Engineering Design Handbook Helicopter Engineering Part Two Detail Design AMCP 706-202 (Helicopter Engineering, Detail Design)

U.S. DEPARTMENT OF COMMERCE National Technical Information Servi;e AD-A033 216 ENGINEERING DESIGN HANDBOOK HELICOPTER ...

Author: US Army Materiel Command

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U.S. DEPARTMENT OF COMMERCE National Technical Information Servi;e

AD-A033 216

ENGINEERING DESIGN HANDBOOK HELICOPTER ENGINEERING.

PART TWO

DETAIL DESIGN

ARMY MATERIEL COMMAND,

ALEXANDRIA,

JANUARY, 1976

Reproduced From Best Available Copy

VIRGINIA

351058 AMCPAMPIT

AMCP 706-202'-

ENGINEERING DESIGN

HANDBOOK HELICOPTER ENGINEERIN\

PART TWO DETAI L DESIGN

IIEADEQMRE-USP 1 AINYV MATERIEL COMMAND NATIONAL TECHNICAL INFORMATION SERVICE U.S. DEPARTMENT OF COMMIE," iMNWacw, VA. 2231W

JANUAR IC76

AMCP 706-202

AN1( Parrnpbkei

No. 714S-01

ENGINEERING I)ESIGN HANI)BOOK 20Jnay17 HELCOPERENGINEERING, PART TWO DETAIL. DESIGN

TABLE (N-' CONTENTS Paragraph

EE LISTOF ILL.USTRATIONS................ ......................... I 1ST OF TABLES ... .... .... ................................... FOREW ORD .. .. .. . . .. . . .. . . ....... . . .. . . . . . .Xxxviii PREFACE ... . . . . . . . . . . . . . . .

Xxviii

Xxxi, XXX%

CHAPTEKR I INTROI)1UCTION (HAPTERI

2-I

)

2/

INTRODULCTION ....... M.Al.'E. A.I..........S.

. . .-

2-2 ~METALS .................. ............. ..... .... ....... ...... 2-1 2-2.1 FERROUS METALS.............................................. 2-1 2-2. 1.1 General .................................. ......... ............. 2-1 2-2.1.2 Ca!'bOn Steels .............. -............. ...... ................ 2-1 2-2.1.3 Allov Steels.........................-2 2-2.1.4 StainessSteels....................... ............................ 2-2 2-2.1.5 Precipitation Hardening Sicels ...................................... 2-2 2-2.1.6 Maraging Steels .................................................. 2-3 NONFERROUS METALS ............................. .......... 2-4 2-2.2 2-2.2. 1 General.............................. ........................... 2-4 2-4 2-2.2.2 Aluminum Alloys........... I................I.................. 2-5 Magnesium Alloys....I............................................ 2-2.2.3 2-2.2.4 Titanium Alloys......................................... ......... 2-6 2-2.2.5 Copp.,r anid Copper Alloys ......................................... 2-6 2-7 2-2.3 ELECTROLYTIC ACTION OF DISSIMILAR MFTAI.S................ 2-3 NOMETALLIC MATERIALS........................................2-4-7 2-7 ............................ GENERAL .......................... 2-3.1 22-3.2 THERMOPI ASTIC MATERIALS ................................ 2-3.3 THERMOSETTING MATERIALS......... ................ ........ 2-9 ...... 2-10 2-3.4 ELASTOMERIC MATERIALS ............................. .. ........... 2-10 2-3.5 WINDOW MATERIALS ............................ COMPOSITE STRUCTURES.........................................21-Il 2-4 .. ........... 2-Il FIBERGLAS LAMINATES .......................... 2-4.1 2-4.1.1 Design Considerations ................ ............................ 2-1l 2-4.1.2 Resin Systems ................................................... 2-12 2-12 2-.Ž;Polyesters ...................................................... 2-4.1.2.2 Epoxies ........................................................ 2-12 2-4.1.2.3 Phrnolics .......................... ............................ 2-12

AMCP 706-202 TABLE OF CONTENTS (Continued) Paragraph

Pagc

2-4.1 3 2-4.1.3.1 2-4.1.3.2 2-4.1.3.3 2-4.1.4 2-4.1.4.1 2-4.1.4.2 2-4.1.4.3 '2-4.1.5 2-4.2 2-4.2.1 2-4.2.2 2-4.2.3 2-4.2.4 2-4.3 2-4.3.I 2-4.3.1.1 2-4 .3.1.2 2-4.3.1.3 2-4 .3.1.4 ,,•

Types of Reinforccment .............. ........... .................... Nonwoven Continuous Filaments .................................... Woven Fabric .................................................... Chopped Fiber ..................................................... Fabrication M ethods ................... .............................. O pen Mold H and Layup .............................................. Sprayup ................ ........................... ...... Matched D ie M olding ................................................ Surface F inishes ....................................................... FABRIC LAM IN ATES .................................................. Reinforcem ent Selection ............................................... R esin Selection ........................................................ Special T ypes ......................................................... Specifications ......................................................... FILAMENT COM POSITION .......................................... Types of Reinforcem ent ............... ............................... E-glass ................................................... S -g lass ...................................................... ........ Boron Filam ents ................ ....................... ........... .............. 11lG ..rap hite .............................................

, C

"" ,..

.

.

......

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

2-4.3.3

Manufacturing Processes ..................... ....

2-4.3.4

A pplications ...................

2.44 2-4.5 2-4.5.1

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

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

2-5 2-5.1 2-5.1.1 2-5.1.2 2-5.1.3 2-5.2 2-6 2-6.1 2-6.2 2-6 .3 2-6 .4 2-7 2-7.1 2-7.2 2-7.3 2-7.4 2-7 5

1t

2-19

2-19o .

,

2-20

HONEYCOMB AND SANDWICH CONSTRUCTION . A R MO R MATERIALS ................................................ Available Materials .......................................... Design ADHESIVES AND SEALANTS ........... ........................ BO N D ING AG EN TS ................... .............................. Structural Adhesives ............... ............................... Nonstructural Adhesives ....................................... Processing O perations ................................................. -Dcsgi, ofonded Structurc......................................... SEAL.ING CO MPO UN DS ............................................. PA IN IS A N D FIN ISH ES ................ .... .......................... PAINTS AND COATINGS (ORGANIC) ............................... SPECIAL. FINISHES ............................................... P LAT IN G ......................................................... ... T A P ES ........................... ................ ................... LUBRICANTS, GRI-ASES. AND HYDRAUI.IC FLUIDS ............... G E N ER AL ......... ............................ ..................... DESIGN OF LUBRICATION SYSTEMS ............................ G RE-A SE S ...................................... ...................... DRY FILM AND PERMANENT LUBRICANTS ..................... H YD R AU LIC FI.U IDS ....... ............... ....................... R EF ER E NC ES ...................................................... ...

""4

2-12 2-13 2-13 2-13 2-13 2-14 2-15 2,15 2-15 2-16 2-17 2-17 2-17 2-17 2-17 2-18 2-18 2-18 2-18 2-18

2-20 2-272-29 2-30 2-30 2-30 2-30 2-32 2-33

•

2-33

2-33 2-34 2-34 2-35 2-36 2-3 7 2-38 2-3K 2-38 2-38 2-38 2-40 2-40

(0i1A'ITR 3 PROPII. ION SIBSYSTEM I)IESIN'N

L IST Of: SY M BO LS .....................................................

3-0

3-1

ii

.. . .. .

.

.

.

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r-

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"AMCP 706-202 TABLE OF (ON1I EN'IS(( ontinud) Page

Paragraph 3-1 3-2 3-2.1 3-2.1.1 3-2.1.2 3-2.1.3 3-2. 1.4 3-2.2 3-2.3 3-2.4 3-2.4.1 3-2.4.2 3-2.5 3-2.5.1 3-2.5.2 3-2.5.3 3-2.5.3.1 3-2.5.3.2 3-2.5.3.3 3-2.6 3-2.6. 1 3-2.6.2 3-2.6.2.1 3-2.6.2.2 3-3 3-4 3-4.1 3-4.2 3-4.2.1 3-4.2.2 3-4 .2.3 3-4.2.4 .3-4 2.5 3-4 3-4 .22.6 .7

"3-4.2. 3-4 .3 3-5 3-6 3-7 3-8 3-8.1 3-8.2 3-8.2.1 3-8.2.2 3-8.2.3 3-8.2.4 3-8.2.5 3-8.3 . .3. I

3-

INTRO D UCTIO N .....................................................

3-1

ENGINE INSTALLATION .............................................. G EN E R A L ............................................................ Subm erged Installation ................................................ Sem icxposed !nstallation .............................................. Exposed Installation ................................................... D esign C hecklist .......................................... .......... ENGINE M OUNTING ................................... ............ ENGINE VIBRATION ISOLATION ................................... FIF EW A LLS .......................................................... Fire Detectors ......................................................... Fire Extinguishing ..................................................... ENGINE AIR INDUCTION SUBSYSTEM ............................ Air Induction Subsystem Design ....................................... Inlet Protection ....................................................6.. A nti-icing ............................................................. Electrical A nti-icing .................................................. Bleed Air Anti-icing ............................................... Anti-icing D em onstration ............................................ EXHAUST SUBSYSTEM .............................................. ....... ............... Exhaust Ejectors . .......................... Inftaid (IR) Radidtiui Suppr, .sior ................................... 1R Suppression Requirements . ...................................... Exhaust Suppressor ................................................. PROPULSION CONTROLS ............................................. .................. ... ....... FU EL SU BSYSTEM ................... ................................... ... GENERAl . ............. FLUEL SUBSYSTEM COMPONENTS ................................ Fuel Tanks ................................................. ........ . Fuel Tank Venis ....................... .............................. . F uel G aging ........................................................... Refueling and D efueling ............................................... F uel D um ping ........................................................ ................... Engine Feed System .............................. Fue l l)r iti ....................... ..... .... .........................

3-i 3-1 3-1 3-I 3-3 3-4 3-4 3-5 3-5 3-5 3-6 3-6 3-6 3-6 3-7 3-7 3-7 3-7 3-7 3-8

....................... Controls and Instrumenlation ................ T EST IN G ............................................................. LUBRICATION SUBSYSTEM ......................................... COMPARTMENT COOLING .............. ACCESSORIES AND ACCESSORY DRIVES ............................ ................................ AUXILIARY POWER UNITS (AIU', G EN ERA L ......................................................... APU INSTALLATION DETAILS ..................................... Method of Mounting .................................................. Inlet D ucting ........................ ............................. ... Exhaust Ducting .................................................. ................................ APU Bleed Air Ducting ........ Cooling ..................... . .. ....................... . ............. APU SU BSYSTEM S ....... ......................................... .............................. C ontrols . ...................

3.13 3-13 3-14 3-14 3-15 3-15 3-15 3-15 3-15 3-16 3-17 3-18 3-18 3-18 3-18

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

Electrical

3-8

3-1 3-9 3-9 3-9 3-9 3-10 3-10 3-11 3-11 3-11 3-13 3-13 3-13

Iiii

AMC? 7W0620

___

T'ABI.ELF

(ONr'' "I %I

i( onalinucd)

Paragraph

3-8.31. 3-8v .1.2 3-8.3.1 3 3-8.3.1.4 3-8.3. 1.5 3-8.3.2 3-8.3 2.1 3-8.3.2.2 3-8.3.. 3-8.3.4

3-8.3.5 3-8.4 3-8.5

____

Pdgv

Sequencing Controls ................................................. Protective Controls ....... ............................ .............. O utput C ontrols ..................................................... Electrical Control Location ....................................... ... Electrical Power Requirements ...................................... Fuel System Controls .................................................. Rated Speed G overning .............................................. Filtering Requirements ............ ................................ APU Lubrication Subsystem .......................................... A PU Reduction Drive ..... ............................... ...........

3.18 3-18 3-18 3-19 3-19 3-19 3-19 3-19 3-20 3-20

AP Starting.................

3-20

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

R ELIA BILIT Y ........................................................ SAFETY PROVISIO NS ..................................... .......... R EFER EN C ES ..........................................................

3-20 3-21 3-22

CHAPTER 4

TRANSMISSION AND DRIVE SUBSYSTEM DESIGN 4-0 4-I 4-1.1 4-1.2

4-1.2.1 4-1.2.1.1 4-1.2.1.1.1 4-1.2.1.1.2 4.1.2.1.1.3 4-1.2A. 1.4 4-1.2.1.2 4-0.21.3

"4-1.2.1.4 4-1.2.1,4.1 4-1.2.1.4.2 4-1.2.I 4.3 4-1.2.1.4.4 4-1.2.1.4.5 4-1.2.2 4-1.2.2.1 4-1.2.2.2 • : 4-1.2.2.3 4-1.3 4-1.3.1 4-1.3.2 4-1.3.3 4-1.4 4-1.4.1 4-1.4.2

"4-1.4.2.1 4-1.4.2.2 4-1.4.2.3 iv

liSTOF SYMBOLS .............................................. INTRO D U CTIO N ....................................................... G EN ER Al. ........................................................ ... REQUIREM ENTS ................................................... . G eneral Requirements .... ......................................... Pe-fortrance ......................................................... Subsystcni Weight ........................ ..................... Transm ission Efficiencr y ................................. ........... Si Ce Levels. ..................... ........ .......................... N oiseLevels ....... . .............................................. R eliability ..... ................................................. Maintainabifity.................................................. Survivability ............................................... R edundancy ........................................................ Dcsigu Configuration ............................................... Self-sealing Sum ps .................................................. Emergency Lubrication ......................... ............ ....... A rm or .............................................................. Drive System Configurations ......................................... Single M ain Rotor Drive System ..................................... M ultilifting-rotor Drive System s ...................................... Compound Helicopter Drive Systems ................................. TRANSMISSION DESIGN AND RATING CHARACTERISTICS ..... Power/Life Interaction ................................................ Transmission Overhaul Life Rating ..................................... Transmission Standards and Ratings ................................... QUALIFICATION REQUIREMENTS ................................ Component and Environment............... ........................ Dcvelopment Testing ...................... .......................... Static C asting Tests ................................................... Deflection Tests ...................................................... C ontact 1 :sts ........................................................

4-1 4 3 4-3 4.3 4-3 4-4 4-4 4-4 4-I1 4-11 4-12 4-16 4-7 4-I1 4-18 4-22 4-22 4-23 4-23 4-23 4-23 4-25 4-25 4-25 4-27 4-28 4.29 4-2'9 4-29 4-29 4-29 4-30

( .•

4 :.

AMCP 706-202 lARIl.01- (()%'I IKNIS (Oininuvdli

Paraigraph 4-1.4.2.4

4-1.4.2.5 4-1.4.2.6 4-1.4.2.7 4-1.4.3

.L4-1.4.4

4-2 4.2.1. 4-2.1.1 4-2.1 .2 -4-2.1.2.1 4-2.1.2.2 .2.3 4-2.2 4-2.2..1 4-2.2.1.1

It4-2.1

4-2.2. 1.2. 1 A

-Bending Fatigue Strength ............................

...........

ailuir .................................................................

4-30 4-30 4-30 4-31 4-31 4-32 4-32 4-32 4-32 4.34 4-34 4-34 4-34 43 4-34 4-35 4-36

Cattie Failure.................................................

4-44

4-2.2.1.2.3.2

Classic or Pitch Line Fatigue ....................................

445

4-2.2.1.2.3.3

Wear Initiated Failure

446

4-2.2.2.2.3 4-2.2.2.3 A-2.2.3 4-2.2.3.1 4-223.2 4-2.2.3.3 4-2.2.3.4 4-2.2.3.5 -4A-2.2.4

4-2.2.4.1 4-2.2.4.2. 4-2,2.4.3 4-2.2.5 4-2.2.5.1 4-2.2,5.2

e:4-2.3 4-2.3. 1

.........

I...........I.............

r~

44.

4-2.2.1.2.3.

''4-2.2.2.2.2

)

Assembly and Disassembly ....................................... Lubrication System Debugging.................................... Incremental Loading and Efficiency Tests ........................... Thermal Mapping Tests.......................................... Overpower Testing............................................... Other Life and Reliabil~ty Substantiation Testing .................. .... TRANSMISSIONS ................................................ FAILURE MODES .............................................. Primary Failure Moocs ........................................... Secondary Failure Modes ......................................... Overload Failures ............................................... Debi is-caused Failure ........................................... Environmentally Induced Failures ................................. DYNAMIC 1C.CNPONENTS .................................... Gears Limi.....ations............................................ Gear Lnialyios ................................................ aiukgl;I

4-2.2.1.3 4-2.21.2 4-2.2.2. 1. 4-2.2.2..1 4-2.2.2.1.2 4222.3 4-2.2.2.1.4 '4-2.2.2.2 4-2.2.2.2.1

Y

Pp

Gear Drawing and Specification................................... 4-4 Bearings ....................................................... 4-48 Lubrication Deighnq.......s.....................................4-4O AMpluctiong Driesign ......................................... .. 4-48 Mounrctiong Practices .......................................... 4-48 Internal Characteristics ......................................... 45 Skidding Control .............................................. 4-54 Life Anm lysis ................................................. 45 Assumptions and Limitations .................................... 4-55 'Modification Faiflot Approach to Lifc Prediction..............I...... 4-55 Complete Elastic and Dynamic Solutions........................... 4-56 Drawing Controls........................................ ...... 4-56 Splines.........................................................44-7 Face Splines ......................................... ......... 4-386L Concentric or Iongitudinal Splineq ................................. .5 :Propcrtics ofSplineb ............................ ................ 4-58 Spline Strength A nalysis..................... .................... 4-59 Drawing Design and Control ..................................... 4-60) 4-60 -Overrunning Clutches ............................................ Sprag Clutches .................................................. 4-61 Ramp and Roller Clutches........................................ 4-62 Self-encvgizing Spring Clutches.................................... 4-62 Rotor Brakes........................ ........................... 4-62 R,ýquirernents and Limitations.................. .................. 4-62 Design arid Analysis ........................................... 4-63 STATIC COMPONENTS.................................... ..... 4-64 Casez end Housings.............................................. 4-64 v

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0 AK 70_2o2 TABLE 0 CONTENTS Iedlnutd) P lParagraph

Pug-

4-2.3.1.1

4-4.2 4-4.2. I 4-4.2.2 4-4.3 4-5 4-5.1 4-5.2 4-5.3

4-64 4-66 4-67 4-W 4-0. 4- 3 4-72 4-72 4-72 4-73 4-73 4-74 4-76 4-76 4-76 4-80 4-81 4-81 4-82 4-82 4-83

J 3.rt..C. a.^CO*

44 13.3

--

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

M ate:ials and Pro .............................................. Quills ...................................... .......................... SPECIAL CONSIDERATIONS ........................................ Vibr'ation Control ..................................................... D iagnostics ........................................................... DRIVE SHAFTING AND INTERCO iNECTION SYSTEMS ............ GENERAL. REQUIREMENTS ........................................ Engine-to-Transmission ............................................... Interconnect Shafting .................................................. Tail Rotor or Propdlcr Shafting ....................................... Subcritical Shafting ........................................ .......... Supercritical Shafting ................... .............................. COMPDONEi-NT DESIGN ............................................. C ouplings ............................................................. Bearings ................................................... Shafting .............................................................. LUBRICATION SYSTEMS ....................................... OIL MANAGEMENT ............................................ Function .............................................................. Comporent and Arrangement .........................................

Desi.5n ai'd Analys-u

4-2.3.1.2 4-2.3.2 4-2.4 4-2.4.1 4-2.4.2 4-3 4-3.1 A-3. I. I 4-3.1.2 4-3.1.3 4-3.1.4 4.3.1.5 4-3.2 4-3.2.1 4-3.2.2 4-3.2.3 4-4 4-41 4-4.1.1. 4-4.1.2 .

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......

COOLING REQUIREMENTS ......................................... Heat Exchanger Sizing ................ ................................ Cooling Fan Sizing ........................................... EMERGENCY LUBRICATION.................................. ACCESSORIES......................... ............................ PAD LOCATION AND DESIGN CRITERIA................ ......... ACCESSORY DRIVE DESIGN REQUIREMENTS .................... REQUIREMENTS SPECIAL F.........................................

4s

4-86 4-36 4-87 4-87 4-88 4-88 4-89 4-89

QC

(

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K

CHAPTER 5

ROTOR AND PROPELLER SUBSYSTEM DESIGN

vi

5-0 5-1

LIST OF SYM BOLS .................................................... IN TRO D UCTIO N .............................................. ........

5-1 5-2

5-2 5-2.1 5-2.1.1 5-2.1.2 5-7-!.3 5-2.1.4 5-2.1.5 5-2.1.6 5-2.1.2 5-2.1.8 5-2.2 5-2.3 5-2.4 5-3 5-3.1

DESIG N PARAM ETERS ................................................ HOVER Disk Loading and Induced Power ...................................... Blade Loading ....................................................... Blade Tip M ach N um ber ....................... ..................... N um ber of Blades ......................... ........................... Tw ist .............................................. .................. A irfoil Sections ......................................... .............. Hovering Th vust Capability ............................................ ........................................... Guideines .......p HIGHu SPEED LEVEL F. IGHT ...................................... HIGH-SPEED MANEUVERING FLIG:T ........................... IN E R T IA .................................. ........ .................. ROTOR SYSTEM KINEMATICS ....................................... G EN ER A L ..................................................... ... ...

5-3 5-3 5-3 5-4 5-5 5-5 5-5 5-5 5-5 5-5 5-6 5-6 5-7 5-7 5-7

(

-W

W4

706-j22

_________AlMiýP

SA OLL OF (CONTENTh I(vatinuLedI PararaphPatic

5-3.

11 LICOPTER CONTROLa.......................

5.9

5-3.j

ARTICUL,%TID R0104.........................

59

5-3.4

GtNMBALED(TEI-TERING)HO1TOR

5-10

5-3.5

HIINGELESS ROY OR .........................

5-3.5.1 5-3.5.2 5-3.6 5-3.6.1 *

-5-3.6.2

5-3.6.3 5-4 5-4.1 5-4.1.1 54.1.2 5-4,1.2.1 5-41.2.2) 5-4.1.2.3

-.

Fr~tigue Tests ....................................................

5-4.3

GROUND RES.NONANCE

f

>

AS

rrQ

wlS.................................

To-blded otorWithHinged Blades ............................... Two-bladrd Riatocs WVitbout Hinges .................................. M ultibiaded Rotors ........ .... .. ... ......... FLUTT FR ASSESSMENT .......................................... Current Criteria................................................... Design Considerations ............. ..................... .......... Helicopter.............. ........ ...................... ......... 5-4.4.2.1.1 FaAed SystEm ................................................... 5-4.4.2.1.2 Ro'ating System ... ............................................ 5-4.4.2.2 Compound ...................................................... 5 -4.4.2.2. i Fixed System ................................................... 5-4.4.2..72 Rotating System ................................................ 5-4.5 ACOUSTIC LOADING ............................................ 5-4.6 GUST LOADINGS................................................ 5 4.6.1 Discussion of the Gust Prohi -m...................................... 5-4.6.2; Guist Design Considerations......... ............................... 15-4.7 TORSIONAL STABILITY ............................. ............ 5-4.7. 1 Discussion of Problem . ...... .. . . . .. ... ........ 5-4.7.2 De!sign Considerations ....... .. . . . ... .. ........ 5-5 BLADE RETENTION .......................... 5-51I RETENTION SYSTLM DESIGN CONSIDElRATIONS................. 55 .1Articulated Rotors ... . . .. . .. . . . . . . .. .. . .. . . . Articluated Rotor Considerations ............... -51.1Typical 5-5.1.1.2 Reversed Hinge Articulation ...................... 5-5.1.2 Gimbaled and Teetering Rotors ... ................. 5-5.1.2.1 Gimbal-mounted Hubs .......... .............. 5-5.1.2.2 Teetering Hubs . . . . . . . .. . . . . . . . . . . . . . . . . .. . 5-5.1.3 Rigid Rotor . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 5-5.2 COMPONENT DESIGN CONSIDERATIONS ............. 5-5.2.1 Rolling Element Bearings ..... .................. 5-4..1

5.4.3.2 5-4.3.3 5-44 5-4.4.1 5-4.!.2 5-4.4.2.1

5-11

XH-51 Rotor System .......................... 5-12 OH-6A Rotor .. .... ... .. . . .. . . ... .... ... 5-12 ROTOR SYSTEM K;NEMA IC COU?U]N!G ......................... 5-13 Pitch-lag Instability ................................................ 5-1; Pitch -flap Instability .............................................. 5-14 Flap-1iit Instability ............................. .................. 5-14 ROTOR SYSTEM DYNAMICS ....................................... 5-!6 OSCILLATORY LOADING 01- ROTOR BLADES .................... 53-16 Hyotheicator LoadDsigcdinCofsid oratibns..............................5-16 Osilltoypohtiad DsesaigConsidrations V ratoy.Load....................51 .. Rotor Oscitlatory Load Calculation....................... .......... 5-17 Drawing Pomid Phase ....................... .. .................. 5-17 Flight rests ..................................................... 5-18

5-4.1.2.4

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_ ................

