for (a)0" lamina degradation alone; (b) added

Fig. 35.18 Assumed degradation of fiber strength for the sample laminate.

effect of matrix cracking; (c) added effect of fiber degradation (e.g.by oxidation). cycles, according to the rate shown in Fig. 35.17 (from cracking rates that must be measured or estimated). With this internal stress redistribution, only, added to eqn (35.2), the damage tolerance changes to curve (b) in Fig. 35.16. Of course, if creep occurs in the matrix (perhaps because of increased temperature), in which case the local fiber stress will increase again as a function of cycles to change the form shown in Fig. 35.17. Finally, suppose creep rupture is occurring, driven, for example, by oxidation of the fibers that is reducing the diameter of the fibers, D, in eqn (35.1), in the manner shown in

FiberDirection Stress

40

Stress increase due

'

I 1

4

I

4

5x10 9.999~10 Cycles

Fig. 35.17 Assumed increase in stress in the 0" ply

due to matrix cracking.

Fig. 35.18. Then the strength model, eqn (35.1), correctly integrates that micro-change into the global calculation, and eqn (35.2) shows the damage tolerance to be curve (c) in Fig. 35.16 for that situation. Hence this 'micro-kinetic' approach has the capability to estimate durability and damage tolerance for very complex situations involving combinations of many time and cycle dependent phenomena in composite systems, using a mechanistic approach. An example follows. Using the methods described above, the rate of matrix cracking and the unidirectional SN curve of a carbon fiber reinforced PEEK matrix composite were determined, and used to estimate the remaining strength and life of several different laminates made from that material. Figure 35.19 shows an example of the predicted and observed life for several load levels of a quasiisotropic laminate made from such material and Fig. 35.20 shows comparisons of the predicted and observed remaining strength of such laminates for two load levels and cycles of load application. It can be seen that this approach can produce quite useable results. Many such predictions have been compared using the MRLife performance simulation code based on this

Estimation of remaining strength and life 807

AS-4lPEE K (APC-2) Quasi-Isotropic Notched Fatigue (R=-1) 0.75

0.45

3

I

1

I

4

5

6

7

Log N (Cycles) Simonds B Stinchcomb MRLife (1 989) Prediction 0

Fig. 35.19 Predicted (line) and observed life for a quasi-isotropicAS-4/PEEK notched coupon under fully reversed loading.

Residual Strength at 0.70 Suit

Residual Strength at 0.90 SI,,

11.05

1.05

1.w

1.00

r

r

P E!

F E!

5

5

3

3

0.95

; 2 .-

0.w

d ......................................................

x1 N

0.w

€0 z

0

z 0.85

0.05

P

P

.....................................................

0.85

......................................................

0.80

0.80

2

Cycles M E h Expepnt

5

1 0 2 0

5 0 1 0 0 2 0 0

Cycles Mfih Expepml

Fig. 35.20 Predicted (lines) and observed residual strength of AS-4/PEEK specimens subjected to under fully reversed cyclic loading.

808 Durability and damage tolerance offibrous composite systems 35.6 SUMMARY

This has been a short outline of the physical behavior associated with the durability and damage tolerance of composite material systems, and a few modeling approaches to the estimation and prediction of that behavior have been indicated. It should be noted that there is every evidence that composite materials are remarkably durable and damage tolerant. Fatigue allowables for carbon fiber reinforced polymer composites, for example, exceed those of structural steels, and the durability and damage tolerance of ceramic composites make them the only choice for ultra-high temperature applications in turbines, etc. Understanding of this subject, which is admittedly incomplete, has reached a level that is sufficient to support engineering applications of composites to even the most demanding situations in the most severe environments. In fact, that is exactly the situation in which the application of composites is most beneficial and cost effective. Composite material systems can provide many new opportunities to design for damage tolerance and durability. REFERENCES 1. Life Prediction Methodologies for Composite Materials, NMAB-460. Washington, D.C.: National Academy Press, 1990. 2. High-Temperature Materials for Advanced Technological Applications, NMAB-450. Washington, D.C.: National Academy Press, 1988. 3. Horton, P.E. and Whitehead, R., Damage Tolerance of Composites, Vol. I and 11, Air Force Wright Aeronautical Laboratories, AFWAL TR87-3030, 1988. 4. Daniel, I.M. and Ishai, O., Engineering Mechanics of Composite Materials. New York: Oxford Univ. Press, 1994. 5. Reifsnider, K.L. and Highsmith, A.L., Characteristic damage states: A new approach to representing fatigue damage in composite laminates. In Materials: Experimentation and Design in Fatigue, Guildford, U.K.: Butterworth/IPC, 1981, pp. 246-260.