5-18 5-19

5-2 1 5-22 5-22 5-23 5-23 5-23 5-23

:

1 '

5-23 5-23 5-23 5-23 5-24 5-24 5-24 5-24 5-25 5-2ý 5-26 5-1) 5-27 5-27 5-27 5-27 5-29 5-29 5-29 5-30 5-30 5-30 5-30

.

"

AMCP 706-22 "1AII: kI F OFl

()N"t I-NS t( outinu'd i

Piaragraph

5-5.2.1. 5-5.2.1.2 5-5.2.1 3 5-5.2.2 5-5.2.3 5-5.2.3.1 5-5.2.3.2 5-5.2.4 5-5.2.5 5-5.2.6 5-5.3 5-5.4 5-5A.1 5-5.4.1.1 5-5.4.1.2 5-5.4.2 5-5,,4,3 5-6

5-6 IF

5[ -t,.t 5-i. 1 2 5-6.1.3

5-6.2 5-11! 2. 1 5-6.2. I .I 5-6.2.1.2 5-6.2.1.3 5-6.2.!.4 5-6.2.1.5

5-6.2.1.6

Ig

C_')hndrical Roller Brarinis ........................................... Tapered Roller Bearings .............................................. Angular Contact Ball Bearings ........................................ Tcflon F-abric Bcaring. ................................................ Flexing Elk nent.' ...................................................... Tension-torsion Strap Assemblies .................................... W ire Tic-bar Assem blies ............................................. Flastom criý" Bcarings ..................... ............................ Lag D)ampcrs, Lead-lag Stops ......................................... Droop and Flap Stops and Rcsti ainers ................................. CONTROL SYSTEM CONSIDERATIONS ............................ BLA D E FO LD ING .................................................... D esign Requirem ents .................................................. M anual Blade rolding ........................................... .... Pow er Blade Foiding ................................................. O perational Requirem ents .............................................. System Safet) Considerations .............. ............................

R OTO R BLA D ES ........................................................

N q, I

5-6.2.4

5-6.2.5 5-6.2.6 5-6.2.7 5-6.2.8 5-6.3 5-6.3.1 5-6.3.2 5-6.3.3 5-6.4 5-7 5-7.1 5-7.2 5-7.2.1 5-7.2.2 vWi

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

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

BLADE CONSTRUCTION ........................................ S p a r ................................... .............................. H ollow Extrusion .................................................... Solid Extrusion ..................... ................................ Formed Sheet Metal .................................................. R ound S,ecl T ube ...... ............ ................................ M olded Reinforced Plastic ............................................ le I UvI I .........................

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

.

5-37

5-38 5-38 5-39"•

(

1 : ':'

, ':

5-41 5-4 1 5-41 5-41i-_ 5-42 5-42

, S,

5-42

I ....... .

C ontinuous Skins .................................................... Segm ented Skins ............................... ................ .... W raparound Skins ................................................... ............................... Root End Retentions ................ Tip Closures and Hardware .....

5-37

5-42

Form ed M etal Tube .................................................. .2...rt

5-6.2.2.1 5-6.2.2.2 5-6.2.2.3 5-6.2.3

. ......

"Tw ist ................................................................. Planfurni " apt~r ... ..... .............................................. A irfoil C ross Section ................... ...............................

5-31 5-31 5-31 5-31 5-32 5-32 5-32 5-32 5-34 5-34 5-35 5-35 5-35 5-35 5-3t 5-36 5-36

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

T rim T abs .................... ....................................... T uning W eights .............. ........................................ D esign Requirem ents .................................................. Tooling and Quality Control P equirements ............................. BLADE BALANCE AND TRACK ..................................... Effect of D esign ............... ....................................... Com ponent Lim it W eights ............................................. 1 rack ................................................................. ROTOR BLADE MATERIALS ........................................ ROTOR SYSTEM FATiGUE LIVES .................................... G EN E R A L ........... ................................................ ENDURANCE LIMIT TESTING ...................................... G encral ............ .................................................. N onm etals ........ .................... ..........................

5-43 5-4 3 544

5-44 5-44

5-44 5-4 5 5-4 5 5-45 5-46 5-4 6 5.47 5-4 9 5-50 5-53 5-53 5-54 5-54 5-55

'

.-

,

1

F~ AMCP ""1 ABI IV!

(

706-20

%'11 iN SI( onfilud

PIr.iagraph

,

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

5-57

5-8

PROP

II. RS .......

.... ..

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

5-H. I 5-8.2 5-3.2. 5-8.2 2 5-8.2.3 5-8.2.4 5-8.2.5 5-8.3 5-8.3.1 5-,.3. 5-8.3.1.2 5-8.3.1.3 5-S.3.2 5-8.3.2.1 5-8.3.2. 1.1 -

GI NL.RAI ............... .................................... PROPELLER SYSI ENI )YNAMICS .................................. Vibrator) L.oads ...................................................... . Critical Speeds and Resiporisc .................. ....................... ....... ............................... G usis and M aneuvers ...... Stall F luttcr ....................................... .................. Prop Iller Roughness ...... ....... ............................ .... PROPI-I.iI-IR HUBS, ACTUATORS, AND CONI ROLS ............... Propeller Barrel and hladc Rctention . ................................ B arrel L.oading ....................................................... Louding D efinition .......................... ....................... Barrel Structural Tests ........................................... Propclcr Actuators and Controls ...................................... C ontrol Configurations ................... ........................... Constant-speed G overnors .......................................... .. ........................................ BnaContr,..

5-57 5-s5 5-57 5-60 5-62 5-63 564 5-65 5 65 5-65 5-66 5-66 5-66 5-66 5-66 5-67

5-8.3.2.3 5-8.3.2.4 -5-8.3.2.5 5-8.4 5-8.4.1 5-8.4.2 5-8.4.2.1 5-8.4.2.2 5-8.4.2.3 5-8.4.3 5-8.4.4 5-9. 4.4.1 5-8.4.4.2

5-67 5-68 5-68 5-18 5-68 5-70 5-70 5-7.2 5-73 5-73 5-7. 5-74 5-75

5-8.4.4.4 5-8.5 5-8.5.1 5- .. .S

Au xiliary Functio s .................................................. ......................................... Control Performance C ontrol R eliability ................................ .................. PROPELLER BLADES ......... ................................ Blade Geom etry ...................................................... Blade C onstruction .................................................... ade Constructior .......................................... Types of BM M anufacturing Processes and Tooling ................................. Q uality C ontrol ........................................... .......... Blade and Propeller Balance ....... ....... ........................... Bvide M aterials ................................. ..................... tl low B lades ....................................................... C om posite M aterials ................................................. ................. F iller M aterial ..................................... Structural Adhesives ....................... ......................... PROPELLER BLADE FATIGUE LIVI:S ............................. Enduranct Limit and O0her Structural T',:sting .......................... ... .......................... ........... . T i A TIst C... I.

5-75 5-75 5-76 5.7

5-8.5.1.2 5-8.5.2

F ull-scale Tests .................................... .................. tigu Life Dctcrminat on ................ Flight Loads Test D-aon

5-76 5-76

IO p erpretation of ResutsI ..............................................

5-77

5-9.4.1 5-9.4.2

Tractor C onfiguration ................................................. ................... Pusher C onfiguration ...............................

5-79 5-79

5-9.4.3

O perational Considerations ............................................ D irection of R otation ..................................................

5-79 5-79

5-,.4.4.3

5-H.5.2. I

5-8.5.2.2

)•

5-56 5-56

Structurai M em bcr I. .............. 5-7.2.4 Dctermination of Fatigue .ifc

5-8.3.2.2 --

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

5-7.2.3

5-9 5-9.1 5-9.2 5-9.3 ,5-9.4

S5-9.4.4

Itydraulic System

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

A ircraft T ests ........................

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

. ........

. .......

............................ AN -IITORQ U E ROTORS ................... . ........... ........... .................... G E N E R A L ............... .......... TYPICAL. ANTiIORQU)E ROTORS ........................ TAIL ROTOR DESIGN REQUIR!M ENrS ........................... .... ....... INSTAILLATION CONS!IDERATIONS ....................

5-67

5-75

5-76

5-7"7 5-77 5-78 5-78 5-79

.,

AM^P 706-202

f 1 AlI.t OE ( Oi11 -N'i S Iollhunitvd (

Paragraph

Pap" Engint E.xhaust

5-9 4 5

5-9.5 5-9.5.1 5.9.5.2 5-9.5.3

"5-9.5.4 5-9.5.5 5-9.6

5-9.7 5-9.7.1 5-9.7.2 5i-9.7.3 5-".7.4 5-9.7.5

* I

-. -_.

6-0 6-i 6-1.1 6-1.I.I 6-1.1.2 6- .1.3 6-1.2 6-1.3 6-2 6-2.1 6-2. 1.1 6-2.1.2 6-2.1.2.! 6-2.1.2.2 6-2.1.2.3 6-2.1.2.4 6.2.1.3 6-2.1.4 6-2.1.5 6.2.2 6-2.3 6-3 6-3.1 6-3.1 .1 6-3.1.2 6-3.1.3 6-3.1.4 6-3.1.5 6-3.1.6 6-3.1.7 6-3.2 6-3.2.1

.A

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

. 80

TAIL ROTOR DESIGN PARAMETERS .............................. Tail Rotor D isk I oading ............................................... T ail Rotor Tip Speed .................................................. ....................... .......... Bladc Nuinocr and Solidity ..... T wist ............. ................................................... 0lade Airfoil t........................................... ...... TAIL ROTOR PI RIORMANC ....... ......... STRUCTURAL CONSID'ERATIONS ............. ............. Struclural D ynam ics ............................ ..................... . Structural Loading ....................... ....................... .. .. Blade Structural Analysis ............. .. .. ......................... A .- uelasticit% ............ ............................................ Flutter and Divergence .......................................... R EFER EN C ES .. ............................... ..... .................

5-89 5-80 5-80 55-8I 5-RI 5.2 5-82 5.82 5-82 5-83 5-83 5-83 5-83

(IAlv IKR 6 FI.V;HT CONTROL S. 1•SVS,,iLIST O t SYM BO LS .................................................... GENRA, .L................................................... ......... ........... D ESIG N M ETIIO ). ............... I Point of Dcpartuie ............... .............................. Mission Requirements and Fligh" Envelope .......................... Basic ttclicoptc" D ata ............................. ........... ...... ANALYTICALTOOLS .......................................... SIMULATION AND TESTING .................................... STABILITY SPECIFICATIONS ........................................ C RITERIA AND METHOD OI'.\V'AI YSIS ........................... Control Power and Damping .......... .............. ............... C haracteristic R oots ................................................... R om Plot, .............................................. .......... Modes and Required Damping ..................................... Inherent Airfram e Stability . ....................... ................. Variation of Para;m eters ... ................... ...................... T ype of C ontrol ............................. . . ..................... Transient Response .. ......... ............. ................. Other Factors ................................................ AIUTOROTATION INTRY . ..................................... SYSTEM FAI URLS ............................ ............ ...... STA BILI FY AUGMENTATION S%Si'i-MS .......................... .. GENERAI ..................... .......... ............ ..... B ell Stabiliter Bar ............... ..................................... H Servo GRotor ............. ......... . ............ ................... ......... M iller echanical -,ro .............. ...................

6.6 1 6-i 6-I 6-1 6-2 6 -6-2 6.2 6-2 6-2 6-i 6-3 6-4 6-4 6-6 6-6 6-6 6-6 6-7 6-8 6-8 6-9 6-9 6-9 6-9 6-9

Lockheed C ontrol Gyro ............................ .................. Elcctrohydraulic SA S ...................... ......................... . Fluidic and Hydrofluidie SAS ....... ................................. Flapping M oment Feedback ....... ......... ..................... . CRITERIA FOR SELECTION .................................... A ugrm entation Requirem ents ................................. .........

6-10 6.13 6-10 6.10 6-10 6-10

(

r

AMCP 706.202 TABLE OF CONTENTS (Continued) Pagt

Paragraph

6-3.2.2

H elicopter Size ........................................................

6-10

6-3.2.3 6-3.2.4 6-3.2.5

Type of Rotor System ................................................. Helicopter Configuration .............................. ............... Suppression of Structural and Rotor Mode Responses, Vibrations. o; G usts ............................................................. SAS RELIABILITY .................................................... ... ..................................................... S afety ..... SA S Fa:lures .......................................................... Fail-safe Principles .................................................... Battle Damage, Vuhnerability ........................................... .......................................... C OST ....................... Developm ent Cost ... ................................................. Production C ost ....................................................... M aintenance Cost ..................................................... TECHNICAL DEVELOPMENT PLAN ................................ PILO T EFFO RT ............................ ............................ CRITERIA FOR POWER CONTROLS ................................ Control Forces .............................................. Vibration lFeelback ............................................ .......................................... K in .inatic Effects ........ Control Stiffness ........... .......................................... HANDLING QUALItY SPECIFICATION ............................ HUM AN FACTORS .............. ................................... Control Force C ues ............. ......... .................... ....... Developm ental Test ................................. .................

6- 1 6-11

6-3.3

S6--?.r`.l 6-3.3.2 6-3.3.3 6-3.3.4 6-3.4 6-3.4.1 6-3.4.2 6-3.4.3 6-3.5 6-4 6-4.1 6-4.1.1 6_6-4.1.2

S ,. A

.

) ,

6-4.1.3 6-4.1.4

6-4.2 6-4.3 6-4.3. 1 6-4.3.2

AUTOMATIC CONTROL INTERFACES .............................

6-4.4

VU LN ERA BILITY .................................................... R ELIA BILIT Y ........................................................ M EC H A N ISM S ......................................................... YSTEMS .............................................. ROTATING a Design Factors ........................................................ T est R esults ........................................................... Bench Tests .......................................................... .................. . Test Loads .............................. Instrur"-,ntation ..................................................... Quantity and Selection of Specimens ............................. lnterpretation of D ata ............................................... Flight T ests ........................................................ Required Instrum entation ........................................... Flight Cond1 .ions ........................................... NONROTATING SYSTEM ........................................... Pilot's Controls to Power Actuator ......................... .......... Power Actuator to the Swashplate ...................................... T R IM SYST EM S ..................................... ................ D isconnectT ri ...................................................... C ontinuous rrim ................................ ..................... Parallel and Series T rim ................................................ SYSTEM DEVELOPMENT .............................................. ....................................... G E N E R A l . .................... MATHEMATICAL MODEL IMPROVEMENT ......................

6-4.5 6-4.6 6-5 6-5.!i 6-5.1.2 6-5.1.2.11 o-5.l.2.1.1 6-5.1.2.1.2 6-5.1.2. 1..1 6-5.1.2.1.4 6-5.1.2.2 6-5.1.2.2.1 6.5.1.2.2.2 6-5.2 6-5.2. 1 6-5.2.2 6-5.3 6-5.3.1 6-5.3.2 "• 6.5.3.3 6-6 6-6.1

6-11 6-12 6-12 6-12 6-13 6-13 6-13 6-13 6-13 6-13 6-13 6-14 6-14 6-14

6-1A 6-14 6-15 6-15 6-17 6-17 6-17

6-17 6-18 6-18 6-18 6-18 6-ig 6-21 6.21 6-.21 6-21 622 6-22 6-22 6-2", 622 6-22 6-23 6-24 6-25 6-25 6-25 6-26 6-26 6-26 6-26 xi

-- m.'ll .-.

~

Il

- ...- •t&

'.1

*

0.

I~Ai • ae&ih

"

,,l.a.&

Mkh~msl,.

.-.-...-.-..

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.

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.

-. i

AMCP 7QW5V2 TABIIV OF ('ONI 1.N*1S t CoEn~inut i) Paragraph 6-6.2.1 6-6.2.2 6-6.3 6-6,4 6-6.5

Pagc ..................... Wind Tunnel Test............................ .. I......................... Hardware Bench Tests ................. GROUND-BASED PILOTED FLIGHT SIMULAT ION ................ FLIGHT TESTS................................................... DESIGN REVIEW ................................................ REFERENCES .. ..................................................

CHAPTEFR 7 ELEC'TRICAL SUBSYSTEM DESIGN 7-0 LIST OF SYMBOLS................................................. 7-I INTRODUc-riON.................... ....... ....................... 7-1.1 GENERAL....................................................... 7-1.2 SYSTEM CHARACTERISTICS..................................... 7-1.3 LOAD ANALYSIS ................................................ 7-1.4 LOAD ANALYSIS PREPARATION ................................. 7-1.5 MANUAL FCRMAT .............................................. 7-1.6 AUTOMATED rORMAT ............................. ........... 7.1.7 SUMMARY...................................................... 7-2 GENERATORS AND MOTORS ...................................... 7-2.1IC GEEAI........................... .:-. ................ 7-2.2 AC GENERATORS (ALTERNATORS).............................. 7221Eiectrical Design................................................. 7-2.2.2 Mechanical Dcsign ................................................ 7-2.2.3 Cooling........................................................ 7-2.2.4 Application Checklist .............................................. 7-2.2.5 V aria ble-freqqtency AC Generators................................... 7-2.3 srA RTER/G EN ERATO RS, DC G IN ER ATORS, AND STA RTERS 7-2.3.1 Starter/Generator ...................................... ......... 1-2.3.2 DC Geuerators....................... ............................ 7-2.3.3 DC Starters ...................................................... Boost Starting System .............................................. 7-2,3.4 1-2.4

7-2.5 7-2.5.1 7.2.5.2 7-3 7-3.1 7-3.2 7-3.3 7-3.4 7-3.5 7-4 7-4A 7-4.1.1 7-4.1 2 7-4.1.3 7-4.2 7-5 7-5.1 7-5.2

ELECTICAL MOTORS............................

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

ELECTRICAL SYSTEM CONVERSION............................. AC to DC Converters ............................................. DC toAC Converters .............................................. ...................................... BATI'LRIES ................. Bm rERY CHARACTERISTICS ................................... ~ GENERATOR CONTROL BATTERY CHARGING................... .......... UTILIZATION LOAD ANALYSIS........................ HEAVY CURENT STARTING REQUIREMENTS .................... MAINTENANCE ......... ..................................... VOLTAGE REGULATION AND REVERSE CURRENT RELAY ... DC VOLTAGE REGULATION ..................................... Voltage Regulator ....................... ............ ............ Reverse Current Relays........................................... Overvoltage Relays ..................................... .......... AC VOLTAGE REGULATION ..................................... OVERLOAD PROTECTION ......................................... GENERAL....................................................... OVERLOAD PROTECTION DF-VICES ............................ .

.26 6-27

6-27 6-2ý, 6-28 6-28

7-1 7-1 7-1 7-1 7-2 7-2 7-3 7-3 7.4 7-4 7-4 7-6 7-6 7-6 7-7 7.8 7-8 7-9 7-9 7-10 7-11 7-1l 7-13

7-14 7-14 7-15 7-15t 715 7-15 716 7-17 7-18 7-11... 7-18 7-18 7K 7-19 7-19 7-19 7-19 7-20

l

A

F .AMCP

706-202 1IABI 1. 01 (()N'II:NI I,(

nttinued)

Paragraph 7-5.2. 1 7-5.2.1.1 7-5.2.1.2 7-5 2.2 7-5.2.3

"7-5.2.4 7-5.3 7-6 7-6.1 7-6.2 7-6.3 7-6.4 7-6.5

"7-6.5.1 7-6.5.2 7-6.5.3 7-6.5.3.1 7-6.5.3.2 -6.5.3.3 7-7 7-7. i ?7-7.2 7-7.3 '• 7-7.4 7-7.5 7-7.6 7-7.7 7-7.8

"7-8

7-8.1I 7-8.1. 7-8.1.1.1

7-8.1. 1.2

7-8.1.1.3 7-F8..1.4 7-8.1.1.5 7-8.1.1.6 7-8.1.1 .7 7- . Q.2 7-8.2 7-8.2.1 7-8.2.2 7-9 7-9.1

"1-9.2 7-9.3 7-9.4

Circuit Breakers ................................................... 7-20 Therm al C ircuit Breakers ............................................. 7-20 7-20 ............... M ars.czic Circuit Breakers .......................... Remote Control Circuit Breakers ....................................... 7-20 C urrent Sensors ...................................................... . 7.20 F uses .............................................................. . . 7-20 OVERLOAD PROTECTION APPLICATION ......................... 7-21 ELECTROMAGNETIC INTERFERENCE (EMI/EMC) ................. 7-21 G E N E R AL ............................................. ............... 1-21 ACCEPTABILITY REQUIREMENTS ................................. 7-21 INTERFERENCE SPECIFICATIONS ................................ 7-21 !NTERFERENCE SOURCES ........................................ 7-21 INTERFERENCE SUPPRESSION ................................... 7-22 Interference-free Components ......................................... 7-22 Equipment Isolation and Cable F outing ................................ 7.22 Source Suppression and Susceptibility Reduction ..................... 7-23 G rounding and Bonding .............................................. 7-23 Shielding ............................................................ 7 -23 F ilters ............................................................... 7-24 ELECTRICAL SYSTEM INSTALLATION ............................. 7-24 -'% N E R A i .. ............................................................ h IE EQUIPMENT INSTALLATION ....................................... 7-24 ELECTRICAL WIRE BUNDLES ....................................... 7-25 TERMINAL STRIP INSTALLATION .............................. 7.25 ENGINE COMPARTMENT WIRING ................................. 7-26 DOOR HINGE WIRE BUNDLE ROUTING ........................... 7-26 WIRING TO MOVING COMPONENTS .............................. 7-26 BATTERY INSTALLATION .......................................... 727 CO M PO N EN TS .............................. ......................... 7-27 7-27 W IR E ................................................................. W ire Insulating M aterials ...... ....................................... 7-27 Polyethylene ......................................................... 7-27 Polyvinylchiorid¢

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

Fluorinated Ethylene Propylen ................................. Polychlorotrifluoroethylene ............... ..................... Polybexamethylene-adipamide ............. .......................... Tetrafluorocthylene .................................................. Dimethyl-siloxanc Polymer M ilitary W ire Specifications ........................................... F I-f IN G . ............................................................ Term it,al Strips ........................................................ Connectors ........................................................ . . LIGHTNING AND STATIC ELECTRICITY ............................ G EN ER A L ............................................................ LIGHTNING PROTECTION FOR ELECTRONIC SUBSYSTEMS .... STATIC ELECTRICITY ............................................... LIGHTNING AND STATIC ELECTRICITY SPECIFICATIONS ...... R EFER EN C E ...........................................................