6. Damage in Composite Materials: Basic Mechanisms, Accumulation, Tolerance, and Characterization, STP 775, American Society for Testing and Materials, (ed. K.L. Reifsnider), 1982. 7. Highsmith, A.L., Stijfrzess Reduction Resulting from Transverse Cracking in Fiber-Reinforced Composite Laminates. Master of Science Thesis. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1981. 8. Reifsnider, K.L., Some fundamental aspects of the fatigue and fracture of composite materials. In Proc. 14th Ann. SOC. Engng Science Mg. Bethlehem, PA: Lehigh University, 1977. 9. Highsmith, A.L., Damage Induced Stress Redistribution in Composite Laminates. PhD Dissertation. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1984. 10. Bader, M.G., Bailey, J.E., Curtis, P.T. and Parviz, A., The Mechanisms of Initiation and development of damage in multi-axial fibre-reinforced plastic laminates. In Proc. 3rd Intern. Conf. Mechanical Behavior of Materials. Cambridge, U.K., 1979. 11. Reifsnider, K.L., Damage and damage mechanics. In Fatigue of Composite Materials, (ed. K.L. Reifsnider), Amsterdam: Elsevier Science Publishers, 1990. 12. OBrien, T.K., Characterization of delamination onset and growth in a composite laminate. In Damage in Composite Materials, ASTM STP 775, 1982, p. 140. 13. OBrien, T.K., Analysis of local delaminations and their influence on composite laminate behavior. In Delamination and Debonding of Materials, ASTM STP 876, 1985, pp. 282-297. 14. OBrien, T.K. and Hooper, S.J., Local delamination in laminates with angle ply matrix cracks: Part I, Tension tests and stress analysis. In N A S A TM 104055,1991. 15. Razvan, A. and Reifsnider, K.L., Fiber fracture and strength degradation in unidirectional graphite/epoxy composite materials. In Theoretical and Applied Fracture Mechanics, 1991, 16,81-89. 16. Razvan, A., Bakis, C.E. and Reifsnider, K.L., SEM Investigation of fiber fracture in composite laminates. In Materials Characterization, 1990,24, 179-190. 17. Lorenzo, L. and Hahn, H.T., Fatigue failure mechanisms in unidirectional composites. In Composite Materials: Fatigue and Fracture, ASTM STP 907, American Society for Testing and Materials, Philadelphia, PA, 1986, pp. 210-232.

References 809 18. Ishida, H. (ed.), Controlled Interphases in Composite Materials, New York: Elsevier, 1990. 19. Warren, R. (ed.), Ceramic Matrix Composites, New York: Chapman and Hall, 1992. 20. Reifsnider, K.L. (ed.), Fatigue of Composite Materials, London: Elsevier Science Publishers, 1991. 21. Jamison, R.D., Fiber fracture in composite laminates. In Proc. lntl. Con$ on Composite Materials VI, 1987, no. 3, pp. 185-199. 22. Jamison, R.D., Schulte, K., Reifsnider, K.L. and Stinchcomb, W.W., Characterization and analysis of damage mechanisms in tension-tension fatigue of Graphite/Epoxy Laminates. In Efects of Defects in Composite Materials, ASTM STP 836, American Society For Testing and Materials, Philadelphia, PA, 1984, pp. 21-55. 23. Tiwari, A., The Development of an Interpretive

Methodology for the Application of Real-Time Acousto-Ultrasonic NDE Techniques for Monitoring Damage in Ceramic Composites Under Dynamic Loads. PhD Dissertation. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1993. 24. Wilkinson, S.P., Liptak, S.C., Lesko, J.J., Dillard, D.A., Morton, J., McGrath, J.E. and Ward, T.C., Toughened bismaleimides and their carbon fiber composites for fiber-matrix interphase Studies. In Proc. 6th Japan-U.S. Conf Composite Materials, 1992. 25. Hedrick, J., Patel, N.M. and McGrath, J.E., Toughening of epoxy resin networks with functionalized engineering thermoplastics. In ACS Advances in Chemistry Series, no. 233, Toughened Plastics I: Science and Engineering, (eds. C.K. Riew and A.J. Kinloch), 1993, pp. 293-304. 26. Swain, R.E., Reifsnider, K.L., Jayaraman, K. and El-Zein, M., Interface/interphase concepts in composite material systems. J. Thermoplastic Comp. Mater., 1990, 3, 13-23. 27. Case, S.W., Micromechanics of Strength-Related Phenomena in Composite Materials. MS Thesis. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1993. 28. Carman, G.P., Eskandari, S. and Case, S.W., Analytical investigation of fiber coating effects on shear and compression strength, symposium on durability and damage tolerance, ASME WAM, (in press), 1994. 29. Jayaraman, K. and Reifsnider, K.L., The interphase in unidirectional fiber-reinforcedepoxies: effect of residual thermal stresses. Comp. Sci. Tech., 1993,47, 119-129.

30. Jayaraman, K., Gao, Z . and Reifsnider, K.L., The interphase in unidirectional fiber reinforced epoxies: effect on local stress fields. J. Comp. Tech. Res., 1994,16(1):21-31. 31. Jayaraman, K., Reifsnider, K.L. and Swain, R.E., Elastic and thermal effects in the interphase: Part 11. comments on modeling studies. J. Comp. Tech. Res., 1993,15(1):14-22. 32. Jayaraman, K., Reifsnider, K.L. and Swain, R.E., Elastic and thermal effects in the interphase: Part I. comments on characterization methods. J. Comp. Tech. Res., 1993,15(1):3-13. 33. Gao, Z . and Reifsnider, K.L., Micromechanics of Tensile Strength in Composite Systems. Fourth Volume, ASTM STP 2256, (eds W. W. Stinchcomb and N. E. Ashbaugh), Philadelphia, PA: American Society for Testing and Materials, 1993, pp. 453-470. 34. Xu, Y. and Reifsnider, K.L., Micromechanical modeling of composite compressive strength. J. Comp. Mater., 1993, 27(6):572-587. 35. Reifsnider, K. L. and Gao, Z., Micromechanical concepts for the estimation of property evolution and remaining life. In Proc. Intern. Con5 Spacecraft Structures and Mechanical Testing, Noordwijk, the Netherlands, 1991, pp. 653-657. 36. Curtin, W.A., Theory of mechanical properties of ceramic-matrix composites. J. Amer. Ceram. SOC.,1991, 74(11),2837-2845. 37. Reifsnider, K.L., Performance simulation of polymer-based composite systems. In Durability of Polymer-Based Composite Systems for Structural Applications, (eds A.H. Cardon and G. Verchery), New York: Elsevier Applied Science, 1991, pp. 3-26. 38. Reifsnider, K.L. and Stinchcomb, W.W., A critical element model of the residual strength and life of fatigue-loaded composites coupons. In Composite Materials: Fatigue and Fracture, ASTM STP 907, (ed. H.T. Hahn), Philadelphia, PA: American Society for Testing and Materials, 1986, pp. 298-313. 39. Reifsnider, K.L., Use of mechanistic life prediction methods for the design of damage tolerant composite material systems. In ASTM STP 2257, (eds M.R. Mitchell and 0. Buck), Philadelphia, PA: American Society for Testing and Materials, 1992, pp. 205-223. 40. Reifsnider, K.L., Evolution concepts for microstructure property interactions in composite systems. In Proc. IUTAM Conf. Microstructure-Property Interactions in Composite Materials. Aalborg, Denmark, 1994.