7-27

7-28 7-28 7-28 7-28 7-28 7-28 7-29 7-29 7-29 7-29 7-29 7-30 7-31 7-32 7-32

xiii

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r• =.-''"

TABLE OF CONTENTS (Conlinued) Paragraph

8-11 I.;.. 8- L2 8-1.3 8-1.4 8-2 8-2.1 8-2.2 8-2.3 8-3 8-3.1 8-3.2 8-3.3 8-3.3.1 8-3.3.2 8-3.3.3 8-3.3.4 8-3.3.5 8-3.3.6 8-3.3.7 8-3.4 9-3.5 I.8-3.6

8-4 8-4.1 8-4.2 8-4.3 8-4.4

Pagc CHAPTER 8 AVIONIC SUBSYSTEMS D~ESIGN INTRODUCTION ................................................. 8GENERAL.................................................... 8i ELECTROMAGNETIC COMPATIBILITY PaOGRAM .. ............. &I DESIGN CONSIDERATIONS..................................... 8-2 ENVIRO1NMENTAM. ASPECTS.................................... 8-2 COMMUNICATION EQUIPMENT ........................ ......... 8-3 GENERAL ..................................................... 8-3 MICRt)PIONE-HEADSET ....................................... 8-4 INTERCOMMUNICATION SELECTOR BOX......................-., NAVIGATIONAL EQUIPMENT;................................... .8-4 GENERAL ..................................................... b-4 TERMINAL MANEUVERING EQUIPMENT ....................... 8-5 EN ROUTE NAVIGATION EQUIPMENT........................ h5 Automatic Direction Finder (A.DF) ................... ...... ....... 8-5 Distance-measuring Equipment (DMW) ................ I............. 8-5 Tactical Air Navigation (TACAN).................................. 845 Lona-ranize Navistation (LORAN).................................. t-6 Compasses ..................................................... 8-6L Doppler Navigation Systems ....................................... 8-6 Inertial Navigation Systems ................................. ...... -E; 6 INTERDICTION EQUIPMENT ................................... 8_/ 8-7 1.0W-LIGHT-LEVEL NAVIGATIONAL EQUIPMENT ............... STATION-KEEPING EQUIPMENT .............. ............... 9-7 FIRE CO1ý'TROL EQUIPMENT ......................... .......... 8GENEt. AL................................. .................... 8-7 INSTALLATION .................................. .. ........... 8-8 SIGHTING STATION ............................................ 8-8 SENSORS ...................................................... 8-8 ..

k

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T~rvDW

O At C'~k

8-4.6 8-4.6. 1 8-4.6.2 8-4.6.3 8-4.7 8-5 8-5.1 8-5.2 8-5.3 8-5.3.1 8-5.3.2 8-5.3.3 8-5.3.4 8-5.3.5 8-5.3.6

IX

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. . .

FIRE CONTROL ACCURACY................ I.................... Inertial Stabilization.............................................. Fire Control Datum Planc ..... ................................... Harmonization .................................................. COMPONENT LOCATION......................... .............. ANTENNAS ...................................................... GENERAL ............................. ..... .............. .. ANTENNA DEVELOPMENT..................... ................ LOCATION AND INSTALLATION OF ANI ENNAS ......... Communication Antenna Considierations ............. .... Low Frequency (-1) ........................ .............. ...... H igh Frequency (H F) ý...... .................................... Very High Fre~quency (VHF).......... .................... ... .... Ultra High Ficqucncy (UHF)........... ..... ............ ........ SpeCi3l Purpose ................................................. REFERENCES........................................... .........

r

I

)Lk

8-9 8-9 8-9 8-10 8-10 8-10 U-11 912 f 3 l-l3 : 8-14 8-14 8-;4 g9.14

tA

[_,0

TABLE OF CONTENTS (Continued) ParRAUapC *LISTOF

N

.-

9-2.1 9-2.1.1 -9-2.1.2 9-2.2 9-2.2.1 9-2.2.2 9-2.2.3 9-2.2.4 9-2.2.5 9-2.3 92.. 9-2.ý.2 9-2.3.3nu 9-2.4 9-2.5 9-2.6 9-2.6.1 9.2.6.1.1 9-2.6.1.2

I9-2.6.1.3 9-2-6.1.4 ý92.6.2 '92.. 92.6.4. V-ZAA.1

4\

AND:NEUATI

SUBSYSTEMS DESIGN

SYMBOLS ...............................................

9-1

FLIGHT CONTROL POWER SYSTEMS.......................... Central Hydraui2c Sysiemn......................................... Flight Control Subsystems........................................ UTILITY HYDRAULIC SYSTEMS ................................ Engine-starting Subsystems........................................ Cargo Door and Ramp System ..................................... Cargo and Pnrsoniicl Hoist ........................................ Rotor Brake ..................................... .............. Wheel Brakes................................................... H-YDRAULIC SYSTEM RELIABILITY ......................... ... Flig ht Control Redundancy........................................9Utility System Redundancy........................................ R elinhilitv Asnacte ....................... 41 YDRAULIC SYSTEM STfRENGTH- CONSIDERATIONS............ H-YDRAULIC SYSTEM TEMPERATURE CONSIDERATIONS .... HYDRAULIC SYSTEM DESIGN ......................... ........ Sun'vability. Reliability, and Safety Trade-offt........................ Reservoir Leve! Sensing ................ .........................

91 9-1 .9-2 9-2 9-2 9-3 9-4 9-4 9-59-S

Systemt Switching Concepis

..........

..

-- 8

......

Return Pressure Sensing ....................................... .. Switching and Return Pressure Sensing..................... ...... Operating Pressure Considerations.................................. Selection of Fluid Medium ............... ......................... Filtration (Contamination)........................................ Fluiud

Fiiirutiun. ..............................................

9-5 9-5 9-6 9-6 9-7 9-7 9-8

.

1.. K 1

9-9 9-9 9-10 9-10 9-104

9-!0

9-Il 9-2.6.4.2Ground Operation Filtration ..................................... 9-2 6.4.3 Filtration Lcvel ................................................ 9-11 9.2.6.4.4 iex~rnal Contamination................... ...................... 9-11 9-2.6.5 Fittings ....................................................... 9-1 i 9-2.6.6 Dynamic Fluid Connections.................................. ..... 9-12 9-2.6.7 Peak Power Leve~ls............................................. 9 9-2.6.8 A1PU ant' Engine Starting ................................... ... . 9-2.6.9 Syskem I-eat Rejection Characteristics ................. ........ ... 9-13 9-2.6. 10 System Analysis........................... ...................... 9-13 9-2.7 HYDRAULIC COMPONENT DESIGN AND SEL ECTIO ... ......... ý~ 9_1I ............. Actuators ... L...................................... 9-2.7.1 9-2.7.1.1 Rip-stop Protection ............................................. 9-0 9-?.7. i.2Endurance Testing Requirements.................................. 9.4g 9-2.7.1.3 Seal Alternatives ....................... ........ ............. 9 92.7.1.4 Materials and Stres!; Corusidci tions ................ ............. .. 9-16 9-2.7.1.5 General Requirements...................................... ..... 9- 1f 9-2.7.2 Hydraulic Pumps;.............................. .. ... ............ ~* ~ k9-2.7.3 Accumulators................................................. 92.. Reservoirs ............ L................... ....... ........ ..... 4-20 ....... .. ........ ... e-2c 9.2L-7.5 Pressure Relicf............................ Nx

4

V

'

AMCP 706-202

,T".I Tmo0 (ONvl I-"N'IS ,

qmtinued u

Paragraph 9-2.7.6 9-2.7.7

"9-2.7.8 9-2.7.9

9-2.7.11 :9-2.7.13

9-2.7.12 9. .7. ,z

9-2.7.14 9-2.8 9-2.8.1 9-2.8.2

"9-2.8.3 9-2.8.4

9-2.8.5 9-2.9 9-2.9.1 9-2.9.2 9-2.9.3

"9-2.9.4

,9-2*.

:

9-20 921 9-21 922 9 .22

Control Sclttor Valves ......................... ................. R estrictors ............ ............................................ S~parauc Secvos a a c e v , ............................. ............ .............

9-22 9-24 9-25;

... ..

9 -25.

AIlovable External Leakage ........................................... HYDRAULIC SYSTEM INSTAILLATON ........................... U se of Hoses and Sw ivels .............................................. Ma-ntrtcance Access ............................................ Hard Vetsus Soft !h.s,stllativns ......................................... Component M ou zingnConctnts .......................................

9-25 9-25 9-25 9-25 9-26 9-26

Miscellaneous Instahiation Conside-ati 1 ... ......................... MISCELLANEOUS DESIGN CRI IERIA ............................. Actuators and Assoziated Equipment Design ...... ................... Brake Design .......................................................... Control System Design ................................................ Electrical Design ................................. ....... ............ , ..........................................................

9-27 9.27 9.27 9-28 9-29 9-29

9-2.9.6

Fittings Design . ......................................................

9-;2c).D 9-30

9-2.9.7 9-2.9.8 9-2.9. 9-2.9.10 9.2.9.11

G age and Indicator Design ............. ............................... H osP D esign .......................................................... s .......................................................... R eservoir Design ...................................................... Valve D esign ..........................................................

9-30 9-30 9 -30 9-31 9-31

9-2.9.12

Lubrication ........................................................... PNEUMATIC SYSTEMS.....G ..................................... PNEUMATIC SYSTEM DESIGN...................................... System A nalysis. ... .... ...................... .....................

9-32 9-32 9-32 9-32

"9-3 9-3.1 9-3.1.1

9-3.1.2

System Redundancy

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

9-33

"9-3.2.1

COMPONENT DESIGN ......................................... A ir C om pressors ............ ........................................ Positive D isplacem ent ................................................ Dynamic Displacement ..........................................

9-33 9-33 9-333 9-34

9-3.2.2 9-3.2.3 9-3.2.4

Compressed Air Supply System Selection and Operation ................ Moisture Separators ................................................... Dehydrators ..........................................................

9-34 9-34 9-35

9-3.2.5 9-3.2.6 9-3.2.6.1

F ilters ................................................................ Valves ...................................................... Check Valves ................................................

9-35 9-36 9-36

9-3.2 9-3.2.1.1 9-3.2.1.2P

9-3.2.6.2 9.3.2.6.3 -9-3.2.6.4 9-3.2.6.5 v -3.2.7

,.'.., •~

Pressure Regulation ................................................. Filters .................... .................................. C heck Valves ....................... .......................... ....... Pressure Switches .............................................. PressureTr~nsm !92.7.10 itters ............................ ................ ...

R elief Valves ......................................................... Pressure-reducing Valves .................................... Pressure R egulators .................................................. Directional Control Valves ................................... u ........................................................

9-3.2.8

A ir Storage Bottles ....................................................

9-3.2.9 9-3.2.9.I

Subsystem Com ponerts ............................................... Actuators ............................................................

9-37 9-38 9-38 9-38 9-41 9-42

9-43 9-43

" .

Li

AM0CP706-202 "IAB1.4 01 ()%'

IN' I

SI ( ontinucd)i

Paragraph 9-3.2 9.2

9-3.2.9.3 9,3.2.9.4 9-3.3 9-3.4 9-3.4 .1 9-3.4.2 9-3.4.3 9-3.4.4 9-3.4.5 9-3.4.6

.

10-1 10-2 10-2.1 10-2.2 10-2.3 !0-2.4

-

10-2.4.1 10-2.4.2 10-2.4.3

10-3 10-3.1 10-3.2 10-3.3 10-3.4 10-3.5 10-3.6 10-4 i-4.i

10-4.2 10-4.3 IO.5 10-5.1 10-5.2 10-6 10-6.1 10-6.2 10-6.2.1 10-6.2.2 10-6.2.3 10-6.3 10-7 10-7.1 )10-7.2

10-7.3

Page Brake V alves ..............

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

9-44

Pneum atic Fu ,cs ..................................................... Q uick-disconnect. ............... ............................. ...... PNEUMATIC SYSTEM INSTALLATION AND QUALIFICATION ... PITOT-STATIC SUBSYSTEM DESIGN ............................... A ltimr et rs ................................... ........................ Rate-of-clim b Indicator ............................................... A irspccd Indicators .................................................... Total-pressure Sources ................................................. Static Pressure Sources ................................................ Pitot-static T ubes ...................................................... R EFE R EN C ES ..........................................................

9-44 9-44 9-44 945 9-4 5 9-46 9-46 9-47 9-47 9-4 8 9-48

CIIAPTER 10 INSTR I EMiNTATION SIUBSYSTEM I)ESIGN IN T RO D U C TIO N ............................. ........................ INSTRUMENTATION LIGHTING REQUIREMENTS ................. G EN E R A L ............................................................ LIGHTING INTENSITY CONTROL .................................. LOW INTENSITY READABILITY ....................................

10-1 10-1 10-1 10-1 10-2

WARNINGC. CAUTION, AND ADVISOR Y S!ONALS

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

W arning Signals ....................................................... C aution Signals ... ................................. .................. A dvisory Lights .......................................................

10-2 10-2 10-3

FLIG HT INSTRUM ENTS . ............................................. ( EN ER A L ............... ............................................ AIRSPEED INDICATORS ......................... .................. A LT IM ET ER S ........................................................ TURN-AND-BANK INDICATORS ................................... ATTITUDE INDICATOR ............................................. RATE-OF-CLIMB INDICATORS ..................................... NAVIGATIONAL INSTRUMENTATION .............................. ENEAL...........................................................

1.7

10-3 10-3 10-3 10-3 10-3 10-3 10-5 10-5

....

TYPESOF INSTRUMENTS ........................................... M A P DISPLA YS ...................................................... HELICOPTER SUBSYSTEM INSTRUMENTATION .................... G EN E R A L ................... ............................ ........... INSTRUMENTATION REQUIRED .................................. WEAPON SYSTEM INSTRUMENTATION ............................. G EN ER A L ............................................................ DESIGN REQUIREMENTS ........................................... Arming, Fuzing, and Suspension and Release Control Design ............ Human Fa -tors Considerations ......................................... Indicator Dcsign ...................................................... WEAPON SELECTION CONTROLLLR/PROGRAMMER ........... TYPES Of INSTRUMENT .............................................. IN STA LLATIO N ...................................................... V IBR A T IO N .......................................................... ACCESSIBILITY AND MAINTENANCE ............................. R EFER EN C ES ............................................ .............

10-5 10-7 10-7 10-7 10-7 10-7 10-7 10-8 10-" 10-8 10-8 10-9 10-9 10-10 10-10 10-10 10- 10 xvii

TABLE OF CONTENTS (-]onitnuvdi Paragiaph

I -f0 I-oI 11-2 11-2.1 11-2.2 1a-2.3 11-2.4 11-2.5 11-2.6 11-2.7 11-3 11-3.1 11-3.2 11-3.3 11-3.4 11-3.5 11-3.6 i i-3.7 11-4 11-4.1 11-4.2 I1-5 11-6 11-7 I-8 11-8.1 I1-8.2

12-0 12-1 12-1.1 12-1.1.1 12-1.1.2 12-1.1.2 .1 12-1.1.2 .2 12-1.1.2.3 12-1.1.2.4 12-12 12-1.2 ,1 12-1.2.2 12-1.2.3 12-1.3 12-1.3.. 12-1.3.2 12-1 3.3

"xViii

Page CHAPTER i i AIRFRAME STRUCTURAL DESIGN LIST O F SYM BO LS ..................................................... INTRO DUCTIO N ....................................................... I)ESIGN CONSIDERATIONS ........................................... W EIG H T .............................................................. SURFACE SMOOTHNESS ............................................ STIFFNESS AND RUGGEDNESS .................................... FATIGUE SENSITIVITY ............................................. C O ST .................................................................. M A T ER IA LS .......................................................... SU RVIVA BILITY .................................................... DESIGN AND CONSTRUCTION ....................................... FITTINGS .................................................. SU PPO RT S ............................................................ FRA M ES ............................................................. BU LK H EA DS ......................................................... SKIN SUBSYSTEM S .................................................. CORROSION PROTECTION ........................................ FiEC.TRWiAL BLONDiNG ................................................ CARGO COMPARTMENT .............................................. STA T IC LO A D S ....................................................... C RA SH LO A D S ...................... ................................ TRANSPARENT AREAS ................................................ D EV ELO PM EN T ............ ........ .................................. M A N U FACTU R E ................. .......................... .......... SU BSTA N TIATIO N ..................................................... A N A L Y SIS .............................................................. T EST IN G ................................................................ R EF E R EN C ES .......................................................... CHAPTER 12 LANDING GEAR SUJBSYSTEM LISTO SYM BO LS ..................................................... G EA R T Y PE S .......................................................... W H EELG EA R........................................................ G enera .......... ...... ................ ............................. Comrponent Design and Selection .................................... Tires .................. .............. .............................. W hec ls .............................................................. Shock Struts ................... ................................. B rakes ............................................................... SK ID G EA R .......................................................... G enera l ............................................................... G round-handling W he ............................ ................. Scuff P late% ....... ................................................. RETRACTABLEG EAR .................... .......................... G eneral ............................................................... A ctuatio n ............................................................. Em ergency Extension ..................................................

II-I I1-1 Il-1 11-1 I1-! l111-2 I1-2 11-2 11-2 11-4 I1-5 i-5 1l-5 11-6 11-6 11-6 .11-7

11-7 11-7 11-9 li-Il 11-12 11-12 11-12 1 1-13 1 1-13 11-13

12-I 12-1 12-1 12-1 12-3 12-3 12-4 12-5 12-8 12-8 12-8 12-8 12-9 12-9 12-9 12-9 12-9

AMI;P 706-202 TABLE OF ('ONTFNTS (('ouilnued)

Paragraph

Page

I2- 1.4 12-1.4.1 12-1.4.2 12-2 12-3 12-4 12-4.1 124 .2 12-4.3 12-4.4 124.5

R EFER EN C ES .'. ..................................

12-14

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

13-0 13-1

13-2 13-2.1.1

PERSONNEL ACCOMMODATIONS ...... ......................... General Vision Requirennents .................................... .1

13-2.1.2.2 13-2.1.3 13-2.1.3.1 13-2.1.3.2 13-2.1.4 13-2.2 13-2.2.1 13-2.2.2 13-2.2.3 13-2.2.4 13-2.2.5 13-2.2.6 13-2.3 13-2.3.1 13-2.3.2 13-2.3.3 13-2.3 4 13-2.3.5 13-2,3.6 13-2.3.7 13-2.4 13-2.4. I 13-2.4.2 13-2.4.3

',

12-9 12-9 12-10 12-11 12-11 12.12 12-12 12-12 12-12 12-13 12-14

CHAPTER 13 CREW STATIONS AND CARGO PROVISIONS LIST OF SYM BOLS .............. t ................................ INTRODUCTION ...................... .............................

13-2.1.2 13-2.1.2.1

)

SKIS AND BEAR PAW S ............................................. G eneral ..................................................... . ... . . Installatior .................................................. LANDING LOAD ANALYSIS AVOIDANCE O!"GROUND RESONANCE ............................. WATER-LANDING CAPABILITY ...................................... G EN ER A L ............. .............................................. PRIM E CAPABILITY ................................................. ADDITIONAL CAPABILITY ........................................ EMERGENCY FLOTATION CAPABILITY ........................... M O D ELTESTS .......................................................

13-2.5 13-2.5.1 13-2...2

i Ik,•.,% F

•II .

..

..

.

..

..

-a

•

¢...

. ..

. ..

. ..

. . . ..

.

..

.

. . ..

Controls .. Pitch C ontrols .....................................

. . ..

.

. . .

. ..

.

..

. .

.

13-I 13-1 ..

.

13-1 132 3-2

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

Directional Control Pedals ........................................... Seats. Belts. and H arnesses ............................................. C rew Seats ........................................................... Belts and H arnesses ................... ............ ................. M ap and D ata C ases .................................................. PASSENGER COMPARTMENT ..................................... Troop and Passcngcr Seats ............................................ C o lo r ............................................... ................. U pholstering and Carpeting ........................................... Sm oking Provisions ....................................... ........... Signal Lights and A larm Bells .......................................... Acrom edical Evacuation .............................................. SURVIVAL FQUIPMENT ...................................... .. Inflight Escape and Survival Equipment ............................... Ground Escape and Ditching Provisions ............................. Em ergency L.ighting Provisions ........................................ L ife R afts ............ ................................................ Survival K its ......................................................... First Aid ................. .................................. Fire Extinguishing Svstern s and Axe .................................... ENVIRONM ENTAi CONTROL ..................................... Ventilation. Heating. and Cooling .... ................................. Windshield Defogging and Deicing Equipncnti....................... .. A coustical Environm ent ............................... ............... SIGHITS AND SIGHTING STATIONS ............................... D irect-view ing Sights ................... .............................. H elm et M ounted Sight ................................................

13-2 13-2 I -3 13-3 13-3 13-4 13-5 13-5 13-5 13-5 13-5 13-5 13-5 13-5 13-6 13-6 13-7 13-7 13-7 13-7 13-7 13-7 13-7 13-7 13-8 13-8 13-8 13-8 13-9

.TABI .01O

(11"VI'Tl'

S I (.'i1niinu.d)

Page

Parjoraph 13-2.5.3 13-2.5.4 13-3 13-3.1 13-3.1.1 13-3.1.2 13-3.1.3 13-3.1.4 13-3.1.5 13-3.1.6 13-3.2 13-3.2.1 13-3.2.2 13-3.2.2.1 13-3.2.2.2 13-3.2.3 13-3.2.4 13-3.2.5 13-3.2.6 13-3.2.7 .3-37.8

Cabin and Compartment Lighting ...................................... ................................ C ockpit Lighting ..................... Utility Lights ............................................... ..................................... Secondary Lighting ............ Panel Lightii• ........................................................ Interior Emergcncy ! ;oht . ....... .................................... Portable Inspection Lights ............................................. Troop Jum p Signal Light .............................................. Warning, Caution, and Advisory Lights ................................ |Inmtrm ent lanel lighting ........................................

13-3.2.9

Cargo Comnartment Lighting .......................................

13-4 13-4.1 13-4.1.1 13-4 .i.2 134 .1.3 13-4.2 13-4.2.1 13-4.2.2

S13-4.2.3 13-4.2.4

13-9 13-9 13-9 13-9 13-9 13-9 13-10 13-10

Indirect Sights ..... ................................................... M is~ile Sighting Stations ............................................... LIG HTING SYSTEM S .................................................. ................................... EXTERIOR LIGHTING SYSTEM A nticollision Light System ............................................. Form ation Lights ..................................................... Landing/Taxi Light ......................................... ........ Searchlight ........................................... ................ ............................................ Floodlight System Position Lights ........................................................ INTERIOR LIGHTING SYSTEM .....................................

3-10 13-10 13-10

.

....

CARGO PROVISIONS .................................................. .................... INTERNALCARGO ............................. Cargo Compartment Layout ........................................... Detail Design ......................................................... Loading A ids ......................................................... EXTERNAL CARGO ................................................. Static Loads ........................................................... ....................................... D ynam ic Loads ............... W inches and H ooks ................................................... .......... ......................... .. System Safety ............

13-10 13-10 13.10 13-10 13.10 13-10 13-10 13-I1 13-I1 13-11

13--il 13-11 13-11 13-11 13-11 13-13 13-14 13-18 13-18 13-19 13-20

t.

.--.......................... .....

CHAPTER 14 14-0 14-1 14-2 14-2.1 14 -2.1.1 14-2.1.2 14-2. 1.2.1 14-2.1.2.2

"14-2.1.2.3 14-2.1.2.4 14-2.1.2.5 14-2.1.2.6 xx

ARMOR. ARMAMENT. AND PROTE('TIVE SUIBSYSTEMS l)E);l(;N .............................. LIST O F SY M BO LS ..................... IN TRO D U CTIO N ....................................................... ..................... ARM IAM ENT SYSTEM S .......................... G U N S ................................................................. T ypes ................................................................. Location ...... ....................................................... Projectile Flight Path ................................................. Blast Effecis ........................................................ Debiis Ejection Path ................................................. External G un Jettisoning ............................................. A ccessibility ......................................................... Dynam ic Forces ......................................................

14-1 14-1 14-1 14-1 14-1 14-2 14.3 14-3 14-3 14-3 14-3 14-4

€

1iil.f f

(OF'O:NT

'ontinucd i NT( l Page

Paragraph 14-:.1.3 14 2.1.3.1 14-2.1.3 2 14-2.1.3.3 14-2.1.4 W14 2.1.5 14-2.1.6 14-2.2 14-2.2.1 14-2.2.2 14.2.2.3 14-2.2.4 14-2.2.5 14-2.2.6 14-2.2.7

Typcs of Installation; .................................................. Pod Installations ..................................................... Turret Installations ................................................... Pintle G uns .......................................................... A m m unition Storage .................................................. A m m unition Feed ..................................................... Boresighting and Harmonization ....................................... G UIDED M ISSILES ................................................... Location of Launcher Installations ..................................... Structural Clearance ................................................... Blast Protection ....................................................... A ccessibility ................ ......................................... Firing C ircuit Testing ............................... .................. Jettisoning ................................................. Effects of Aircraft M aneuvers ..........................................

14-4 14-4 14-4 14-5 14-5 14-6 14-6 14-6 14-6 14-7 14-7 14-7 14-7 14-7 14-7

14-2.2.8

Types of Installations ..................................................

14-7

14-2.2.9 14-2.2. iO 14-2.2.11

L oading .............................................................. Aerodynamic Effccts .................................................. Suspension and Retention .............................................

14-7 14-7 14-7

14-2.2.13 14-2.2.14 14-2.3 14-2.3.1 14-2.3.2 14-2.3.3

Restraining Latch ..................................................... Forced Ejection ....................................................... R O C K ETS ............................................................ Rocket Launcher Installations ......................................... Launch Tube M aterials ................................................ Launcher M ounting ...................................................

14-8 14-8 14-8 14-8 14-9 14-9

14-Z.2.i2

Launch initia.ion

"14-2.3.4 14-2.3.5 14-2. 3.6

.....

14-9

Load Requirem ents ................................................... G round Safety ........................................................

14-9 14-9

Firing Contacts ....................................................... Intervalonicter ........................................................ Launcher Fairing ............................................... ......

14-2.4.1 14-2.4.2 14-2.4.3 14-2.4.4 14-2.4.5 14-2.4.6 14-2.4.7 14-2.4.8 14-3 14-3.1 14-3.2 14-3.2.1 14-3.2.2 14-3.2.3 14-3.2.4

"14-3.2.5

14-9 1i4-i0 14-10

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

14-10

Safety C riteria ......................................................... Fire Interrupters ................... .................................. Contour Followers ............................................... Burst Lim iters ......................................................... C ockpit N oise ................................. ....................... Debris D isposal ................... ................................... Toxic Explosive Gas Prot,:ction ........................................ Turret M as,¢r Power Switch ........................................... PROTECTIVE SUBSYSTEMS ........................................... G EN ER A L ............................................................ DEVELOPMENT OF VULNERABILITY REDUCTION SYSTEMS.. Vulnerability Analysis ................................................ Vulnerability Reduction Checklist ................................... Vulnerability Data Presentation ....................................... Aircrew Armor Configuration Development ........................... Armor M aterial Selection ............. ................................

14-10 14-10 14-1 14-I I 14-I 1 14-I I 14-12 14-12 1412 14-1 2 14-13 14-13 14-16 14-16 14-16 14-16

SAFETY CONSIDERATIONS

'

A

4-4

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

Restraining Latches

14-2.3.8 14-2.3.9 14-2.3.10

"

4-8

N umber of Rockets ....................................................

14-2.3.7

14-2.4

)

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

[

.

xxi

K t.

K

'-'.

.. . . ...

ANCP 706-PM "TABLE

OF CONTENTS (('onlinucd) Page

Paragraph 14-3.3 14-3.3.1 14-3.3.2 14-3.3.3 14-3.3.4 14-3.3.5 14-3.3.6 14-3.36.1 14-3.3.6.2 14-3.3.6.3

15-1 15-2 05.2.1 15-2.2 15-2.3 15-,.4 15 2.5 15-3

ARMOR INSTALLATION DESIGN CONSIDERATIONS ............ A irc, ew Torso A rm or ................................................. Interchangcability ..................................................... .................................. R em ovability ...................... F ,ying Qualities ....................................................... Im m obilization ........................................................ Armor Material Attachmcnt/Installation ............................... M ounting of Arm or Plate ............................................. Installation Design ............................. ..................... Bullet Splash and Spall ........... ................................... R EFER EN C ES ..........................................................

14-18 14-18 14-18 14-18 14-19 14-19 14-19 14-19 14-19 14-20 14-20

('HAPTER 15 MAINTENANCE AND GROUND SUPPORT EQIIPMENT ((;SF) INrFRFA('F IN T RO D U C T IO N ....................................................... DESIGN CONSIDERATIONS AND REQUIREMENTS ................. SA F ET Y .............................................................. A CC FSSIBILIT Y ...................................................... STANDARD IZATION ................................................ HUM AN ENG INEERING ............................................ INSPECTION, TEST, AND DIAGNOSTIC SYSTEM .................. PROPULSION SUBSYSTEM INTERFACES ............................

15-1 15-1 15-1 15-2 15-2 15-3 5-3 15-3

15-3.1

G E N E R A L ............................................................

15-3

15-3.2 15-3.3 15-3.4

INTERCHANG EA BILITY/QUICK-CHANGE ........................ CONNECTORS AND DISCONNECT POINTS ...................... INSPECTION AND TEST POINTS ....................................

15-4 15-4 15-4

OIL, FUEL, AND LUBRICATICN ................................... G RO U N D IN G ........................................................ STARTING ................................. S A R IN .......................... 15 3 . ............................. GROUND HEATERS .................. 53.8 EN G IN E W A SH ....................................................... 15-3.9 15-4 TRANSMISSIONS AND DRIVES ....................................... 15-5 ROTORS AND PROPELLERS ........................................ 15-6 FLIG HT CONTRO LS ................................................... ......................... 15-6.1 ROTATING SYSTEM S ...................... 15-6.2 NONROTATING SYSTEMS .......................................... T R IM SY ST EM S ...... . .............................................. 15-6.3 ELECTRICAL SUBSYSTEMS ......................................... 15-7 A VIO N IC SU BSYSTEM S ................................................ 15-8 15-8.1 COMMUNICATION SYSTEMS .................................... 15-8.2 NAVIG ATION SYSTEM S .. ... ...................................... 15-9 HYDRAULIC AND PNEUMATIC SUBSYSTLMS ..................... 15-9.1 HYDRAULICSUBSYSTEM ...................................... PNEUMATIC SUBSYSTEM ................................... 15-9.2 15-10 INSTRUMENTATION SUBSYSI EMS .................................. ............... 15-10.1 FLIGHT INSTRUMENTS ...................... NAVIGATION INSTRUMENTS ...................................... 15-10.2 AERIAL VEHICLE SUBSYSTI M INS1 RUMt-N]ATION ............ 15-10.3 ........... ................. A IRFRA M E STR UCTU RE ................ 15-11 LANDING GEAR SUBSYSTIM ..................................... 15-12 15-3.5 15-3.6 15-3,7

xxii

.

15-4 15-4 15-4 15-4 154 15-5 15-5 15-5 15-5 15-5 15-5 15-5 15-6 15-6 15-6 15-6 15-6 15-7 15-7 15-7 15-7 15-7 15-7 15-7 15X

L)

"

AMCP 7CO-202 TABLEl- ('ONTENINS C(ontinued) Paragraph

S"

Page

15-13

C R EW STATIO N S ...................

"15-14

ARMAMEN

AND PROTECTIVE SYSTEMS ................

15-8

16-0 16-1 16-2 16-2.1 16-2.2 16-2.2.1 16-2.2.2 16-2.2.3 16-2.2.4

STANDARD PARTI LIST O F SY M BO LS ..................................................... IN TRO D U CTIO N ........................................ r .............. FA ST FN E IRS .................... ....................................... GENERAL .................................................. THREADED FASTENERS ...................................... Screw s ............................................................... B o lts ............................. ............................. ...... N uts .. ...................... ......................................... W ashers ........................ ... .............. ..................

16-1 16-1 16-1 16-I 16-1 16-1 16-2 16-2 16-2

:

16-2.3 16-2 .3.1 16-2.3.2 16-2.3.3 16-2.3.4 16-2.3.6 16-2.3.7 -. 16-3 16-3.1 16-3.2 16-3.2.1 16-3.2.2 16-3.2.3 16-3.3 16-3.3.1 16-3.3.2 16-3.3.3 16-3 3.4 16-3.4 16-3.5 16-3.6 16-3.7 16 -3.7 .1 16-3.7.2 16-4

16-4.1 16-4.2.

r, ARMOR,

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

('lAPTFR 1I)

16-4.2.2 16-4.2.2.1

16-4.2.2.2 16-4.2.2.3

.

NONTltREADED FASTLNERS ................................... R ivets ......................... ...................................... Pins. ...................................................... Quick-release Fasteners ............. ............................... .............. .. Turnbuckiics and Tcfinitaiil. .... Retaining Rings ........ ................................. Clam ps and G rom mets ................................................ Self-retaining Fasteners ................................................ HEARINGS............. ........................ ......I............ G E N ERA l . ................ ........................................... BA LL BEA RIN G S ..................................................... R adial Ball Bearings ................................................... A ngular Contact Bearings ............................................. Thrust Bali Bearings ... ................... ........................ ROLLER BEARINGS ................................................ C ylindrical Roller Bearings ............................................ N eedle Bearing% ............... ................ ........... ........ Spherical Roller Bearings ............................................. Tapered R oller Bearings ................ .............................. A IRFRAM E BEARIN G S .............................................. SLIDING BEARINGS ......................................... LAMINATED EL.ASTOMiRIC BI-ARINYS .......................... BEARING SEALS AND RETAINERS .............................

16-3 16 -3 K.3 16-3 16.3 b16-4 164 16-4 16-4 16-b 16-8 16-9 16-10 16-10 16-10

S eals ....... .. . . .............................. ............... Bearing R tcnrtion ..................................................... ELECTRICAL FITTINGS ...................

16 -15 16-15 16-16

16-11 16-11 16-12 16-12 16-14 16-15

G E N E R A L ....................... ... ..... .......................... CONNECTORS AND CABLE ADAPTERS ............................ Connector Selection .N... A ............................. C ircular C opnectors ................. .......... .....................

16-4.2.1

16-16 16-16 16-16

16-18

Termination Seals ............................... C able A dapters ......................................................

16-19 16-19

Connector Couplings ...............................................

16-19

16-4.2.3 16-4.2.4

Rack and Panel Connector . ...................................... Flat Conductor Cable Connector ....................................

16-4.2.5

Printed W iring Board Connector .r ................. ....................

-•--:

15-8

.

16-19 16-19

16-19

"x

xiii

Si.......

'•

I

AM CP 706-202

\EI

(%I

"l.,AHI.I. OI1 ( ()\I i '\1 SI( I ,,niiguud

! Paragraphl 16-4 .3 16-4.4 16-5 16-5.1 16-5.1.1 16-5.1.2 16-5.1.3 16-6 16.6.1 16-6.2 16-6.2.1 16-6.2.2 16-6.2.3 16-6.2.4 16-6.2.5 16-6 2.6 16-6.2.7 16-7 16-7.1 I 16.L 16-7.2.1 16-7.2 2 16-7.2.3 16-7.2.4 16-7.2.5 16-7.3 16-7.4 16-8 16-8.1 16-8.2 16-8.2.1 16-8.2.2 16-8.2.3 16-8 .2 .4 16-8.2.5 16-8.3 16-8.3.1 16-8.3.2 16-8.3.3 16"9 16-9.1 16-9.2 16-9.3 16-9.4 16-9.4.1 16-9.4.2 16-9.4.3 16-9.4.4 16-9.4.5 16-9.5

xxiv

T E R M IN Al S ... ........... ................................. ........ TERMINAl BOARDS ........................................... ELECTRICAL SW ITCHtES .............................................. ........... ........ ................... G EN I'R AI ................. ...... .............................................. T oggle S% itclics Pu.,h-button SAitches ...................................... ... . ..... ......... ....... R otaw S's itchc% ..... ............................... PIPE ANI) TUBE !ITTINc;S ......................................... GENERAl ..... ............................................. TYPES 1F: lTT NGS . . ....................................... . T apered Pipe I hreads ............ .................................... Stritght T hread Fittings ............................................... Flared T ube Fitlings .................................................. Flareless T ube Fitting% ............ .................................... Thin Wall Tube Connectors ........................................ Q uick-disconnect C oupin , ........................................... ..................... Perm anent F ittings .............................. C O N T ROL. PU LL EY S ........................ .......................... G E N E R A L ............................................................ ILLEY , -............................................................ SE.LECTIO P ulle%D iam eter ........................ ............................. Pulle) G roove ... ...... ....... ....................................... P ulle) S treng th .......................................... ............. Pulley Perform ance .................................................... ..... N onm etallic Pule) s .... ........................................ PU IL EY IN STAI LATIO N ......... .................................. PU LLEY G U A R D S ................................................... PUSH-PULL CONTROl S ANI) [ItiXiBII. SIAI"TS ................... G E N E R A L ............................................................ ......................... PUSH-PULL CONTROLS ................... ................. C ontrol T ravel ...................................... ................................ C ontrol L oads ....................... ........................... Core Configurations ................... C o n d ui. ............................................................... End F:itting, ................................................ FLEXIBLE SHAFTS ........................................... Torque C apacity ..................................................... .................... Flexible Power Shafts . .......................... Flexible Control Shafts .. ......................................... CABLES AND WIRES(STRUCTURAI ) ................................ G E N E R A L ............................................................ PREFORMED WRE STRAND AND CABL I . ....................... TYPES 01: CABLE CONSTRUCTION ............................... C A BLE SELECT IO N .................................................. C able Strength ........... ............................................ ............................................. Cable Deflection Operating Characteristics ....................................... . W ire M aterial ...................................................... C ablc(.onstruction ............. ....................... ...... ....... SAFETY WIRE AND COTTIER PINS .................................. R EFER EN C ES .........................................................

6 19 ,16-19 6-20 .I%-20 16 -20 l 20 16 -22 1622 122 16-22

.

16 22 16-23

16-23 16-24 16-25 16-25

..

.

.

.

.

16 25 l ,-25 1 6-25 ,I 2 16-211 16-26 16-2(1 16-26 .1 -26, 1(,-2(, 16-27 16-27 16-27 16,-27 I t,-2,,s 16 -29 16-28 16-28 16-21) 16-29 16-30 16-30 16-30 (1 31 16-3 1 16-31 16-3 1 16-31 1(-3?. 1(0-3I 16 32 1 -320 16 -32 16-32 16-3.

706-202

_____AAMCP

TABI.E 01 (ON'I EN• S i(',.,.ue _Paragraph

Pape (HAIA

T

llR 17

VROCl*:sSFS 17-I 17-2 17-2.1 17-2.2 17-2.2.1 17-2.2.2

IN T R O I)U C T IO N ..................... . .......... ... ................. M ET A LW O R K IN G .............. ............................. ........ G E N E R A I . . .......................................................... C A ST IN G ...... .............................. ....................... Sand C astings ............. ............ ............................. Investment Castings ......................................... ...... Permanent Mold Castin. ........................................ C entrifugal C astings ................................................... FO R G IN G ............................................................ E XT R U SIO N .................................... ..................... SHEET-METAL FORMING .......................................... M achine Form ing .................. ................................. Shop Fabrication ........ ............................................. M M............ E N IN G ........ .. ................... ................

"17-2.2.3 17-2 2 4 17-2.3 17-2.4 17-2.5 17-2.5.I 17-2.5.2 17-3

17-3.1

GENERAL................. ..................................

17-3.2 17-3.3

MACHINING OPERATIONS ...................................... ELEMENTS OF MACHINING DESIGN .............................. JOINING .......................... GEN ER A L ..........................................................

17-4

17-4.1

9.17-4.2

...

..

WELDING, BRAZING, AND SOLDERING .......................... W elding .............................................................. B razing ................... ........................................... So ldering .. ........................... ... . ........................ MECHANICAL FASTENING ......................................... R ivets ................................................................ Bolts, N uts, and W ashers .............................................. Screw s ................................................................ ADHESIVE BONDING - STRUCTURAL .......................... SWAGING AND CABLE SPI ICING .................................. IIEAT TREATM ENT ................... ............................. G EN ERAL ............................. .......... ................ HEATTFREATMtENT METALLURGY ....... .......... .. .. A nnealing ............................................. N orm alizing .. ........................................... ............ Stress Relief ........ Tem pering ............................................................ A ging ........................................ ........................ FERRO U S ALLOYS ......... ............................ ........... NONFERROUS ALLOYS ............................................. A lum inum A lloys ..................................................... C opper A lloys ......................... ............................... T itanium A iloys ................ 0 .. ................................... DESIGN ASPECTS OF HEAT TREATING ............................ WORK HARDENING .................................................. G E N ER A L ............................................................ FO R M IN G ...................... ..................................... ROLLER BURNISHING .............................................. SHOT-PEEN ING .....................................................

17-4 .2.1 17-4 .2.2 17-4 .2.3 17-4.3 17-4 .3.1 17-4-3.2 17-4 .3.3 17-4.4 17.4.5 17-5 17-5.1 17-5.2 17-5.2.1 17-5.2.2 17-5.2.3 17.5.2.4 17-5.2.5 17-5.3 17-5.4 17-5.1.1 17-5.4.2 17-5.4.3 17-5.5 17-6 17-6.1 17-6.2 17-6.3

ST7.6.4 i'

•

17-I 17-I 17-I 17-1 17.2 17-2 17. 17-2 17-2 17-2 17-3 17-3 17-3 17-4

17-4

17-5 17-5 7-6

!7-6

17-7 17-7 17-8 17-10 17-10 17-10 17-1I .7-11 17-12 17-15 17-16 17-16 17-17 17-17 17-17 17-17 17-17 17-17 17-17 17-18 17-18 17-18 17-18 17-19 17-19 17-19 17-19 17-19 17-20 xxv

TABLE OF C17N1 EN"IS (Continued) Page

Piragraph 17.7 .17-7.1 17-7.2 17-7.3 17-7.4

................................. TOOLING ........................ GENERAL ....................................................... SHOP TOOLING.................................................. ......... AIRFRAME TOOLING .................................. T EST TOOLING .................................................. REFERENCES................... ..................................

1720 17-20 17-22 17-22 17-23 17-23

APPENDIX A ]EXAMPLE OF A PRELIMINARY HEATING, COOLING, AND V'ENfILATION ANALY'1S A-1 ...... HEATING AND V~ENTILATION ANALYSIS................... A-1 A-1 DESIGN REQUIREMENTS ........................................ A-1.1 A-1 DESIGN ASSUMPTIONS .......................................... A-1.2 A-I ....... H EAT LOSSES................ ............................ A-1.3 A-1 Cockpit.......................................................... A-1.3.i A-1 ......................... Convection............................ A-1.3.1.1 A-2 Infiltration ................................................... A-l.'.I.2 A-2 Total Cockpit Heat Loss..................................... I...... A-1.3.1.3 A-2 Cabin....................... .................................... A-1.3.2 A-Z A-1.3.2.l Convection ..................................................... AA-1. 3.2.2 Infiltration ........................................................... A-', Tota! Cabini Heat Loss.................................................. A-1.3.2.3 A-2 ........ VENT ILATING AIR REQUIRED ........................... A-I1.4 A-2 A-1.4.1 Based on Number of Occupants and Minimum Ventilating Rate............ A-3 Requirement Based on Maximum Allowable Temperature Difference ... A- 1.4.2 A-3 I............................. Cockpit Requirement................. A- 1.4.2.1 A-3 Cabin Requiremcnt................................................ A-1.4.2.2 A-3 A- 1.4.2.3 Total Air Requirein~eit . .......................................... A-3 I.................................. Total Heat Requirement .......... A- 1.4.3 A-3 HEATER REQUIREMENTS................. ........................ A-1.5 A-3 ........ Heat Gained.............................................. A-1.5.1 A

V

I C~ '%

A-2.3. i

A-2.3.2

Effective A~'s..................................................

A-2.4 I

A-2 4. 1-1 A-2.4.1.2 A-2.4.1.3 A-2.4.1.4 A-2.4.1 xxvi

Iia

................... Heater Size .................................... BLOWER SIZE ................................................... Volume of Air to be Delivered ....................................... Pressure Drop . . . . . . . . . . . . . . . . . . . . . . . . . . . .. COOLING AND VENTILATING ANALYSIS .............. DESIGN REQUIREMENTS ...................... DESIGN ASSUMNiPTIONS ....................... DETERNIINOATION OF EFFECTIVE TFMiPLR,.TURIDIFFERENCLS ASSOCIATED WITH VARIOUS SURFACES OF THlE HELICOPTER .............................................. Effective Solar Tempe~ratures ........................................

A-i.5.3 A-1.6 A. 1.6.1 A-1.6.2 A-? A-2.1 A-2.2 A,-2.1

COCKPIT HEAT GAINS........................................... ionvectiori. lnfiltrat'on, and Solar Radiation........................ Convection Gains................................... ............ Infiltration Gain ...................... ............. ........... Solar Radiation Gajii ................... ........................ Total Hleat Gain Ducto Convection. Infiltrationi and Solar Radiation .. Occupants .......................................................

A~AUI~ IAI

A-4 A-4 A-4 A-4 A-4 A-4 A-4 A-4 A-4

A-5 A-5 A. 5 A-5 A-6 A-6 A-6 A-6

(

TABLE OF CONTENTS I(onfinaed) Paragraph

A-2.4.3 A-2.4.4 A-2.5 A-2.5.1

fAtC

'Electrical System .............. ........... Total Cockpit liea&. Gain .................................... AIR CONDITIONER SIZE ........................................... Conditioning of Ventilation Air ........................................

A-6 x6 A-6 A-6

A -2.5.2

Fan Size and H rit .....................................................

A -7

A-2.5.3

Tons of Refrigeration Required ....................................... -REFEREN CES .......................................................... IND•EX .................................. ..............