ENVIRONMENTAL EFFECTS ON COMPOSITES 36 Ann F. Whitaker, Miria M . Finckenor, H a r y W. Dursch, R.C. Tennyson and Philip R. Young

stabilizers, vertical fins and fairings were flight-tested with annual inspections. Composite usage has increased dramatically Satisfactory performance of the composite over the last three decades due to the advanmaterials was noted over fourteen years, with tages of light weight, specific strength and some parts experiencing more than 39 000 stiffness, dimensional stability, tailorability of hours of flight loads. Also evaluated were properties such as coefficient of thermal composite parts from military aircraft such as expansion and high thermal conductivity. the C-130 center wing box, S-76 tail rotors and Environmental effects on these properties may horizontal stabilizer, 206L fairing, doors, and compromise a structure and must be considvertical fin and the CH-53 cargo ramp skin. ered during the design process. Boron/epoxy, graphite/epoxy, Kevlar/epoxy Because of the variety of uses, the composand Nomex honeycomb were used in these ite environment cannot be exactly defined. aircraft and helicopter components. This chapter details the major environmental concerns for the composite designer, problems encountered with these environments in the 36.3 ENVIRONMENTS AND EFFECTS past, and some materials or protective systems effectively used. The use of trade names, how- 36.3.1 BIOLOGICAL ATTACK ever, does not constitute endorsement, either Biological attack on composites may consist of expressed or implied, by the authors. fungal growth or marine fouling. As reported in the literature, fungal growth does not 36.2 HISTORICAL PERSPECTIVES appear to be as damaging as the wet conditions that promote growth. Fungicide has been Use of composites in commercial aircraft mixed in with resins to retard this growth. increased under two NASA programs, the Marine boring organisms do not appear to Flight Service Evaluation Program and the attack glass-reinforced composites. Even Aircraft Energy Efficiency Program, begun in though marine organisms will grow on comthe 1970s. These programs included evaluaposite surfaces, mechanical properties do not tion of environmental effects on the composite appear to be affected, and the fouling can be parts of the Boeing 727, McDonnell Douglas by scraping (Fried, 1969). removed DC-10 and Lockheed L-1011 commercial airComposites with graphite fibers have been craft. Elevators, rudders, ailerons, horizontal used in medical applications for both internal and external purposes. Internal composite Handbook of Composites. Edited by S.T. Peters. Published structures, such as artificial joints or plates for in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7 bone fracture support, must be biocompatible 36.1 INTRODUCTION

Environments and effects 811 or the material may degrade over time. External composite designs, such as artificial limbs or orthotic braces, may experience impact damage, flexural and torsional loading during use. 36.3.2 FATIGUE

Fatigue, either through mechanical loads or acoustic vibrations, can cause crack growth or local defect formation. Fatigue design depends not only on the load, but also on the use temperature range and amount of moisture present. Very cold temperatures (below -50°C (-58°F)) may increase the stiffness of some composite materials, thereby increasing the susceptibility to fatigue damage (Staunton, 1982). Destructive effects of fatigue vary with the composite system tested. One example of fatigue resistance is the B-1 horizontal stabilizer torque box, an all-mechanically fastened hybrid composite structure (Staunton, 1982). Acoustic fatigue testing produced no degradation, nor did the service environment of moisture, mechanical fatigue, and temperature cycling from -12 to +167"C (10 to 260°F). 36.3.3 FLUIDS