A-7 A -7 . J-i

.9

'22

S",,,.XXV8I

ýLIST OF ILLUSThAA 9ýNS it.,$u No,

2

Titie-ct~y

fig. 2-I Sdwc Srcue. ......... ..................... -2 Fig. 2-2 'Weight Compai ison of Matcri,0%ýor Equial 1ttii. ....... .. ......... 2-22 -Fige. 2-3 Comparative Sonic Fatigue Rnsistancý: fltZpnvcjiti~oawA bud S-4ndwichi4 '.Structures . . . . . . ... . . . . . ... . . ... .. 2-22 4Fig. 2-4 -Common Honeycomb Conf;ý urations ............................. .... 123 fig. 25 trperdecs of Balss Wood - Comprcssive Strengtl i,\ rbnsity.............. -24 t%4ig. 2-6 Properties of Balsa Wood - "L" Shear Strength vs Denwy ................ .$24 3Z tig.44 27 Typical Stabilized Comprcsaivc Strength ......... I......... 2-24 2-24 ............. ............................. Strength Typical "1" Shear ~'r. 2-8 vig. 2-9 .Typicrsl "L" Sheur Modulus........................................ (24 IFtg. 2-lb Modes fr(ailu.- of Sandwich Componimc Under Edgcwisc Loads ........... -,-.2-7 ' Fig. 3-; Submer-ged Engine Installation (Exantplc) .............................. .3-2 'wFig. 3-2 Scmicxposcd Engine lnstclladon (Example).............. ............... 3-2 Fir. 3-3 Exposed En-im: Installation (Exiampc)................. .33 KFig. 3-4 Typical Fuel Subsytcmn.................. ... .......... .............. 3-10 fig. 3-5 Typical Fuel Subsystem Wit Pressure Rcfuel~ng.......... ............... 3-12 rig. 3-6 PFerformance Cci rrczions for Duct Lovsses................. ....... .3-17 *fig.3-7 Allowable Combined lvde;ý asnd Exhaust Duct Pressure LOSSeS................ 3-17 fig. 4.1 4Helicopter Main GeiL box Weignt vs Takeoff Power ....................... 4-4 f ig. 4-2 tower Loss to Hcfat vs lnpui Povwrc -- Typical Twin -engine-driven Gearbox 4-5 ""A"............ I .......... ........ 4-7 t.ig. 4-4 Elastic Body Contact Pressurt IlIbstribution and interface Contour ........... 4$Fig.4.5 .Fri,.Iion Coeffhieint vs EI-1I) Parameters - Regions I and 11 ............. 4-8 f'ig.- 4-6 :"Angle Of Enga.;emcnt ............................................... 4-9 _l-ig. 4-7 Coefficient of Friction vs S9iting Velocity ............................... 4-10 S4-8 of &urface 'texturt and Lay on Friction and Scuffing Behavior ..... 4-10 ~ ~ ~ N~unibci of Failures vs howurs Srice Overhaul - MTBF- 500 hr............4 4-4 't' Number of Failurc:s vs HWoes &nuriv Operation- MTBF 5000 hr........... 4-14 m.-I 4rotm~aily of Survival vs L/rF,)Ratio...................... ............ 4-15 r'Cig j i'ý ~ pli~ivs Hertz Stress...........................................41 F.i 4 -Ii 14cibull Plot - Spalling Life vsGear Population Rank .................... ...... 4-"I ig44 'T pica 1 aiVPRotnýr Gearbox - Vulnerable ...................... 4A-19 ~f%4 '~~iUoto~earo-%X n zxo ....-. ...... 1........................4-% jW IC Zvý 0 ote,; Gcarbois - 12.7 Mmn Proof.............................. 11-21 .0,1% Typicwil Spects 1Powe.rFuncticn ......................................... 4-25 4S !2ý Li~~ee~ feCultvc,; ............................. ....... 42 I 4j4ig419 Shaft Horspowei ~cr His-Lograms................... ......... 4c$I~4-20 Orioph cRlationship - Fwilnte- Modes - Load v., 'eoi hy ............. 4-35 -A-f iraphic- Rel~ationship -- Failure Modes - Load vs Tjooth Siie........ 4-35 'Fiag. 4-22 . S5ingki VYoat!) Pusorf Gear Fatigue Test Results ........... I......I........ 4-41 f4. *-23 z vs SýWini: Velocity - Synchroniizd ani Ui~ytchreiiae Dics 442 -Tig4ý--_, C4i L -,,týiAtV>'wabke *i Suibsurfacxý Shear........ ........... ....... 4-45 4-46 -* Unsynchror.z. ................................ s V50tr.:b'cnt 4?i

.

34

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

:C.Tect~

,~

.

.

:

-4-26;~

j

-0ý1

2 4-

Cr~~ ear--Inrc~Ring Fit vw peratinTime ................ ........ c s ater Ring -ILinc~r Fit ReC$,4tion...... ............... -.:nGtigy 4 With ittcer Ring Expansion .................. -.-

7zrn al lieoad-

OF vs. Dl?.....................

Sp~t!nul on lh'Ttuhant Bearing fotccs................... Clearmnat LLv - lasic.anti -syj 4

..

1

4-49 4

45

-56

4

~-A

LIST 01- 1 I.V34TA~ll .o 0 0- Caef~iwwd) Fig. No,

*

Fig. 4-33 F-ig. 4-34 Fiwg 4135 Fig. 4-36~ Fig. 4-37 Fig- 4-38 Fig. 4-39 Fig. 4-40 Fig. 4-41 Fig. 4-42 Fig. 4-43 Fig, 4-44 Fig. 5-1 Fig. 5-? Fig. 5-3 Fig. 5-4 Fig. 5.5 5-6 Fig. 5-7

IFig.

RP

iý 5-8

*

Involute Spline Data 1, vs;N ... ....... ............ 4-60 Radial Mode Resonance 17ý_owiic vN Cear t'ooth Meshing Sipted ...... 41 Typical Spiral Damper Ring App~k~aions ............. .......... ........ 4-70A Relative Shaft Speed vs Pantive Vib~ration Amplitude......................4-74 Typical Bcaring Flangc~r As-_cmbl.V - Subci itical Shaf! Assembly .............. 4-75 Flc'-ible Diphragm Coupirr,... .................... ...........I... 4-77 Bosskcr Coupling ....................................... ............ 4-18 Elastom,:ric Coupling ................. ............................ 4-78 Hooke's Joint (Universal) ......................... ................. 4-9 Gear Coupling ................................................ ...... 4 Breakraway Sliding Foice vs Misalignment for Various Spline Devices ..... 4-79 Oil System Seheniawc.............I........... ............ ......... 49 Vectot Diagram of Swirl in 1-over ...................................... 5-1 Control Momyent for Basic tRotor Types. ..................... .......... 54 Articulated Rotor Schematic...................................... 5-9 Coincident Flap and Lag Hinge Rotor................................... 5-10 Girnb3led Rotor Schematic................ .......................... 5-if) Tcctering Rotot Schematic .............. ... .......................... 5-Il Teetet ing Rotor..................................... ................ 5-1; i-ingcicss Rotor Schcunaiti:.......................................... XH-51 Rotor System .................................................

5-13

Fig.5-11 Fig.5-12 Fif$. 5-13 Fig. 5-14 Fig. 5-15 Fig. 5-16 Fig. 5-17

Mechnismof Pitch-lag Instability.......... ......................... Pitch-nlap Coupling of Rotors ...................................... Mechanism of Flap-lag Instability ...................................... Typical Plots of Rotor Natural Frequency vs"OperatinF Speed ............... Single-degree-of-freedom Cole-man Plot............... ................ Two-degree-of-freedom Coleman Plot ............................... ... Two-degree-of-fteedvim Coleman Plot Showing Se~tisfaction of Minimum Frequency Criteria fur Two-blade IlingEcless Rotor...................... Gus* Load FactorComputec! for the UII-IB Helicopter Using I.ineai-Theory Gust-alleviation Factor (M IL-S-8698)................................. Rotor Limits as a Function of Advance Ratio........................... Results of a Load Gust Study Compared With Military Speciricauion Requiruments.......................................... Arliculatcd Rotor (Boeing Model 107) ............ ........ ............. CH--46 and CH-47 Tension-Torsion Strap Assemblies ...................... Torsionally Stiff and Flexible Wire-'vound Tic-bar Assemblies............... Elastomeric Bearings ................................................. Hydraulic Lag Damper............................................. CH-46 Power Blade Folding Mechanism................................. Typical Helicopter Rotor Blade Airfoils.................................. Track WiVth Varying rpm (Zero Collective Pitch) ........................... Tr-ack With Varying Collective Pitc!h (Constant Rotor rpm) ................. Alternating Stress Superimposed on Steady Stress ......................... Alternating Stress vs Cycles at Various Steady Stress Levels (Crv!ss Plotted from Fig. 3-12, MIL-HDBK- 17 for Notched Specimens of 181 Glass Fabric With MIL-R-7575 Polyester Resin) ................................. Propeller Flow Field for- Compound Helicopters........................... Comparison of P-order Excitations ......................................

51 i 5-16 5-16

Fig. 5-22 F;g. 5-23 Fig. 5-24 Fig. 5-25 Fig. 5-26 Fig. 5-27 Fig. 5-28 Fig. 5-29 Fig. 5-30 Fig. 5-31 Fig. 5-32 Fig. 5-33 Fig. 5-34 xxx

Title

Fig. 5-9

Fig. 5-18 Fig. 5-19 Fig. 5-20 Fig. 5-21

.

~

5-20

5-21 5-21 5-22 5-24 52 5-25

5-25 5-28 5-32 5-33 5-33 53 5-37 5-40 5-49 5-49 5-54 5-55 5-58 5-58

4.-

P 706- 20

4.,

LIST OF ILLUSTR ATIONS (Confinued)

Propeller IP Loads from Nonaxial Inflow.............................. 1IP Excitation Diagram for Typical STOL Aircraft......................... 1IP Excitation Diagram for Helicopter With Pusher Propeller................ Propeller Critical Speed Diagram ...................................... Propeller Vibration Modes ........................................... Stall Flutter Design Chart ............................................ Airfoil Characteristics and Stall Flutter ................................. Propeller Control System Schemnatic.................................... ............................ Simplified Propulsion System Block Diagramn Linearized Propellev Control Block Diagrai'n............................. Typical Blade Cross Sections.......................................... Typical Spar-shell Blade.............................................. Blade Materials and Weight Reduction ............................... Fatigue Strength Diffcrence Between Specimen and Full-scale Tests...... ..... Typical Stress Summary Curves ....................................... Geometric Da3ta.................................................... Fin Separation Distance/Rotor Radius ................................. Sideward Flight Velocity ............................................. Tail Rotor Performance, Four Blades................................... Typical Variation in Tail Rotor Noise Level.............................. Cutimpriisati-on" Ilgtie PD,-h n- Moment With Conine Angle and Blade CG....................................................

5-58 5-59 5-59 5-60 5-61 5-64 5-64 5-67 5-69 5-70 5-71 5-72 57 5-76 5-77 5-78 5-79 5-80 5-80 5-81

Fig. 6-I Fig. 6-2

Typical Control Function Scheduling for a Tilt-rotor Aircraft................. ....................... Characteristic Root Plot............ ............

6-3 6-5

Fig. 6-3 Fig. 6-4 6-5 Fig. 6-6 Fig. 6-7 Fig. 6-8 Fig. 6-9 Fig. 6-10 Fig. 6-11I F ig. 6- 12 Fig. 6-13 Fig. 6-14 Fig. 7-1 Fig. 7-2 Fig. 7-3 Fig. 7-4 Fig. 7-5 Fig. 7-6 Fig. 7-7

Allowable Pitch Czintrol System Residual Oscillations ..................... Control Mixing Schematic ............................................ Mechanical Mixing Assembly......................................... Powered Actuators (Tandem Helicopter)............ .................... ....... Artificial Feel and Trim Schematic.............................. Rotating Controls .................................................. Typical Pitch '-ink Rod End .......................................... Centrifugal Force Deflections......................................... Pitch Link Adjustment Provisions ..................................... Relative Pitch Link Rod End Position................................... Instrumented Pitch Link ............................................. ..................................... 1Instrumentcd Drive Scissors .... Typical DC Power Distribution System ................................. Typical AC Power Distribution System ................................. Example Load Analysis AC Left 1-and Main............................. Examiple AC Load Analysis Format ................................... ........... Typical Automation Flow Chart ........................... Typical AC Generator With Oil-lubricated Bearings....................... _............ DC Starter/Generator .................................

6-8 6-IS 6-15 6-16 6-17 6-19 6-20 6-20 6-20 6-21 6-22 6-22 7-3 7-4 7-5 7-6 7-7 7-8 7-10

Fig. 5-35

~Fig. 5-36 ;

*Fig. *Fig.

Fig. 5-37 Fig. 5-38 Fig. 5-39 Fig. 5-40 Fig. 5-41 Fig. 5-42 Fg5-43 Fig. 5-44 Fig. 5-45 5-46 5-47 Fig. 5-48 Fig. 5-49 Fig. 5-50 Fig. 5-51 Fig. 5-52 Fig. 5-53 Fig. 5-54 Fill. 5-55

)Aft

*Fig.

-AFig. )

5-002

Clast-cooled DC Generator ...........................................

7-12

Fig. 7-9 Fig. 7-10 Fig. 7-1l Fig. 7-12

DC Starter Motor With Solenoid-operatcd Switch ......................... Prolotype Cartridge-boosted Electrical Starter Systemn..................... Sample Set of! tilization Loads ....................................... Gases Emitted from Nickel-Cadmium Sintered Plate Cell During Overcharge

7-12. 7-13 7-18

Fig. 7-13

PermiisihleClamp Deformation

7-8

*

Page

Title

Fig. No.

at

.7-19

70'-75*F.........................................................................

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

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

7-2 5

a.

AMCP 706-201 LIST OF ILI-LJSTRATlONS8Conlinued)

Fig. No.

Title

Page

Fig. 7-14 Fig. 7.15 Fig. 7-16 Fig. 8-1 Fig. 8-2 fig. 8-3 Fig. 9-I Fig. 9-2 Fig. 9-3 Fig. 9-4 Fig. 9-5 Fig. 9-6 Fig. 9-7 Fig. 9-8 Fig. 9-9 Fig. 9- 10 Fig. 9-l11 Fig. 9-12 Fig. 9-13

Terminal Strip Installation............................................. Typical Connection to Grounding Pad................................... Typical Lightning Electrical Circuit Entry Points .......................... Block Diagram of Classical Communication System ................. ...... Typical Intercommunication Selector Box ................................ Typical Cnimunication Antenna Layout.................... -.... I...... Central Hydraulic System ........... .................... ............. Dual System Hydraulic-powcred Flight Control Actuators .................. Dual-powcrcd Stability Augmentation Systcm............................. Dual-po%-:rcd Stick Boost I ydraulic System.............................. Hydraulic Starting; Energy-limited System ................................ Hydrauix Starting. Power-iirnited System............. ................... APU Starting System................................................. Cargo Door and Ram~p System ......................................... Cargo and Personnel Hoist (Constant Prcssurt) System ..................... Rotor Brake System .................................................. Wheel Brake System .................................................. Combined Spool Switching Vakve ....................................... Pressuie Check Valves Plus Power Return Switching........................

7-26 7-27 7-31 8-4 8-4 8-13 9-I 9-2 9-2 9-2 9-2 9-3 9-3 9-3 9-4 9-4 9-5 9-8 9-8

Fig. 9-14

Pressure Check Valves Plus Inline Return Relief Valve......................

9-9

Fia 9! 5

Fig. 9-16 Fig. 9-17 Fig. 9- 18 Fig. 9-19 Fig. 9-20 Fig. 9-21 Fig. 9-22 Fig. 9-23 Fin. 9-24 Fig. 9-25 Fig. 9-26 Fig. 9-27 Fig. 9-28 Fig. 9-29 Fig. 9-30 Fig. 9-31 Fig- 9-32 Fig. 9-33 Fig. 9-34 Fig. 9-35 Fig. 9-36 Fig. 9-37 Fig. 9-38 Fig. 9-39 Fig. 9-40 Fig. 9-41 Fig. 9-42 Fig. 9-43 Fig. 9-44 xxxil

Irinein Mehnicafy LockeC-out

-cifVie................

Cam-operated Poppet Switching Valvec....... ................... ....... Switching Valve...................................................... Hydraulic System Ground Fill Provisions ................................ Rosar, Boss Fitting ................................................... Use of an Articulating Link ................... ................ ........ Use of Protective Cover on H-oses ....................................... Typical Mission Requirement Profile .................................. Examples of Parallel and Series Control Modes............................ Schematic of Jet Pipe Electro-hydraulic Control Valve ..... ................ Schematic of Flapper Electro-hydraulic Control Valve...................... Typical Master Control Valve.......................................... Anticavitation Approaches ............................................ Feedback Techniques ....................... ......................... Hydraulic Pump Flow vs Pressure Characteristics .......................... Hydraulic Pump Soft Cutoff Characteristics .............................. Hyaraulic Pump Case Drain Flo%%Characteristics ......................... Suction Line Length - Reservoir Pressure Characteristics............ ...... Hydraulic Pulsation Suppressor ........................................ Filler Element Dirt-holding Characteristics ............................... Filter Element Performrince ........................................... Hydraulic Valve "Trail" Configurations ................................. Hydraulic Valve Configurations........................ ................ Direct-operated Valve ......................................... ....... Pilot *operated Valve ................ ................................. Valve Operating Time................................................. Solenoid-operated Valve Incorporating Rcl urn Pressure Sensing .............. Power and Spring Main Section Valve Return to Neutral.................... Typical Separate Servo Actuator ........................................ Dual Seals With Return Vent ............................ ..............

-

9-9 9-10 9-1l 9-1l 9-12 9-12 9-12 9-14 9-IS 9-IS 9-15 9-17 9-17 9-18 9-18 9-18 9.19 9-19 9-21 9-21 9-22 9-22 9.23 9-23 9-23 9-24 9125 9-26 9-27

cP

LIST OF ILLUSTRATIONS (Continued) Fig. No.

Title

Page

Fig. 13-8 Fig. 13-9

Methods of Raising the Suspension Point .. Helicopter Load Dynamics Schematic .................. Study Input Variables ................................................. Equivalent Steady Load for Combination of Steady and Vibrator), Loeds (Nonreversing) .................................................... Equivalent Steady Load for Combination of Steady and Vibratory Loads (Reversing) .................................. Mounting of Duplexed Bell Bearings..................................... Spherical AicatB aig .........................

13-16 13-18

16-6

Common.....f onetos........................

16-18

*Fig.

14I Fig. 16-1 Fig. 16-2

Fig. 16-3 Fg 164

Fig. 16-5 Fig. Fig. Fig. Fig.

Fig. Fig. Fig. Fig. Fig. Fia.

16-6 16-7 16-8 16-9 16-10 16-11 16-12 17-1 17-2 17-3

Fig. 17-4 Fig. 17-5

~i.176 * * *

Fig. Fig. Fig. F.g. Fig. Fig. Fig.

17-7 17-8 17-9 17-10 li-Il 17-12 17-13

Fig. 117-14 15 Ig

xxxiv

(

014-17 C

16-7 16-9 16-13

Tapered Pipe Thread Fittings........................................... 16-22 Straight Thread Fittings ............................................... 16-23 Flared lube Fittings.................................................. 16-24 Flareless Tubc Fittings ................................................ 16-24 Cable Alignment and Pulley Guai d Location.............................. 16-27 Pus~i.pull Cablus and End Fittings..................................... 16-29 Right-handed Thread Application of Safety Wire......................... 16-33 Standard Bend Radii PrActice -Minimum Bend Radii.......I...... I........ 17-5 Weld Contour and Stress Concentration ....................... 17-7 Welding Symbols.......................................... 17-8 Reprcsentative Buit Juivinb................................................ 17-8 Representative Corner Joints........................................... 17-9 Representative Tee Joints..............................................19 Rivet Spacirg ........................................................ 17-10 Types of Loading for Bonded Joints..................................... 17-12 Lap Shear Joint Deflection Under Load .................................. 17-13 Typical Rotor Blade Design - Alternate I................................ 17-13 Typical Rotor Blade Design - Atlernate 2................................ 17-13 Hioneycomb Sandwich Structure ........................................ 17-14 Addition of Doubler-' to Honeycomb Structure............................ 17-14 Balance Bar Design ............................. ..................... 17-15

as........eWngRr

Fig. 17-16

I...................

*2

.r

--

D....................

...........

Cable Splicing..................................

... ..

..

.. ..

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

176

17-16

-

". 't

LIST OF TABLES Table No.

Title

Page

TABLE T"ABLE YABLE TABLE TABLE

2-I 2-2 2-3 2-4 2-5

Mechanical Propcrtii.s of 18 Ni Maraging Steels ............................ Comparative Mechanical Properties for Selected Nonferrous Alloys ......... Grouping of Metals and Alloys (MIL-Si D-889) ............................ Position of Metals in the Galvanic Series ................................... Process Comparison Guide for GRI' Laminates ............................

2-4 2-4 2-7 28 2-14

TABL E 2-6

General Properties Obtainable in Some Glass Reinforced Plastics ............

2-15

TABLE 2-7 TABLE 2-8

TA BLE 4-2 TABLE 4-3 TABLE 4-4 TA BLE 4-5 TABLE 4-6 TAB! F 5-! TAB[E 5-2 TABLE 5-3

Common Resin Rcinforcemen't Combinations of Thermoset Laminates ...... 2-16 Typical Values of Physical and Mechanical Characteristics of "Reinforcement Fibers ................................................... 2-18 Nominal Composition of Glass Reinforcements ............................ 2-19 Typical Unidirectional Composite Properties Based on Commercial Prepregs 2-20 Properties of Rigid Foams ................................................. 2-25 Common Adhesives in Current Use ....................................... 2-26 Shear Bond Strengths of Adhesives ..................................... .. 2-26 Useful Temperature Range and Strength Properties of Structural Adhesives .. 2.2k Armor Material Design Data and Physical Charactetistics .................. 2-28 Fabrication Data for Lightweight Armor Materials ........................ 2-2') Typical Properties of Commonly Used Structural Adhesives ................ 2-32 Helicopter Lubricants and Hydraulic Fluids ................................ 2-3) APU Types for Main Engine Starting Environmental Control, and E lectrical Supply .................. .................................... 3-16 APU Reliallity ..................................................... 3-20 U S Army Heliconters - Transnmission ,ind nrive-Sytem. On.y 4-17 Maintenance Workload .......................................... External N oise Level ......................... .......... ..... ........... 4-18 He!icopter Drive Subsystems - Single Main Rotor. ....................... 4-24 .. 4-33 Life Modification Factors - Surface Durability ......................... Shear Stress vs Depth ............... ...................... ............. 4-45 Helicopter Transmission Case Materials and Application Data .............. 4-68 The Relative Effects of Various Parameters on Gust Response ............... 5-26 Example of Nominal Weight and CG Locations ........................... . 5-48 Rotor Blade Balance (Sample) .......................................... .. 5-48

TABLE 5-4 TABLE 5-5

Comparison of Material Properties ................................... Aerodynamic Characteristics of Several Airfoil Sections Suitable for

TABL -5-6 TABILE 6-1 "TABLE7-1 TABLE 7-2 TABLE 7-3 TABLE Il-I TABLE 11-2 TABLE 12-1 T A BLE 13-1 TABLE 13-2 TA BL E 14-1 TABLE 14-2 T A B LE 14-3 TABLE 16-1 TABLE 16-2 TABLE 16-3

Summary of Tail Rotor Excitation Sources ................................ Maximum Amplitudes of Limit-Cycle Oscillations .......................... Outputs of Converters Relative to Continuous-Current G-ncrator ........... Typical Characteristics of 24V. 34 All Battery Systems ................. ... Alternative Charging Methods ....................................... Cost Impact. Airframe Detail Design ..................................... M aterial Selection - Airframe Design .................................... Load Factors for Helicopter Tire.s ......................................... C oefficients of Friction .......................................... ......... Standard Cargo T iedessn Devices .......... ....... ....................... Typical H elicopter G uns .................................................. Vulnerability Damage Criteria Data Sumnmar% ............................ Vulnerability T able ....................................................... Life I actors for Antifriction Bcaring M aterizis ............................. Cost vs Tolerance Class for Antifriction Bcari.igs .......................... Standards for Airfrname Control Annular Ball bearing. .....................

"TABLE 2-9 TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE 'ABLE

,'

2-10 2-11 2-12 2-13 2-14 2.15 2-16 2-17 2-18 3-1

TABLE 3-2 TABLE 4-1

T ll

__\

II

RiB.tU d esic.........

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

2

5-51

. .........

5-82 6-8 7-15 7-16 7.17 . -. , 11-3

12-4 13-14 13-15 4-2 14-14 14-14 16-6 lo-7 16-12 \•

mm

-9.•

'••

•,.,•,

•

.,.:.:.••.

..

S.11

AMOP MG-202 I.T01:"1AI ABIS ("outinutcd

Table No. TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE

16-4 16-5 16-6 16-7 16-8 16-9 16-10 16-I 16-12 17-1 17-2 17-3 17-4

TABLE 17-5 TABLE 17-6

*

_

(.p

Titlc

Pagc

Standards for Airframe Control Rod End Bcaring.i ......................... Standards for Sphe'ical Roller Airframe Bearings .......................... Properties of Sliding Bearing Materials for Airframe Use .................... Specifications and Standards for Self-Lubricating Slide Bearings ............ Militery Specifications and Standards for Connectors for Aircraft ........... Other Military Spccificationi and Standards for Connectors ................ Military Specifications and Standards for Crimp-Style Terminals ............ Military Specifications and Standards for Switches ......................... Military Specifications for Cables .................................... Bentd Characteristics of Selected M etals .................................... Unit Horsepower Values for Represerntative Metals ....................... Representative Surface Finishes Obtained in Machining Operations ......... Values To Be Added to or Subtracted from Base Dimension for "Holesand Shafts To Calculate Tolerance ................................ Representative Heat Treat Temperatures .................................. The Effect of Shot-peening on the Fatigue Properties of Selected Samples

16-12 16-12 16-13 16-14 16-17 !(-18 16-20( 16-21 16-31 17-4 17-6 1'7-6

, -1

17-7 17-i 17-21

I

S

-,-.

\\~

!. -

1 ~-~-

AMCP 706,201

The IIhli,•p:er Lngineering llandbook form. ai part-of the 'ingincering Design Handbook Series shich presents engineering data for the design and construction of A -y equipment. This volume. AMCP 70M-202. Delail De.%ign. is Part Two of a three-part Enginering Design Handbook titled Helicopter Engineering Along with AMCP 706-201. Preliminary Design. and AMCP 706-203, Qualification A.tsurance, this part is intended to set forth explicit design standards for Army helicopters, to establish qualification requirements. and tu provide technical guidance to helicopter designers, both in the industry and within the Army.

a

This volume, AMCP 706-202, deals with the evolution of th', vehicle from an approved preliminary design configuration. As a result of this phase of the developmcnt. the design is describcd in sufficient detail to permit construction and qualifica"tion of the helicopter as being in compliance with all applicable requirements, inchiding hce approved system specification. Design requirements for all vehicle subsystems are included. The volume concists of 17 chapters and the organization is discussed in Chapter 1. the iutroduction to the volume. AMCP 706-201 deals with the preliminary design of a helicopter. The characteristics of the vehicle and of the subsystems that must be considered arc described.

ir

and possit '- solutions at ' suggested. The documentation necessary to describe the

J

"preliminarydesign

in sufficient detail to p,-rniit evaluation and approval by the procuring activity also is described.

A

The (hird volume of the handbook, AMCP 706-203, defines the rcquirements Ifor airworthiness qualification of the helicopter and for demonstration of contract cornpliance. The test procedures used by the Army in the performance of those additional tests required by the Airworthiness Qualification Program to bt performed by the Army also arc described.

xxxvii

F.

AMCP 706-202

PREFA(I: This volume, AMCP 706-202, Detail Design. is the %econd section of a three-part -ngitieering handbook, Heficopier Engineering. in the Engineering D'%sign Hand"book series. It was prepared by Forge Aerospace. Inc.. WAshington. D.C., under subcontract to the Engineering Handbook Office, Duke University, Durham. NC. The Engineering Design Handbooks fall into two basic c-'tegorics. those approved fot release and sale, and those classified for security reasons. The US Army Materiel Commano policy is to release these Engineering Design Handbooks in accordance with current DOD Directive 7230.7. dated 18 September 1973. All unclassified Handbooks can be obtained from the National Technical Information Service (NTIS). Prozedures for acquiring these Handbooks follow: a. All Department of Army activities having need for the Handbooks must submit their request on an official requisition form (DA Form 17, dated Jan 70) directly to: Commander Lettcrkenny Army Depot ATTN: DRXLE-ATD Chambcrzburg. PA 17201 (Requests for classified documents must b,. submitted. mith appropriate "Need to Know" justification, to Letterkenny Army Depot.) DA activities will not requisition Handbooks for further free distribution. b. All other requestors. DOD, Navy. Air Force. Marine Corps. nonmilitary Government agencies, contractors, private industry. individuals, universities, and others must purchase these Handbooks from:

.

National Technical Information Service Department of Commerce Springfield, VA 22151 Classified documents mey be released on a "'Need to Know" basis verified by an official Department of Army representative and processed from Defense Documentation Center (DDC), ATTN: DDC-TSR, Cameron Station, Alexandria, VA 22314. Users of the handbook are encouraged to contact USAAVSCOM, St. Louis, MO, System Development and Qualification Division, with their recommendations and comments concerning the handbook. Comments should be specific and include recommended text ¢hangcs and supporting rationale. DA Form 2028, Recommended Changes to Publications (available through normal publications supply channels) may be used for this purpose. A copy of the comments should be sent to: Commander US Arny Mateviel Development and Readin•es Command Alexandria, VA 22333 Revisions to the handbook will be made on an as-required basis and will be distrbutcd on a normal basis through the Letterkenny Army Depot.

t

AMCP 706-202 (CHAPTER I

INTRODUCTION AMCP 706-202, Engineering Design Handbook, Helicopter Enginre.-ing. Parr 7wo. Detail Design, is the second part of a th:¢¢-volume hlicopter engineering design h~ndbook. The preliminary design (covered in AMCP 706-201) is *vclopcd during the proposal phase. at which time all subsystems must be defined in sufficient detail to determine aircraft configuration, weight, and pcrformance. The detail design involves a reexamination of all subsystems iri order to define cach clement thoroughly with the aims of optimizing the aircraft with regard to mission capability as well as cost considerations. Detailed subsystem specification requirements are the basis for in-depth analysis and evaluation of subsystem charactcerstics and interfaces. Based upon complete system descriptions and layouts, performance, weight, end cost trade-offs arc finalized. Periodic reviews of the design are conducted to evaluate mairtainability, reliability, safety, producibilhty. and .oriforniancc with spc.i.fication requiremeents. Development testing may be required to permit evaluation of alternate 5olutions to design problems or to obtain adequate information for trade-off investigations. Appropriate consideration of human engineering factors often requires evaluation of informal mock-ups. SWeight control is an important element of the detail design phase. Subsystem weight budgets, prepared on the basis of the preliminary design group wCight hreakdown, arC aasiýzsed at the initiation of

the detail design phase. The continui-ig evaluation of compliance with the budget as an essential part of the manzgement of the project and the awsurancc of cornpliance with weight guarantees of the helicopter detail specification are described in conjunction with the discussion of the Weight Engineek'ing function in ANICP 706-201. The requirements and procedures for airworthiness qualification and proof of contract compliance for a new model helicopter for the US Army are defined and discussed in AMCP 706-203. which ikthe third volume in this handbook serits Qualification is not time-phwed. but is a continuing part of the acquisition program. A number of qualification requirements are integral parts of the detail design effort. Desipr. reviews by the procuring activity are re(quired during the definition of subsystem configuraons as well as during the final design of assemblies --.

and tinstallations. Evrluation of a full-scale mock-up of the complete helicopter is a major part of the design review process. The requirements for this review are described in dctail in AMCP 706-203; but the construction and inspection of the mock-up must be completed at the earliest piactical point of the detail design phase to permit the contractor to complete the desibn and manufacture of helicopters for test and for operational deployment, with reasonable, assurance tOat the configuraion is responsive to the mission ecquircments. Also completed during the detail design ohase are a variety of analyses necessary to substantiate the comp!ianci of the physical, mechanical, and dynamic characteristics of subsystems and their key cornponents with applicable design and performance requirements, including structural integrity. The analysis required during the design, development, and qualification of a given model helicopter are those specified by the applicable Contract Data Requireicis List ('Dr , L-). This volume reviews the functions pcrformre by the major helicopter subsystems and outlines the requirements app!icaolr to the design and installation of eaich one. Principal documentation of the detail design phase is the final drawings of the helicopter in sufficient detail for procurement, fabrication. assembly, and installation. This volume, therefore, aiso iincludes discussion of materials and processes pertinent to the ),-nnufacture of bclicopter components. This volume is intended to pruvide designers. engmnrr,

relaltvelv

rew to the

helicorptev tech-

nology, and program managers a general design guide covering all of the helicopter detail design specialties. however, it is not intended as a source of detailed design procedures for use by the experlcnced design engineer in his specially. Throughout this volume the mandatory design reouiremcnts have been identified with the contractual language which makes use of the word ".shall". To assist in the use of the handbook in the planning or conduct of a helicopter development program, the word "shall" has been italicized in the state.ment of each such requirement. Since: the mission requirements for individual helicopters result in variations between subsystem configurations and performance requirements. the procuring activity will specify in its Request for Proposal (RFP) the extent to which the design requirements of this handbook are applicable to the acquisition of a given helicopter. I-I

.. • ,

AMCP 7_06-202

CHAPTER 2

MATERIALS

)

2-1 INTRODUCTION

factors, standard mill products, and cost data. tl,.IL-

This chapter addresses the properties of the various materials used in the construction of helicopters, These materials include ferrous and nonferrous metal. nonmetallic materials, composite structures, adhesives and seplants, paints and finishes, lubri., cants, greases, and hydraulic fluids. Among the ferrous me:als are carbon steel, stainless steel, and al:oy steels. The nonferrous mttals include aluminum, magnesium and titanium alloys, beryllium, copper, brass, and bronze. Thermoplastic and thermosetting plastics, elastomers, woods, fabric, and fluoroplastics are reviewed. Composite structures, including filament laminates, fabric laminates, and filament wound, and honeycomb and sandwich construction ake discussed aý are the adhesives used for bonding of primary structure, honeycomb and compo3ites, fahris, rbhhr, e'astomers, g1.s; and rlastics. Sealhng compounds, such as putty and pastc;, also are de.ai!ed. A discussion of paints and coatings, special finishes, plating, and tapes is included, as well as a review of th,. most commonly used lubricants and their applications. The designer will find that a good working relationship with vendors wii! help him to keep abreast of new rtiatvrials and processes with possib!t applications to helicopters. New materials are being iiitroduced continually, and new processes alter the cost and performance relationships among older

and provides additional detail design data. For technical data and information pertaining to wrought iron, carbon steels, and low-alloy steels, an exccllent source is N4IL-IIDBK-723, which covers some of the more practical aspects of -ecta! forming and joining. Finally, for design data and metallurgical details, the designer should consult the various American Society of Metals (ASM) Handbooks. Because of weight considerations, it is desirable to restrict the heavier ferrous metals to those applications where very high strength, a high modulus of rigidity, high resistance to fatigue, and high modulus of elasticity are reqaired. The more expensive high-performance steels often are more economical in terms of weight, cost, and fatrication processes than are the lower-cost ferrous products. Anplications for these materials include high-sirt-ss parts such as rotor drive shafts, masts, hubs. vertical hinges, flapping hinges, tie cables, ti'bular frames, an6 control cams, keys, gears, and hydraulic cylinde-.s.

, -.

.

HDBK-5 is a source of mechanical properties data

..

conform to one or more Government specifications, and many manufacturers take step, to keep their products on the qualified-products lists, where such are requiret. Propei regazd for and awareness of these concerns in design callouts will simplify procurernent, fabrication, and qualification of hard•arc.

2-2 METALS FERROUS METALS

2-2.1

R"

2-2.1.1 General Thik dis•:ussion provides a brief review of ferrous metalk and their application to the construction of helicopters, as well as of some of the parametcrs go)vernling the choice of a particular ferrous metal for specifi as a AP•'','•, moreconpreheisivc discussion, as well as detail \,dcsign data, will be found in Chapter 9, AMCP 70%hich describes such items as mater'al selection r

2-2.1.2 Carbon Steels The carbon steels are a broad group of iron-base alloys having small amounts of carbon as their principal alloying element. Commonly, the carbon content falls between 0.03 and 1.2%. The American Iron and Steel Institute (AISI) codt usually is used for designating s•ate ,. Tins ...ystm i.i..i Uup.. a Or five-digit number to designate each alloy, with the first digits referring to the alloy and the last two digits giving the carbon content in points of carbon. where one point is equal to 0.01%. Thus, 1045 steel is one of a series of nonsulphurized carbon steels and has 0.45% carbon. Other carbon steel s.ies are the II XX series, which are resulpherized; the BIIlXX series, which are acid Bessemer resulphurized, and the 12XX series, which are rephosphorizcd. Low-carbon stee's range from 0.05 to 0.30% carbon, mediumcarbon steels range from 0.30 to 0.60% carbon, and high-carbon steels range from 0.60 to 0.95% carbon. The machinability of low-carbon steels is poor. They tend to drag and smear and to build up on the cutting edges of tzols, generating considerable heat and decreasing cutting cffiricncy. Medium-carbon steels machine .better, although the cutting pressures are higher. High-.arbon steels are too hard for good machining, but they are used where fine finish and dimensional accuracy are required. Hot- and coldfuui--

accrac

27I

-:

-'

AMCP 706-202 rolied stels machinc better than do annealed steels, and the macbining properties of the low-carbon steels arc improved by 1he addition of sulphur. phosphorus. or lead. The low-carbon steels have excellent forming properties, and can be worked readily by any of the normal shapin3 processes. Their ready formability is due to the fht that there is less carbon to interfere lans th sam tokn, heidr-. withthe ofslip witulty plao increafsesp. witheinmereasing carbon df ok i a wtough ctn content. Plain carbon steel is the most readily welded of all materials. Low-carbon (0.15%) steel presents tke least

rosion resis,.anc is about the same. hih I lie AISI designation system is used for alloy steels also. This is illustrated by 4130 steel, -hich Is an alloy steel coltainin.g chromium. molybdenum, and 0.30% carbcn. 2-2.1.4 Stainless Steels All stainless stees contain at least 10.5% chrochro1s at least eellent nless Ai from which excellent corrosion resistance is obtained. Apparently, a very thin, transparrnt, and film of oxide forms upon the chromium surface. This film is inert, or passive, and does not react upon exposure to corrosive Materials. There are three broad types of stainless steels. as defined by thee

.

cf. content as the carbon difficulty. a result toof 0.30%, rpidmartensitic. some martcnsite may form as increases coling. ftheyare coyfole too arapidlt aftr ldi, Austenitic stainless steels, which have an austenicooling. If they Arc cooled too rapidly after welding. tic structure at rom temperature, arc known as the 3W series (AISI). These materials have excellent ducmedium- arid hil[.-carbon steels may harden, but pretilit, at heating to tM00 0F or post-heating to I 100'F will very low temperatures, the highest corrosion t remove britde nicrostructures. resistance of all steels, and the highest scale reThe yield strength of low-carbon steels is on the temperatures. strength at elevated itancc and steels is of high-cat bon psi, -while that be order of 46.0000,000 m achine, but steels are difficul, ustcnitic hncr A eelstioof fselastio canok modulua s ie toot4rt fom. psi.while psi.thtThe modulus 4of of 150,000 theorder order or. or the formd when carc is given to the r-te of workv miiiion for aii m in tension remains at 3S0 city ban steels. Core (Brinell) hardness ranges from 4 haideniiig. They art not harderiable by heatt rcatfrobon steels.Cove 40Brfnlhihardnessranges c rbom, ment. Welding is 'one best in an inert atmosphere, becae f the low thermal conductivity, care must be for low-carbon to 400 fo higher carbon. 2-2.1.3 Alloy Steels Alloy steels are those that contain significant amounts of such alloying metals as manganese, molybdenum, chromium, or nickel, which are added in order to obtain higher mechanical properties with heat treatmoent, especially in thick sections. A family of extra-high-strength, quenched, and tempered alboy steels has come into wide use because these c...a. rial.s havc y..ld s;reingths. of more than IArlflW psi.

The alloy steels have relatively good resistance to fracture, or tough-ess. Weldability is good, and machinability and castabihty are fair. The alloy steels generally can be hardeneC to a greater depth than can unalloyed steels with the same carbon content. Many of the alloy steels are available with added sulphur or lead for improved machinability. However, resulphurized and leaded steels are not recommended for highly stressed aircraft pa. ts because of drastic reductions in transverse properties. The alloy steels arce somewhat more difficult to forge than are the corresponding plain carbon steels, and the maximum recommended forging temperatures are about 50%

taken to avoid cracking. Carbide precipitation is minimized during welding by selecting one of the stabilized grades, e.g., 321 or 347. Ferritic stainless steels are magnetic and have good ductility. Because of the low carbon-to-chromiumn ,atio, the effects of thermal transformation are eliminated and the steels are not hardenable by heat treatment. They also do not work-harden to any grcuit e-tent. are machined easily,. and arc formed type readily. A general-purpose ferritic stainless is istp 430. Martensitic steels have a higher carbon-to-chroimiurn ratio and are hardenable by heat treatment. They are characteriLed by good ductility, hardness, and ability to hold an edge. These steels are. magnritic in all conditions, are tough and resistant to impact, and attain tensile strengths of up to 200,000 psi when hardened. Martensitic steels machine very well. Type 410 is the most widely used steel in this group. .,A

2-2.1.5 Precipitation Hardening Steels Precipitation hardening (PH) steels arc those that harden at relatively low temperatures due to the pre-

lower

Lipitalion of copper, aluminum, or titanium inter-

Cold-forming. it performed, is done in the annealed con6ition because of the high strength and limited duaility of heat-treated materials. Notch toughness of alloys in the heat-treated condition is much better than that of the carbon steels. Cor-

metallic compounds. They may be nonstainless or stainless. The best known is 17-4 P11, which is stainlcss by composition and is used fcr parts requiring high strength and good resistance to corrosion and oxidation at temperatures of up to 6 00F. 17-4 PH is

2-2

"

,•

:% "".

4,4

4-

.enerl-p

.

AM,,i /06 12 inartensitic in nature, but other precipitation hardening steels may be austenitic. Forming properties are .nuch the same as for stainless steels: forming must he accomnplished before heat treatment, and allowance must be made for the dimensional changes that occur during the hardening process. Strcrngth properties ate lowered by exposure to temperaturrs abosc 9750: for longer than 0.5 hr. The heat-treating procedures are specified in MIL-H-6875. •

--

2-2.1.6 Maraglng Steels The maraging steels are not treated in the refcrences given in par. 2-2.1.1; henc¢ they arc discussed in somewhat greater detail here. The term "maraging' is derived from the capability o! the material for age hardening in the martensitic condition. The distinguishing features of the 18% nickel maraging steels are that they arc designed to be martensitic upon cooling te room temperature after hot-working or annealing, and t- be agehardenailc to ultra high strengths in that condition. The 18% nickel maragirg steels essentially are wrought alloys. The nominal yield strengths of four "well-eitablished grades are 2Q., 2M), 3M).). and 35( ksi, The ability of these steels to transform into martensite upon cooling from elevated temperatures is im'paried by their nickel content. The transformation, which begins at about 310'F and ends at about 210°F, is of the diffusionless or shearing type. The formation of martensite in these steels is noi disturbed by varying the cooling rate within practicalec limits. Hence, section size is not a factor in the process of martensite formation, and the concepts of hardenability that dominate the technology of the quenched and tempered steels are not applicable with the maraging steels. The l8Ni maraging steels may be cut with a saw in the annealed or hot-worked condition. Alternatively, oxyacetylene and plasma arc torches may be used. Hot-rolled or annealed maraging steels can be sheared in much the same manner as can the quenched and tempered structural steels that have yield strengths in the vicinity of 110 ksi. In grinding, these steels behave in a manner similar to that of stainless steels, using a heavy-duty, water-soluble grinding fluid. The maraging steels can be hot-worked to finished "or semi-finished products by all of the standard methods of forming that are used for other steels. To avoid carburizing or sulfidizing, the metal should be free of oil, grease, and shop soil before heating. Fuel "withlow sulphur content is preferred. The meta! can be press- or hammer-forged at temperatures ranging 'from 23000 down to I500°F. Forging is completed at ý,"latively low temperatures. The objective is to refine

S~2-3

the grain struchtre, thereby enhancing ihe strenpth and toughness of the sleel. A minimum redution of 25'v. in thickness during the finil forging cycle is recommended to produce optimum mectianical propcrtices in the finished product. Ho, bending, hot drawing, and hot spinning are accomplished at 1500O- 180O F:. Cold-forrnirg operations are performed on the annealed material. Even in the annealed condition, the 18Ni maraging steels have yield strengths of up to 120 ksi, approximately four times those of deep-drawbody stock The tensile elongations of these steels in the form of annealed sheet may be as little as 3-4%. These factors impose limitations upon forming the sheet metal by tensile stresses. On the other hand, these steels work-harden very slowly, making them well suited to formirL methods dominated by shear. They can be cold-reduced by 80% or more, and shape. are formed readily by rolling or spinning. Flat-bottom cups can be deep-drawn to considerable depths. Roundtd shapes are formed mi re readily by means of the flexible die process. Cold-rolled, solution-annealed material is preferred. Rolling and welding of shwc, strip, and plate are cmm1on mithods of making cylindrical shapes. In the annealed condition, the l8Ni maraging steels are machined quite easily. In the age-hardened condition, machining is difficult because of the hardness imparted by the aging process. Although these steels have been welded by all of the common welding processes, the toughest welds are produced by the gas tungsten arc process or the electron beam kEB) process. For maximum toughness, the carbon, sulphur, silicon, phosphorus, and oxygen content must be kept at very low leveis. It is good practice to avoid prolongca times at elevated temperatures, not to preheat, to keep interpass ternperatures below about 250 *F, to use minimum weld energy input, and to avoid conditions causing slow cooling rates. Annealing is accomplished at 1500'F with air cooling. For improved combinations of stren~gth and toughness, the steel may be double-anne.aled. Thprocedure is to heat the material to !600°-l800'F, air cool to room temperature, reheat to 14000 -150O"F, and again air cool. Special furnarie atmosphe:-es art; required in order to prevent carburization, sullidation, or excessive oxidation. Age-hardening is accomplished at 9000)F, the time varying from 3 to 6 hr. Air is used commonly as the heat-treating atmosphere. It is advisable to maintain thc temperature at all part: of the load to within - 100 of the desired temperature. The nominal mechanical properties of the agehardened 18Ni tiaraging st~cts are listed in Table 2-I.

ca:~toas-4s thc: des~'icer in formulainig the re-

I~s~i

Adtona r~r~ daa h~h-t~cdgt

~

~

rnaraging btecis will be round a iv .I 2-.2.2 NON~FRUOUS MFITALS

o

d

E~nparativc

-

mcchfinical properties for mreprsenta-

lees3 ive 1aonluCW1_us alloys arc giken in Tabas- 2-2 Im;nIAly tlicii ols&.d Abrief review of nonferieus i-' af l Aus-a ý1 wco hk:,mprlrs kiusiv applicatkio to the cosru~m cur-wniirJ a as of some of [thc paramr.ters j,o,-vcrning *lýc Jioic; i.ofD~*{~ cus.ý. r of v'.-inum a~lcys, along with design at one metal amnoneg many for a a tý~la appficatiot,. smay1stdrdization ocz.uolacrIpe alonga with di-tail ticsigni Jita, is foune io AN11 Y nicnis, ir'cludiiig militury. fcdea ad, anid iindustry specilt00 arid in MIl.-1-DBI-5. ficatiowis. ThCSe Npe'.-ification, cover most of the uses Metais siuýh as alkmnrii¾m, in-gflesium, or iiof alulnihiaum in detail and shoutld IN consulted hefore rclativclv their tanium may be sailectcd beciuse of ein poadrgwt lilht weIghts. Other factors in material seteciion inailaminum allos are deexcept~or.s, fw Withi clectrica'i and thernial esstance. crrsin *eud nwogtpo o atn o o indete coraductivftty. lubricity, softness, cost and ease of, d-icts, but not for both. Althouth some genera!fabrcation. hardness. stiffness, and faoigpae resispurpose :-illoys are available. compositions rnoimally tance. UsAually, ii is the sum of a number of factors formulated so as to satisfy tspecific rqirements. arc of that in',1uences a designer to select the sequence mnoic %&idely used And readily avaiwable copiThe a constitute that processes materials and fabrication positions. are covered by G(.?vcinment specifications. design item. This discussion is inteiided to provide Most are adaptable to a variety of applications. The Aluminum Association has devised afor TABLE2-1 digit system -for wrought alloys in which the first 110*v fABL12N, ig% no " ý. A number designazes tnhe major anulloyi ungemiciii. Ti u.N. MARAGING ST EELS I is pure alu-minum, 'A is copper, 3 is manganese, 4 is

2-2.2.