Moisture Moisture effects on composites have been studied for decades. Water acts as a plasticizer when absorbed by the matrix, softening the material and reducing some properties of the laminate. Moisture may also migrate along the fiber-matrix interface, affecting the adhesion. Moisture in composites reduces matrix-dominated properties, such as transverse strength, fracture toughness and impact resistance. Lowering of the glass transition temperature may also occur in epoxy and polyimide resins with an increase in absorbed moisture. Debonding can occur due to formation of discontinuous bubbles and cracking in the matrix. Mechanical properties can be reduced even

further if heat is present or if the composite is undercured or has a large amount of voids. Moisture is absorbed into the composite until a saturation point is reached. This has been described as a non-Fickian process, meaning the rate of relaxation in the material due to water absorption is comparable to the diffusion rate of water. As the material properties change, such as a decrease in glass transition temperature, the diffusion process changes. Swelling stresses due to non-uniform water absorption have been investigated (Ashbee, 1989). Volume expansion due to water absorption can be a few percent at saturation. Moisture absorption is usually dependent on the matrix, but aramid fibers will also absorb water. The mechanical properties degrade in relation to the amount of moisture absorbed, with no further deterioration after saturation is reached. Strength reductions in polyester laminates have been found to be 10-1570, while epoxy resins are less vulnerable. In a few cases, drying of the composite restored the original mechanical properties. Testing of a glass/polyester laminate allowed to dry after ocean exposure at 1700 m (5700 ft) below sea level for three years showed little change in compressive strength and modulus, flexural strength and modulus, or interlaminar shear strength (Fried, 1969). Fiberglass composites with either polyester or epoxy resins have been used extensively in marine structural applications because of their strength-to-weight characteristics and resistance to the marine environment. Glass reinforcement is preferred over carbon fibers due to carbon's electrical conductivity, which may result in severe dissimilar metals galvanic corrosion with sea water acting as an electrolyte. MIL-HDBK-l7B, besides providing guidelines for characterizing materials and designing a composite system, contains a wealth of mechanical property and environmental effects data. The effect of moisture absorption or water immersion on weight,

812 Environmental eflects on composites (150°C (302°F)) environment, microcracking occurred. The amount of moisture absorbed, as measured by weight gain, is directly related to the change in mechanical properties. Salt water, antifreeze and gasoline had the most proAircraft fluids nounced effect. Dqmg did restore some, but The aircraft fluid environment consists of fuel, not all, of the strength and modulus. Volkswagen of Germany tested composite hydraulic fluid, lubricants, deicing comsystems for compatibility with gasoline, oil pounds, and water. Polysulfone has been and coolant for engine use (Beckmann and found to be sensitive to phosphate ester based Oetting, 1985). These materials were tested hydraulic fluids. Some polymer resins, such both as pure resin and in fiber reinforced lamas PEEK, may have lower glass transition inate form. Among the materials tested were temperatures after exposure to fluids with a glass, carbon and aramid fibers, high-temperhigh aromatic content. A study of stressed and ature epoxy, polyimide and polyester resin unstressed composite materials (Dexter, 1987) systems. Mechanical properties of the samples evaluated short-beam shear strength and tenwere measured after each 100 h of immersion sile strength after immersion in JP-4 fuel, 1000 h total. Adverse reactions of the carup to hydraulic fluid, a fuel-water mixture and bon reinforced materials with metal parts, oil fuel/air cycling for 5 years. The composites and combustion residues were noted. tested were T300/5208 graphite/epoxy, T300/5209 graphite/epoxy and Kevlar 49/5209 in (k45)s configuration. The Other fluids fuel-water immersion appeared to be the most damaging, reducing the tensile strength Liquids accidentally spilled on composite surof the T300/5209 and the Kevlar/5209 by 11% faces may also affect the mechanical and 25%, respectively. Fuel-water exposure properties. Methylene chloride, found in paint also degraded short-beam shear strength by strippers, may cause severe damage to epoxy resins and a number of other polymers. as much as 40%. Graphite/polyimide composites samples were immersed in various fluids for 10 min then Automotive fluids tested for flexure strength and modulus The automotive fluid environment consists of (Lisagor, 1979).Slight increases in these propgasoline, oil, battery acid, brake fluid, transmis- erties were noted for samples immersed in sion fluid and coolant. A study by University of hydraulic fluid, nitrogen tetroxide liquid, Michigan and General Motors (Springer, monomethyl hydrazine liquid and unsymmetSanders and Tung, 1981) details the effects of rical dimethyl hydrazine. Samples exposed to prolonged exposure of E-glass/polyester and hot hydrazine vapors were degraded beyond E-glass/vinylester to automotive fluids. the ability for mechanical testing. Solvents, Composite samples were immersed in water, bases and weak acids at room temperature do salt water, No. 2 diesel fuel, lubricating oil, not appear to affect graphite/epoxies and antifreeze or gasoline. Property changes mea- Kevlar/epoxies. Molten metal, such as alusured were weight, ultimate tensile strength, minum or titanium, may react with carbon tensile modulus, short-beam shear strength and fibers. shear (flexural) modulus. Specific materials exposed were OCF-E-920-1 polyester/E-glass, OCF-E-980 polyester/E-glass, and vinyl ester/E-glass. In a moist high-temperature

operational temperatures, and mechanical properties such as stress rupture characteristics is discussed.

Environments and efects 813 36.3.4 WEATHERING

effects on composite materials. Over 35 different types of organic matrix composites were Warm, moist climates may affect the perforflown on LDEF during its 69-month mission in mance of composites. Decreases of 10-20% in LEO. The post-flight testing and analysis of tensile strength have been noted for fibercomposites flown on LDEF has become the glass/polyester and fiberglass/epoxy (Graner, basis of understanding the long-term effects of 1982)where the surface resin has been eroded the LEO environment on composites. Data away due to extended weathering. Resins from the Solar Array Materials Passive LDEF which are more weather-resistant have been Experiment (SAMPLE), the Space developed since this study. In a 10-year study Environment Effects on Spacecraft Materials of real-time weathering of graphite/epoxy, graphite/polysulfone, and Kevlar/epoxy Experiment, the University of Toronto (Dexter, 1987), the materials that absorbed the Institute for Aerospace Studies (UTIAS) most moisture were most affected by UV radi- Experiment, and the Space Exposure of Composite Materials for Large Space ation. Erosion due to rain, snow or ice impact Structures Experiment are presented here and may be a problem for some aircraft parts, in the section on design considerations. such as radomes or leading edge parts. Further information may be found in the Coatings, such as polyurethane, may be used LDEF Post-Retrieval Symposium Proceedings to make composite parts more resistant to (Levine 1991,1992). this erosion. A real-time weathering study was performed by Grumman on fiberglass Atomic oxygen parts from E-2A and A-6A aircraft (Staunton, 1982). Length of service varied from 12 to 19 Atomic oxygen is formed through the dissociyears. Effects of weathering were dependent ation of 0, by UV radiation. It is the on the material used and whether a protec- predominant molecular species at 100-1000 tive coating was intact. When a hygroscopic km orbital altitudes. Its destructiveness is BF3400 curing agent was used, the fiber- caused by its strong chemical reactivity combined with its translational energy of 5 eV glass/Epon 828 epoxy lost nearly half of its J) from the high velocity of the spaceflexural strength. Fiberglass/Epon 828 epoxy (8 x craft. Studies on the effects of atomic oxygen with methylene dianiline/benzyl dimethyl have been performed both on samples amine (MNA/BDMA) curing agent performed well, retaining tensile and flexural exposed to the low earth orbit environment and in ground-based simulators. strength. Tensile and flexural test samples The Mass Spectrometer Incoherent Scatter taken from a fiberglass rotodome demon(MSIS) neutral atmosphere model is generally strated the value of good coatings. Where the used for predicting the atomic oxygen fluence paint was intact, the material retained more during a mission. Orbital altitude, inclination, than 90% of its original strength and 82-94% other orbit parameters and solar activity are of modulus. Where the paint had been eroded used as inputs. The amount of atomic oxygen away, the composite retained only 68% of its received by a surface also depends on its orioriginal strength. entation to the RAM or velocity vector. RAM facing composites (facing into the velocity vector) flown on LDEF were sub36.3.5 SPACE jected to an atomic oxygen fluence of The Long Duration Exposure Facility (LDEF) approximately 9 x IOz1 atoms/cm2 which has provided a wealth of information on the resulted in a thickness loss up to 0.013cm low-earth-orbit (LEO) space environment (0.005 in) of material, the equivalent of