*

Ertc

I

silicon, 5 is magnesium, 6 is magnesium and silicon, and V is zinc. The last two digits are supposed to ULTIMATE TE14SILf 00 designate the aluminum purity, but the exceptions 365. 00 294,000 260,000 210,000 NGTH,psi ST RE destroy. the rule. hlowever, the more frequently used 4 ~ 55,001290,000 255,000 000 SRENTH, si 0.2-YIED become familiar to the designer. The 0.~YILD * TREGTHPSI355000alloys aluminum casting alloys usually are identified by ar125 13.1 11.8 100 ELONATIN,-. ELN A IN,,IU0 T 10 1 . bitrarilv selected com mercial dicsignations of twoand three-digit numbers 2.0 570 6.0 0.0I OF~RE~o * IOREIJC NOT UCHTENSIL STREA' N50 0 6!. 2. Most alumninum alloys used for wrought products contain less than 7% of alloying elements. By regula33 qu00 420nn 1 190m00 I25.000 ~ STEGH ~ TE0IL NOTCH 19.0'0' -t tion of the amounts and types of elements added, the operties of the alumninum can be enhanced and its pr.0 17.019. 1. CHARFY V-NOTCHII-Is working characteristics improved. Special comLITIGU (0DUCYCLES 1000 2,00 115,000 115"U000 positions have been developed for particular fabrication processes, such as forging and extrusion. Wrought alloys are produced in both heat-treatable 1 75 60 55 HRDNCKESLS'~ HARDNESS0 nn nonheat-Lreatable types. The mechanical properties of the nonhecat-treatable materials may be varied 220026,0 8,0 siE1 STRENI~H, YIELDOPE

FSLRIES

-250

-30130

_

-2,

-------

-~

--

-

-

!

I

-

--. ------------

'

-

strain-hardening or by a combination of strain-

TABLE 24by

COMPARr

TABLECIA 2-2ERIE A

h.irdering and annealing&...

I he aluminum alloys specified for casting purposes contain one or more alloying! elements; the maximum . .n~u of any one element must not exceed 12%. designed Sonealoy are deindfor usc in the as-cast con-

COMPR~n E MCHAICALPROERTES FOR SELECTED NONFERROUS, ALLOYS PTY FROPE _

--

_-A

T

A29.C 14

1% STRENGTH. LL

IENSILL STRENGTH,.~

-.-ý

34

ELONGATION,

MOOIQ"US. 10Il

'Wo

Llui

6.5

2017

VANSM 'cr[oun:.

I C-i

I4A,~~IVCR

ition; others are designed to he heat-treated in order to improve their mechanical properties ;nd dimensional stability. High strength with good ductility can be o~btained by selecting the appropriate comnpo-

32 -

17

___

12

5

66

10.4

5.5

-

230 -HANUNESSIB ..

2-4

ndha

-

retet

~sitionanhetramn.

AMCP 706-202

* . S"

-

_. TN•

S,)

The hcat-trcatment and temper desinations for aluminum arc long and complex. The desinations most frequently stamped on products are: F-as fabricated; 0-annealed; H-strain-hardened (many subdivisions); T2- (cast products only). T4-solution lihat-treated and naturally aged. and T6-solution heat-treated and artificially aged. The heat trcstmcnt of aluminum alloys is detailed in MIL-H-6083. The processes commonly used are solution heat treatment. precipitation hardening, and annealing. A .small amount of cold-working aftcr solution heat "treatment produces a substantial increase in yield strength, some increase in tensile strength, and some loss in ductility. Rapid quenching will provide maximum corrosion resistance, while a slower quench-used for heavy sections and large forgings-tcnds to minimize cracking and distortion. Momt forming of aluminum is done cold. The ternperature chosen permits the completion of the fabrication without the necessity for any intermediate annealing. Hot-forming of aiuminum usually is perS formed at temperatures of 300°-400°F, and heating periods are limited to 15-30 min. When nonhcatitreatable alloys are to be formed, the tempef bhuuld be just soft enough so as to permit the required bend radius or draw depth. When heat-ireatable alloys are used, the shape should govern the alloy selected arnd its temper. To a great extent, the choice of an alloy for casting is governed by the type of mold to be employed. In turn, the type of mold is determined by factors such as intricacy of design, size, cross section, tolerance, surface finish, and the number of castings to be produced. In all casting processes, alloys with high silicon ,-onte~nt ee m~fmi in thwprndctinmi of narts with thin walls and intricate design.

mizt distortion due to expansion and contraction. Molten aluminum bsorbs hydrogen easily, and this may cause porosity during cooling. Because they provide a protective inert-gas shield. TIG and MIG welding are common choices. TIG is an incrt-gas shield-arc process with a tungsten electrode, and MIG is an inert-gas, shielded-metal-arc process using covered electrodes. A suitable flux, and mechanical (stainlcss steel brush) removal of the oxide film just prior to welding. are mandatory. Certain aluminum alloys - 2014, 7075, etc. - arc extremely difficult to fusion weld (excluding spot welding) and normally would not be used in structural applications when welded. Brazing is somewhat more difficult, and soldering of aluminum ir extremely difficult. The other joining processes include riveting and adhesive bending, both of which are used extensively in aircraft structures. Applications for aluminum in helicopters include the 4heet-metal exterior surface of the fuselage. framing, stringers, beams, tubing, and other usages where the density, corrosion resistance, and ease of fabrication of alum,'ium give it an advantage ovcr steel and where its high,, star-eg•,,sth, and oadulus properties give it an advantage over magnesium. 2-?.2.3 MNgneslum Alloys MlL-HDBK-693 provides a comprehensive discussion of magnesium alloys and their properties, and also describes design, fabrication, and performance data. Numerous Military and Federal Spccifications covering specific shapes, forms, and p ,,:esses also are summarized. I ,e outstanding characteristic of magnesium is its igl.i weight. This is important in helicopter design,. ,,-•. payvload ratio is a direct function of vehicle ,.,h Magnesium is two-thirds as hevy as atumi..

The most easily machined aluminum alloy is 2011T3, referred to as the frce-cutting alloy. In general. alloys containing copper, zinc, and magnesium as the principal added constituents are machined the most readily. Wrought alloys that have been heat-trcated have fair to good machining qualities. The welding of many aluminum alloys is common, practice because it is fast, easy, and relatively inexpensive. Welding is usclul especially for making ',akproof joints in thick or thin metal, and the r,. - ess c€n be employed w;th either cast or wrought. :,num or with a combination of both. The re' -:;.i. low melting point, the high thermal conductiviy , the high thermal expansion pose problems. •., heating is necessary when welding heavy secti,:. otherwise, the mass of the parent metal will cond,. the heat away too rapidly for effective welding. , rapid welding process is preferred in order to mini-

rm,;w

.1one-fourth as heavy as steel. The low densi• tI effective in relatively thick castings, where ,s :;r -sed rigidity of magnesium is an additional hcnefit or this reason magnesium is used freq••iv., . main rotor gearboxes, motor trans,,:..sings, and many other load-bearing :ilphpii .' .mn helicopters. Most of the helicopter pm C .: . have several hundred pounds of maincsiw, . construction. The ,--.: (American Society for Testing Mcr •1 .-: '•it nclature system is used exclusively in u',: , . ,..,sium alloys. In this sytem, the first ,t I. . .1' the principal alloy elements, while th, n -. ,ate the rcspectivi, percentages. I'. ., .... aluminum, F i -. earth, H •,, . .:.,, ;ým. 1. lithium, M manganese, Q .iye I . -,o ,.'inc. By this designation, AZ91C - Co..i .. i.ioy of magnesium containing 9% i\

-.

..

"'*

'A''

j'. "'

`

AMCP 706-2012 r

.luriiinum. V; zinc, and having a "C" variation.

The heat-treat and temper designations for magnesium virtually are identical to those for aluminum. -:1 he temp~er designations used urc those in ASTM 8296. There are four groups of mugnesium casting alloys. The Mg-A and Mg-Z binary systems are tiesigned Iir use at temperatures belo%% 300'1[ and are ~~iof Ios~er cost. The Mg-F and the Mg-Fl binary s~stents are designed for good strength in the 500'0 8004:~ range. The choice of casting composition is dictated largely by certain features of the design, and by cost and irecthod of production. For magnesium alloys, the important casting proce~ses are sand, per. mnciaret miold, and die. The choice of a casting proc-ess depends upon the size, shape, and minimum seclion thickness of the part, and upon the tolerances, types of surface finish, number of pieces to be produced. and relative cost of finishing the part. Magnesium alloys, both cast and wrought. haie outstanding macmiinability. Greater depths of cut and higher cutting rates can be used with these metals than with other structurai metals. Magnesium does slt

?a

%%,U .

s

U

94

ns

;ý

siderahh higther tenmperatures than aluminum often gise it advantages bor partic:ular applications, as in the hot structures and exhaust ducting for helicopter power s)-steins. Indeed. increased paj loads resulting front %% eight saving~i catmr ~a fseth initial costs, and in the long run titanium may prove less costls for seii plctosta okrpie mtaterials.spih aplcto hnlserrid 1Itnmisaihlt mlc ilsivrou srought shapes and in at side range of' alloyed ane unalloyed grades including billet. bar, extrusions, plate. iheet, and tuh.*ng. The mill products can be grouped into three categories according to the predomninant phase in their inicrostruclure: Alpha, Alpha-Beta. and Beta titanium. There is no single acceptt'd system for the designation or classification of titanium and its alloys as there ure for other metals. Titanium iactual", is easier to machine than the stainless steels becauhe the effects of work -harden ing are far less pronounced. Titaniumi requires low shearing forces, and is not noi-.h-sensitise. Because of these properties, it can he machined to extremely low micro-inch finishes. On the other band, the sharp

,*M4

-

-

-

the shearing point to heat rapidly. At elevated temperatures tItnmtedtoislvayhig within contact, and the cutting tool is dulled revdily. F~urther. the cirhides and oxides oin the forged pie.es are extremecly abrasive to tou's and miust be removed by tiltric-hydrolfluoric acid treatment prior to machining. Osciall, considerable kno%%hoA is required for the economical machining of titanium. Titanium assemblies are joined by spot, scant, flash, and pressure welding technique-.. In fusion welding, the TIG process is used-. heavy welding also requires Incrt g.. k.

other metals, and welding of magnesiam to magnesium can be accomplishmed reliably only by a skilled operator. The metal also cannot be soldered. properly. Thus, electron beam (EB) welding is the mostsatsfator wedin prces, athogh luxdip mostsatsfator wedin prces, athogh iuxdip brazing also may be used; care must be employed in removing all of the flux because of the danger of corrosion. The best method ofjoir .ig magnesium in thin sectonsis ahesve ondng.mal y

welding is quite saiLffactory. 2-2.2.5 Copper and Copper Alloys

2-2.2.4 'Titansium Alloys MIL-HDIIK-697 contains a comprehensive description of titanium alloys arnd their properties, and discusses design, fabrication, and performance. In addition, seven Military Specifications for specific forms of titanium will be found in Refs. 2 and 3. Although titanium is relatively costly, its high strength-to-weight ratio, excellent corrosion. resistance, and capability of performing at con-

tageous in inserts, studs, bushings. etc., where owk load is desired. Beryllium copper is useful for sprink's and oiler applications where its good modulus, hardness, fatigue resistance, and ease of formiing are advantageous. tlo%%e~er, copper alloys are the hcaviest of the common structural metals, and, therefore, have at %%eight disadvantage in aiiborne applications. The various types of copper and its alloys are

2-6

.

.

I! 1* A

Ionn. C;11;%-*

polish the material in order to obtain an. extremnely fine finish. The chips from machining readily clear ihe work and the tools. Because of its position in the electromiotive series, magnesium is subject to more corrosion than are the other structural metals. The many corrosion problerns associated with the use of magnesium severely limit its use in rotary aircraft. Magnesium alloys shall not be used for parts that are not readily accessible for inspection, application of protective finish, and replacement. magnc.ýiuin cannot be wcided satisiucturily to

A 1;oniprehensivc discussion of coppeF and copper alloys and their prtop-rties. design and fabrication characteristics, and design and performance data is contained in MIL-l-ll)BK-698. h aiu om fcpe n otraly have found only limlited use in helicopters Their therand electrical conductivity properties are ad.-an-urct

.

tl -A

1

*I

AMCP 705-202

hetter known hy namec than by code number 1 lie' lerin copper is used when the material cxceds 99.4'4 purnty. The principal alloying agent of bra... is zinc, v.hile tin isthe principal alloying agnin in bronie. Thc beryllium coppers have sinall percentajges of herylhum,. producing

u remarkably hard. hiph-modulus.

high-strength, inonsparking mat erial. The copper and copp~r alloys arc ca.st readily in all of the various castling processes. The allo.ýs are coldformed catsly, and are capahleof being rolled, drawn. spun, and flanged. In hot-%orking they arc rolled, cxtruded. pierced, and forged. The machinability ofcopper alloys is excellent. For sand castings, low speeds and coarse feed-- are used for remosing the scale in order ito increast tool life. It is better ito remove the scale by sand blasting and pickling, The copper alloys are %elded readily by all the sselding processes, although their high thermal conductivit) i%a problemi. The,, arc adaptable ito brating and are the eiasiest of ;ill metals to solder. 2-2. FL:cTOI.Y~c (1IN OFfl~;SIII2-2.3 MEITALST AmeO O 1SM1Dissimilar metals, as defined in MIL-STD.889. should not be used together in helicopter applicathc m ating surfaces are insu lated adetinis unless When tape is used between two dissimilar . qt\ quately. S metals, such as in the mounting of a magnesium gearbox toa~n adurnintm airframe, the contractor miust insure that there will be no loss oi- mounting torque as a result of normal usage and vibrations. Metali can be grouped in four categories, as shown in Table 2-3 Metals grouped in any one of the calegorits in Table 2-3 can rye considered similar to one another, while those metals placed in different groups should be considered dissimilar to one -unoher. The categorization does not apply to fasteners - such as rivets, bolts, nuts, and washers -- that are comrponent parts of assemblies and usually are painted prior ito being used. Instead, the metels referred to arc surface mnetals. F-or example, zinc covers all zinc parts, ineluding castings and zinc-coated paits. TABLE 2-3 GROUPING OF METALS AND ALLOYS (M IL-STD-889)

~~McGraw-tfill,

GFUU' 5I'M AD I Al UM'I.1UKIAt OYS !.",2 Ei`,t, 5351, (;ROUI' 11

IlI 0i~lDUI' GRUPI

6061 AND 1410

IllR At 101 t'ADMIiIM PNI . ANO All A'.~lI AND (INCIIDIND THElAIDIMItRIIK All CIYSIrl GllILIr' 1,

IRN I EAC!).ANU TINANDLHIR~l All (,i lEXlCu'il AIlNI 1SS SIFELS) C01111. I'll MILIM NICKEL Sit VIER(,OLD. P1MINUM TITANIUM. COBA, AND 1`4ODIlUM AND 1110111 AllIOYS SIAINI.LSS STEEl S AtOL itlAP111fI

______

Par. 2-2.2.2 details the use of alumninum alloys in 11clicopter construction. Aluminum alloys used in helicoýptcrs may contain copper ot zinc as an cssent ial const ituent. In sheet form, thesecalloys rcsusceptihlc ito corrosive action resulting in a loss of strength of the material, which becomes brittle withoiut evidence of surface change. Aluminum alloys containing magnecsium, magnesium and silicon. and chomniumn as the essential alloying constituents arc mnuch more stable under prolonged weathering condition% than are the aluminum alloy% containing copper or zun::.4 Cadmium behaves similarly to zinc as a coating mectal hi.-affording electrochemical piotection of ferroins%metals against corrosion. Cadmium plating thus can he: used to put ferrous metals into the same group a%5 aluminum al~oys. giving them a similarity. A dctailed ditscutsion of coating processes can he found in p~ar- 2-tv. In general, when two dissimilar metal surfaces come in contact with one anothcr, a corrosive action called galvanic action can take place. Coating iitals are used as thin layers between dissimilar mietals to prevent this type of corrosion. Table 2-4 illustrates the position of metals with ~.I r fth V lack of su ce ti.lt i.h__r red t to galvanic action.

2-3

1

NONMETALLIC MATERIALS

2-3.1 GENERAL This paragraph discusses the applications of the thermoplasstic and thermosetting plastics, elastomners, fabrics, and transparent maiteials. Other materials - such as glass in light bulbs or optical piping, ceramics and mica in electrical insulation, and cairbon and graphite in lubrication or electrical contacts - also play significant roles in helicopter construction. The nonmetallic materials used in composite structures, reinforced plastics. and other composite mnaterial% are treated in par. 2-4; plastic materials used as sealants and adhesive-; arc covered in par. 2-5. Comprehensive discussions and detu~led design data will be found iii existing documents. Among these are: MIL-H-DBK-700, MIL-HDBK-l7, the Modern Plastics Encyclopedia, published annually by and the Malerials -Selector Isjxuv, published annually by Reinhold Publishing Co. The major disadvantage of plastic's is their low moduitis. which is in the order of a few hundred thousand psi compared to 10 or mome million psi for metals. Thiey also arc more- sensitive to heat, soften-

ing markedly at 400~0 F and below. On the other hand.

plastics can he as strong as steel, be lighter than magnesiurn, aind have better abrasion resistance than

~2-7

.

I

4X metals. Normal corrosion is not a problem. Although they are nonconductors for electricity and poor conductors for heat. they can be exceedingly tough and wear-resistant, and can be fabricated in a variety of waya. When judiciously selected and pioperly used, they often can p.rform better at lower cost than any other material. TABLE24 POSITION OF METALS IN THE GALVAN'C SERIFS

MAGNESIUM MAGNESIUM ALLOY

shapes. plate. sheet, and film and in a wide range of shapes. sizes, and thicknesses. The materials are machined, or shaped readily by thermoforming proccsscs. Many items may be purchased in thc finished form as produced by extrusion or injrction molding: included arc screws, nuts. bolts, inserts, grommets, straps. pins, knobs, handles. instrument facings. housings, boxes, conduits. electrical receptacles, covers, rails, runners, guides, snaps, and slides. Many of these items are supplied as off-the-shelf inventories in a variety of sizes. Nylons and polycarbonates are known as the enginecring plastics. Nylon, a polyimide. has high strength and high elongation, giving it a toughness that many applications depend upon. It has high

ZINC ALUMINUM 1100

modulus in flexure, good impact strength, a low coefficient of friction, and high abrasion resistadcC, as

CADMIUM ALUMINUM 2017 STEEL OR IRON CAST IRON LEAD-TIN SOLDERS

well as good fatigue resistance under vibration conditions. Its primary disadvantages, though not significant, are dimensional change with moisture absorption, and the need for incorporating carbon black in order to )rotect against ultraviolet degradation in outdoor use. Nylon is used in gears, arms and other contact applications, and in pressure

CORRODED END ANODIC.1

SLE)

•

LEAD

:I

TIN -•-;T

.

;

tubing, belting, and we.ar pads. The polycarbonates are aromatic esters of car-

BRONZE

COPPER-NICKEL ALLOYS TITANIUM MONEL SILVER SOLDER NICKEL INCONEL CHROMIUM-IRON 18-8 STAINLESS 18-8-3 STAINLESS *.SILVER

,• L

GRAPHITE GOLD PLATINUM PLATINUM

______

PROTECTED END (CATHODIC]

OR MOST NOBLE)

SOURCE: REFERENCE DATA FOR RADIO ENGINEERS

FEDERAL TELEFHONE & RADIO Co. 3RD

2-3.2 THERMOPLASTIC MATERIALS Thermoplastic materials are those that soften when heated and harden when cooled. Typical of the thermoplazic family are the polyvinyls, acrylics, nylons, polycarbon tes, aitd fluorocarbons- Often, these have linear micromolecular structures. Products of these materials usually are formed by extrusion or by inJection molding, and they are available for manufacturing in the form of rods, tubes, contoured 2-8

boric acid. They have excellent rigidity and toughness. high impact strength, and low water absorption. They are stable dimensionally under a wide range of conditions, are cieep-resistant, and are transparent and stable in sunlight. Probably their major deficiency is that their fatigue resistance is lower than is desirable. Polycarbonates are used in shields, lenses, ammunition chutes, knobs, handles, etc. The acrylic of interest here is polymethylmethacrylate, better known as Plexiglas. This plastic has crystal clarity, outstanding weatherability in optical p, operties and appearance, dimensional stability, good impact resistance, and a low water absorption

.4.

rate. Its major deficiency is its low resistance to

'-

scratching. Its major use is as window glazing and for such applications

as transparent aircraft covers;

covers for signal lights, where its ease of coloring is advantageous; and in other optical and instrumentation applications. Its use as a window material is discussed in par. 2-3.5. For helicopters, the polyvinyls are used largely in the form of sheeting simulating leather or upholstery fabric. These are very tough and wear-resistant. In the transparent form, they are used to make pockets and holders for documents and maps. The fluorocarbon poymers have excellent thermal stability at continuous temperatures of 400°-550*F.