814 Environmental effects on composites I

3

Meteoroiddebris impacts

A spacecraft in any orbit is susceptible to micrometeoroid impact. These small particles, fragments of asteroids or comets, may impact at velocities up to 60 km/s (37 mi/s) but average 17km/s (10.5mi/s). Those spacecraft in near-earth orbit are also susceptible to impact from pieces of space junk or debris, also travelling at high speeds. Damage from impact may reduce the strength of composite structures or cause rupture in filament-wound tanks. Damage may consist of cratering, penetration, including penetration of thermal or protective coatings, and spallation Fig. 36.2 is Fig. 36.1 Cross-sectional photomicrograph of of a classic Whipple-type meteoroid/debris atomic oxygen exposed LDEF graphite-reinforced shield before and after particle impact. OMC showing approx. one ply (0.013 cm/O.OOS in) of erosion.

e-

approximately one ply of laminate (Fig. 36.1). Impact particle For unidirectional reinforced specimens, the reduction in mechanical properties was proportional to the reduction in cross-sectional area. The SAMPLE experiment flown on the / One or more protectivetpmpers LDEF contained tensile specimens of graphite/epoxy systems and S-glass/epoxy. The composite systems flown were HMF Pressure wall 322/P1700 polysulfone in a ( d 5 ) s weave, HMS/934 in both 0" and 90" unidirectional Before imDact configurations, and P75S/934, also in both 0"and 90" unidirectional lay-ups. The Sglass/epoxy was flown with and without Debris cloud aluminized thermal control tape as a protective coating. These samples received direct atomic oxygen, as well as UV radiation, thermal cycling and micrometeoroid/debris impacts. Atomic oxygen reaction efficiency was calculated to be approximately 1 x cm3/ atom. The thickness loss due to atomic oxygen Penetration or erosion of the S-glass/epoxy samples in this 0 0 spallation of f pressure wall may experiment is estimated to be 9.14 pm (0.36 Spallation or may not occur mil). The glass fibers are not susceptible to erosion, and thus protect the underlying matrix. After imDact The atomic oxygen reaction efficiency for these samples was calculated to be 0.13 x lowz4 Fig. 36.2 Hypervelocity impact of Whipple bumper cm3/atom. design.

/

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I

Environments and efecfs The potential hazard of a meteoroid or debris particle is dependent on its size, velocity, density and angle of impact. NASA SP-8042 may be used to predict the meteoroid environment encountered by a spacecraft in near-earth orbit, earth-to-moon space and near-lunar orbit. The space debris environment in earth orbit is continually changing as more debris is being added with every launch. At the time of publication, NASA TM-100471 is being used as the definition of the space debris environment. Prediction of damage caused by meteoroid/space debris impact can be made either by using a variety of models, study of impact sites in flight hardware, or simulation in the laboratory. Hydrocodes such as CTH and HULL use finite elements or finite differences to predict penetration or spallation of a composite laminate. Inspection of composite samples from LDEF (Fig. 36.3) revealed 2-5 impacts of

Handbook of Composites, 2nd Edition

Handbook Of Ceramic Composites

Handbook of Ceramic Composites