*

.-

-

AMCP 706-202

)

They virtually are inert to chemical attack, have exccllent damping properties, and have outstanding electrical characteristics. such as high dielectric strength, low dissipation factor and radio frequency (RF) transparency They arc used widely in microwave components and high-frequency connectors, as well as in wire coatings, gaskets, and electrical tcr•minals. 2-3.3 THERMOSETTING MATERIALS H M TT MT ASfusible. A lthou!h there is a great diversity in the chemical mak:up of thermosetting resins, they have one chaIacteristic in common: once they are cross-linked, tl-ey do not soften undoi heat and cannot be formed by thermofotming processes. With the application of heat, thermosetting resins undergo a series of changes that are irreversible. The polymeritation reaction that occurs results in such a high degree of crosslinking that the cured product essentially is one molecule. In many cases, this results in a highly rigid molecule of good thermal stability. The thermosetting rcsirts usually are used with fillers and rcinforcement. Three of the most widel. used of these materials are the epoxy. phenolic. anid polyc~tei iesius. Thcsc are employed extensively with Fiberglas fabric. %%ith chopped fiber in laminates, in sprayed forms, in filamnent-wound structures, in honeycomb sandwich structures, and in combinations with halsa wood or formed shapes. The epoxy resins arc based upon the reactivity of the epoxide ktroup and generally are produced from bisphenol-A and cpichloiohydrin. Epoxies have a broad capability for blending properties through resin systems, fillers, and additives. Formulations can ne so.i armnex1t.'01c or F-11,U T'are.. ... able as prepolyrners for final polymeritation in the form of powNders and liquids with a %ide range of viscosities, Some cure at room temperature, while others require curing at elevated temperatures. The powders may be transfcr-qilded byi machine, and the liqiids may he cast. Mure-oftcn, the liquid is used to inipregnate materials for bonding. The outstanding characteristic of epoxies is their capability to form a strong bond with almost an) surface. Fur this reason. they are used widely in adhesive formulations. The molded produtcts have high dimensional stability over a %ide range of temperatures and hunridities, excellent inmchanical and shock resistance, good retention of properties at 50091-, and excellen! electrical properties, The phenolics are the oldest and the least cxpensive of the plastics. 1-thermosetting he basic resin is ianufactured by weans of a rc,actiorr bet ncii

phenol and formaldehyde. Thi. resin is blended with dye, filler, and curing agents to make the molding powder, which is called the "A" stage powder. Pokkders such as these are molded for 2 min at '25'F at 15(X) psi pressure. As the granules are warmed by the hot mold, the resin melts; the material flows and fills the cavity, further reacting and going through a rubbery "B' stage. With further cross-linking it reaches the -'C" stage, at which it is hard and inFillers used in typical phenolic molding powders are fibrous in nature; their interlocking fibers act to reduce the brittleness of the cured resin. Wood flour is used most commonly, while asbestos and graphite fibers form the i.onventional heat-resistant plastics. Paper and fabric fillers are used for high impact or shock-resistant phenolics. When the powders are used for lamination or in making composite structurc-:. a solution of the resin in alcohol is used to impregnate the fabric and then "B" st.ged. The layers ot" impregnated fabric are laid together and then cured b) heat and pressure. The advantages of phenolics are their low material processing costs, dimensional stability. excellent hrc ktia fpris 1 lodb1rn teristics, and good weathering properties. They are used in electrical components, receptacles, conduits. housings, etc. IThe polyesters are plastics formed of chains produccd by repeating units of a polyacid and a polyglycol. They may be aliphatic or aromatic. Familiar forms - fibers such as Dacron synthetic fiber or Mylar film - are the so-called linear polye.;ters. More important to applications in helicopters are the thermosetting resins. These arc the three-dimensional or cro.-linked polyesters that are formed by bridging ,i unsaturated polyesters. In this form, the polyester is supplied as a syrupy liquid that - %hen mixed with a small amount of curing agent, ,pplied to a fabric, chopped Fiberglas, or filament tow, and laid over a form - rapidly reacts so as to establish a rigid structure. Such structures are of particular use for ra. domes because of their RF transmission and excellent weaaherability. They have high modulus and impact strength as well as excellent flexural and tensile properties. The polyesters may be cast in order to produce glazing materials. Another highly cross-linked family of polymers consists of the urethanes. 1 hese arc formed by the reaction ot isocyanates with esters and unsaturatcs. In the process, carbon dioxide is evolved and forms a highly porous structure. "[he stiffness ranges from soft, flexible foams to highly rigid foams. The flcxible foams are employed for cushioning and padding 2-9

.

-

1xvln

SECP 706202 .and to reduce noise and shock, as well as for thermal insulation. The rigid foams are used as light-weight stiffeners in structures.

MIL-HDBK-149 presents a comprehensive discussion of the technology of the elastomeric materials and their applications. From the standpoint of durability and performance, natural rubber remains in demand, and substantial quantities arc used in blends with SBR, butyl, and other synthetic rubbers. Natural rubber is a stcrospecific polymer of isoprene. Its applications in pneumatic tires, bumpers, shock absorbers, etc., as well as in belting, gaskets, and seals, arc well known, Substantial quantities con"tinue tc. be employed in hdicopters. Carbon black constitutes about 50% of the w,:ight of il-csc corpositions. A more advanced synthetic rubber is nt-nnr%,,e, a general-purpose synthetic made by emulsion polymerization of chloroprene. A notable characteristic ":t rabte•ri e ris i,.,o gasoline. foles, monrieating oils, and othur solvents, a d its excellent resistance to weather-oxidations, ozone, and ultraviolet light. It has goon tensile strength, tear resislance, abrasion resistance, and rebound character;stics, and excellent adherence to metal and fabrics. It provides average insulation and has excellent dicirctric strength. In helicopters, it is used to coat radomes and the leading edges of the rotors for protection against abrasion by rain and dust. It also is used in boots on other leading edges and areas where wear is a factor; and in transmission belts, hoses, liEcs sas,

&-1 ecica

apl-U----------

Another important family of clastonteric materials is the silicones, which are used in many diverse and seemingly unrelated applications. The silicones are organo-poly-siloxanes, having alternating silicon and oxygen atoms in the backbone of the chain. The silicone resins may be cast, extruded, or injectionmolded so as to form shaped products. They are available in sheet or bulk form; as a range of pastes and liquids for use as adhesives, sealants, and coatings; and as powders for foaming. They are stable continuously at temperatures from - 140' to +600 0 F, and initermittently to 70001. They are weather-resistant, hove high dielectric strength and a low dissipation factor, and are bonded easily to metals, ceramics, and plastics substrates. Aromatic solvents and chlorocoinpounds swell silicones, and they have higher gas permeability than do other rubbert. They are used as foaming agents, encapsulatin mesins, sealants, and in electrical applications, 2-10

2-3.5 WINDOW MATERIALS Gazing materials and methods of attachment are , discussed in detail in MIL-HDBK-17. That document also lists additional Military Spccification" coveting specific glazing materials, resins, cement., and proc-.sscs for the design and fabrication of widow systems. The optical properties of greatest significancefor aircraft glazing are surface reflection, index 9 'f refraction, absorption of light, and transmission of an undistorted image. The thermal properties of primary concern are the coefficient of expansion, the thermal conductivity, and the distortion temperatare. The major physical properties are density and hardness, or scratch-resistance; the major mechanical properties are tensile and compressive strength and the modulus of elasticity. The ideal glazing will be strong enough to withstand structural and operational (wind and water) loads, hard enough to remain unscratched, optically clear after a life of operatiun, unchanged by thermal loads, and unaffccted by the weather. Although no material- passess all of these desirable characteristics, there are several that perform ver) well. The three glazing materials that are employed most often arc glass, cast polyester, and cast aciylic (methylmethacrylate). Polycarbonate, an otherwise strong contender, has not yet been produced economically in large sheets with the required optical properties. Monolithic glas- is used in helicopters only when use temperatures ex,.ecd the performance t(mperatures of the laminated glasses and the poiymeric materials. The lamination of glass with plastit imt provcE. the -cs~s!_i. to- therma and rne-har.ca! stresses, and minimize; the possibility of complete failure of a panel, Splintering of the g'asi is prevented, although the load-carrying capacity of the laminated glass is less than that of plate glass. The plastic interlayer is sel,.cted so as to provide the greatest ability to absorb impact energy. Polyvinyl butyral is the most common interlayer material for both glass and plastic laminates. A new thermosetting, polyester-base, transparent sheet material has becn developed under the trade niame "Sierracin 880". It can be used for aircraft endosuren. that operate at suifitce temperatures of up to 300°F. and is characterized by its two-stage cure. After formiig and post-ctri~ag, the ultimate physical properties of this materia! are obtained. Sierracin 880 generally is used as a laminate with acrylic, and is described in MIL-P-8257. The glazing materiaS u;sed nast widely for helicopters is cast polymethylmethacrylate. In many air-

I

S" craft, it constitutes a major portion of the fuselage walls. For window applications, the stretched, modifled acrylic sheet is preferred, per MIL-P-25690. The nmodified material has slightly higher heat resistance than does heat-resistant polymethylmethacrylatc, along with better resistance to crazing and solvents. When stretched to W-I100% biaxially or multiaxially. acrylic sheets develop increased resistance to crazing, higher impact strength, and improved rcsistance to crack propagation - without detrimental effects upon their other properties except for reduced abrasion rFsistance and laminar tensile and shear strengths. The sheets may be formed thermally to diffcrent contours. Laminated plastic glazing materials are made by bonding two or more layers of acrylic or polyester plagtic sheet to a soft plastic interlayer by means of an adhesive. This process greatly improves the impact and structural strengths of the material, Laminated plastic glazing materials arc defined in MIL-1N5374. Differing thermal expansion rates of glazing materials, edge attachment mmieriais, and metal air-

frames present one of the major probicms

iM I

design of window-glazing. For all types of glazing, an edge wrap is used in order to minimize the propagation of stresses originating in cracks and chips at the edges. The preferred wrap is two or more layers of polyester (Dacron) fabric, ,,oven from twisted yarns and impitgnated and bonded in ?lace with epoxy resin. Th wrap overlaps sufficiently on the glaze material and extends sufficiently beyond the edges to absorb the stresses of aitachment closure and at the same time "o distribute the load uniformly across the window. There are many closure designs, but the prefcrred enclosure will be designed so as to hold the glaze securely in a sliding grip ir. such a manner as to allow for reciprocal longitudinal motion - as the glaze expands and contracts - while always applying a comprehensive load endwise. This may be achieved by placing a compressible, neoprene-imprcgnated tube at the bottom of the closure channel. The closure will be attached rigidly to the airframe. The contacting areas between the closure and the edge wrap will be sealed with a flexible sealant, prefe-rably one made of silicone,

-

24 COMPOSITE STRUCTURES 2-4,1 FIBERGLAS LAMINA'TES *

".

Of all the fibers available for the rcinforcement of plastics, glass is used by far the most widely. Of the various glass compositions, only two are important in "aircraftconstruction: "E" and "S" glass. "'E' glass is used extensivily; "S" glass provides greater tensile

A CF 706.a01

strength and modulus, but is considerably more cxpensive. The major advantages of glass-reinforced plastic (GRP) over isotropic structural mate•ials (primary metalsý include: I. Formability and versatility. Large complex parts and very short production runs are practical. Because there are few limitations on size, shape, and number of parts, design freedom is maximized. In addition, the reinforcement can be oriented as desired in order to increase properties ir. specific directions. 2. Chemical stability. GRP is resistant to most chemicals, and does not rust or corrode. 3. Toughness. Good impact resistance is a feature of GRF. 4. Strength-to-weight ratio. Specific strength of GRP is very high. For example. unidirectional GRP has a specific strength about five times that of the commonly used steel and aluminum alloys. 5. Insulation. GRP is a good thermal and electrical insulator, and therefore, wiil transmit radar and radio waves. Wk . ......... , ..... r .. .atched n he readily and effectively. On the other hand, GRP has certain disadantages compared to other construction materials, namely: I. Nonuniformity. Variations in material properties within a part and from part to part are inherent in most of the fabrication techniques. 2. Low modulus. Stiffness of GRP is relatively low. 3. Slow fabrication. Production rates are low in comparison with most metal-formirg operations. Thus, GRP coiitiUctmon is mnostudan'tgeou•s ar, parts with complex shapes that would be difficult to form from metal, foe parts with anisotropic strength requirements, or in applications where the conductivity or poor dent or corrosion resistance of metals present a probl'm. Sonic typical GRP applications in helicopter construction arc in canopies, covers, 'tad shrouds (for formability, specific strength, dent resistance); rotor blades (for formability, anisotropic strength, and stiffness)- control surfaces (for anisotropic properties, dent resistance, repairability); and antenna

housings (for radio frequency transparency, formability). It is conceivable that an entire helicopter airframc can be constructed from GRP, as has been

done o, several smal, fixed-wing aircraft. 2-4.1.1 Design Colsiderations Design rules and procedui ,. for reinforced plastics do not differ markedly from those for metals. Stress st:ain curves, however, are similar to those for wood 2-11

.

. ,

.

(

AreP 7062D2 (which also is a riber-reinforced composite) in that there is no yicld point. As Part I of MiL.IIDBK-17 contains conlsidcrable prapcrty data on specific materials, only generalities arc considered herce, ~ar hertw cscntal igreiens i glss. 7~reinforced plastics: glass fibers and resin. A finish, or coupling agent. that enhances adhesion between ile glass and resin usu'lly is used as a cobting on the glass, and may be considered as a third component or as part of the reinforcement. The resin system ha roe ~ detrmiingthe geneall th limtin chemicalyhathermlmiand e oletica pdopertiesiofgltm catemwicl, thertyealon, elcrclpoprtientto and of themi reinforcement predominate in determiining the basic --

mechanical poets.esters, propeties.laytips 2-4.1.2 Resin Systems

rtint

*

.

Essentially all GRP laminates arc madc with thcrmosetting re-sins that, when mixed with suitable catalysts or curing agents, arc pormanently converted to the solid state. Reinforced thermoplastics (RTP). whicn contain snort glaass ircrb. arc a rapiduy growing element of the injctlion-molaing industry: such parts. however, are not consider.-d laminates, Probably 95% of all GRP laminates are made from polyester, epoxy, or phenolic resins. For very-high. temperature service (above 50OOT), silicone (MIL-R25506 and MIL-P.25S18) and polyimide resins are available. These, however, have no known applications in current helicopter technology. 2-4.1.2.1 Polyesters These are by far the most widely used resins when thob eniirp rGP intisictry ic -nn-ijere-d. They are !ow, in cost, easily processed, and extremely versatile. Available types range from rigid to flChible: there are also grades thait are fire-retardant. ulhrav'ioletresistant, and highly chemical-resistant. The upper temperature limit for long-term operation of generalpurpose grades is 200*F, although temperature-resisresins are available that are useful up ito 5000F. These can be formulated for rapid curing at room temperature or with long pot-life for curing at elevated temperature. rhus, they commonly arc used for wet layups but prcpregs also are used frequently. Prepregs that cure by ultraviolet light also are available. Disadv:tntages of polyesters include high shrinkage * during cui-c, inhcrently tacky surfact. ii cured in the

presence of air, odor, and fire hazard in wet layup fabrication from styrene monomecr and peroxide catalysts. Requirements for general-purpose polyester larninating resins and lamuinates are contained in M IL-R7575 and L-P-383, respectively; fire-resistant resins 2-12

und lainin.,tes are detailed in MIL-R-215042 and Ml L-IP-25395. 2-4.1.2.2 Epoxies Epoxies probably arc thec resins most frequently used ior atircraft GRP laininates. Although about twice as costly as polyesters or phenolics, epoxy plctos rsn tl jeiepniefrms stapiatint. hcia-rsivefrm rechansticl, 4elietrcaend Mcaiacetiaadccia-eitn prptisreeclc.Adconomstubrae odadcuehinaeadmitraisvr sorption are low. Temperature resisiance of generalpurpose types is intermediate betweca that of polyesters and phenolics. Formulation anid fabrication oy swt psiltesaexrmly vraie

epoxies can be formulated for uses tuch an wet

or prcpregs for room-temnperature or elevatedtempei ature curing, and for fire-rctardancy. The choice of curing agent plays a maijor part in determining curing characteristics, temperature resistancc. chemical resistance, flexibility. etc. In addition, a variety of modifierF and fillers is available to provioc specific qualities. There are relatively few disadsantags-s with epoxy rcsins. However, becaube an~ine curing agents that are commonly used in room tempeature curing forlmulations mady ctiuse severe dci matitis, skin contact must be avoided. MIL-R-9300 contains requirt.ments for epox) lamninating resins, wvhile requirements for epoxy laminates are covered in MI[-P-25421. 2-4.1.2.3 Phenolics Phenolic resins are used primarily in (3RP appli-

an inrinrpnsive material

%xhprc

,

(

th heat. re-

sistance uri to 50001: and/or rionflanimability is required. Excellent elctirical propcrtics also arc obtaircti. Because water is produced and released in the curitig reaction, relatively high molding pressure is req4uired in order to prc~eni porosity in phenolic laminates. Preprcgs ticarls alk-ays are used. Mil.-R-9299 and NIIL-P-25515 cover the requiremncrts for Phenolic laminating resins and phenolic laminates, respectively. 2-4.1.3 Types of Reinforcement Gilass reinforceniem is av;,1lable in several basic form,,. anid in a %ide variety of specific construo'tions ssithirr these basic cztegories. Those formis coidnion1) j.scd in 6RP lamirates include skoven fabric, chropped fiber mat, and! i~onwoven continuous iapes or ro,*ing. Neaýrl) all of these art: dcris,,d fromn continuou% filamrents of 0.00023. 0.00028, or 0.0003h in nominal diamecter. Numecrous standard yarn C(:ii structio11s are available, with %ar)yrtg numbers (if

3

parallel Miaments per strand, strands per yard, and twists pc.- inch of the strands. Likewise. there is a multiplicity of fabrics woven from thc.ic yaifls that vary not only in type and amount of ) arns but also in thc type of weave. MIL-Y-l 140 is an excellent reference for definitions and rcquit emcnts for ti,a various yarns and woven fabrics, The type of reinforcement selctcird will depend upon the mechanical property requirements, part shape, and applicable fabrication technique as discussed in the paragraphs that follow, 2-4.1.3.1 Nonwoven. Continuous Filaments axium Thi renfoceentoffrs fom o mechniclpopeties bu ha miimumfabicaion possibilities due to the difficulty of placing and aligning the reinforcemuent in complex shapes. A big adv'antage where this type of construction is practicable is that the fibers can be oriented in proportion to the stress in anly given direction. Filament winding is the most widely used fabrication technique with nonwoven continuous filaments. This construction is covredmoecmpetey n pr.2-43.As 2-4.1.3.2 Woven Fabric This constru-ctwn proidus good mechaniv-ul properties and formability and is. theref'ore, the niost commonly used reinforcement in aircraft fabrica-

2.4.1.3.3 hopilped F"be The third commo3n form or rcinfoiroment is chopped Fiber mat, as defined by MIL-M-15617. Because thc Fibers arc short inno their orientatIon is completely random. this material is veycm formable. For the same reasons. atid also because or its high bulk - which limits the percentage of glas obtainable in a laminate - mechanical properties are lower than with roving or fabric. Continuous (swirl) strand mat is another veriation and is particularly useful for deep contours. in both types of mai. th-glass is held in place with a small amount of resin binder. Both types are available in weights ranging fromn 0.75 oz to 3 oz per ft1. corresponding to laminl. atd thicnfiesseno absou0.3in. trodue 0.00in.per naldMith rei~nfo 6icemincallso ishproduc ldin promrn-

h

1 A_4

adsetmoigcontewihrsitiss pound (SMC). Chopped fiber parts can be fabricated by the sprayup or Preform t-echniques described subsequiently.

4..

FaratoMeod

previously discussed, each of the common forms of glass reinforcement (roving, fabric, mat) can be niorchip el either dry or preimpreitnated withth ) laminatirip resin, which is cured or dried partially to a ' solid or tacky condition. The latter, called preprep, tion. When wetted with resin, the cloth has conarc advantageous in that they contain a controlled, siderable ability to stretch and conforfm to rather uniform, and readily measurable amount of resin. complcr. contours. Although intended specifically for They arc, therefore, easier to lay up, because wet laypolyester laminates, MIll-C-9084 fabric usually is up operations often arm messy and odorous. Prespecified for larniiat"s made with all resins. IKaquirepregs can be obtained with varying degree of tack so ments for eleven basic fabrics and six subtypes arc deas to suit the spwcific operation. And, because cornfined in the specification. Approximate thickness per~ pleie quality con irol tests can be mitde before the part is fabricated, the. problems of incorrect weigtding and ply ranyes from 0.003 in. for 112 fabric to 0.027 in. mixing of the resin system arc eliminated completely. for 184 fabric. (Still heavies fabrics, woven from :he( fife-, pre-pre" er In o~z rovings rather than yarns, arc avia~hbl Ii- i'll-----ses formulated with curing agents or catalysts thatI of up to 0.045 in. pcr ply-. these are covered in MILrequire heat to cure (generally 250*-3S0*F for at least C-i19663.) Most of these fabrics are balanced %%eavcs. I fir). Under heat, the resin melts initially, and then with nonoinally equal construction in the warp and fill directions: 181 fabric at 0.009 Mi.per ply is the coiiverts chemiically to a thermoset solid. Some pressure almost always is required in prepreg lamistandard balanced fabric upon which most test laminating in order to maintain good contact betwtien nate- and published properties data are based. Repireplies of reinforcement. This pressure results in reater cm, ing the extr,:me of unbalance is 143 fabric. %hich resin flow and, consequently, in higher glass ratios has a warp strungth about 10 times as great as its fill strength. This approaches the nonwo%,en conarid better mechanical properties than are obtained with unpressurizcd wet-layup laminates. struction described previously, sacrificing sonic mechanical properties for improved drapability. Generally, epoxy resins and preprcgs of roving. Most high-strength, glIAss- fabric- based laminates Lre tape, or fabric are associated with components of made from 181 and/or 143 abrics. higher quahlt, cost, and strength, while polyester As with MIL-C-9094, MIL-Y-l 140 originally swas resins and wet-lay up (or SMC) processing of fabric or ) interldeo lor polyester laniinates. However. its renina arc used where maximum required properties do S quliremeflts also usually are specified for fabrics conri)t justif*y the increased costs. Those fabrication taining epoxy compatible finishes. mecthods applicable to construction of laminated *.o

n'

.es-m-l

2-13 u

F

.

W, v

2

j

AMCP W16-202

r

TABILF 2-5 PROCESS COMPARISON ;Ii)1 F('R (;R! lANIINAl -:S*

,

-

, O

POLYESTER EPOXY

CONTACT •IOLOING __

_

VACUUM

.

.

_

BAG

"OPEN MOLD

._

I

__,-

SPRAYUP

_'_

_

POLYESIER

V•A1 FABRIC

30 45

MAT

40

T

I" L VI? R S•, E

RL.

70 I011

70101I10

)

60 10O 220 55 455ONLY 60 70 TO 2,0

POLYESER EPOXY

CONIlOU ROV'NG

I

0

_

I

EPOXY

CHOPPED

30

50 1000 1

70 TO 110

p•,,1

CLOSED MOLD MATCHED-DIE MOLDING S~EPOXY

MAT

FABRIC

PHENOLIC,

MAT

"POLYESTER,

FAPI:,CS, PRI'REG

EPOXY

LIMI ILI? DY SiZL AU I)CLM L Oi AUIOCLAVL 00IOL

100 10300

BYI _

_

3'i

, 4 1• .

••jo

Of

FRVSAEY.QN,1 FR'..V5AEEYHELIEISI TO 1CIt BOATIHULS_ t S-

1,.?..

MELAMINE, SI,1N1 EPOXY[

PHLNOLIC, ,ELAVINE, SILICONE,

__IOW

L'11

_

22510300

CE POA! HUHl •.

50

ER PREFORM I,• EPOXY •i-

10 ()AIL

P

55SKBM _

ROIIUUIS Si"L 01 I'0

0

5 ROLYESTER

FABRIC AOE N ROV ýOVE ,MA 1 MA FABRiK EPOVXY R00 POY WOVEN ROV

Epoxy

••POLYISI

.

j

TRr

40

OVEN ROY

EPOXYFARI__ POY V.OVEN RO\'

PRESSURE

HAND LAVUP BAG .,...AICAEPOLYESTER • -': AUTOCLAVE

:-

_

A

f -IBI RGLAS I'E RR[ F 1.r, WERGA R61 I

RESIN

PROCESS

•

R,

SP.

L1

'25 10 350

100 10 3000 1 ' 6 .ANE f't, NLI S

2Vr'TO 350

IN

O

_

1O

{ "1

ANE LS•'U 1C5

AND NONV"OVEN

_j _._

FROM OWENS CORNING FIBERGLAS CORPORATION TECHNICAL BULLETIN I-PL-1998-B

GRP helicopter components arc jiscu::scd subsequently. General guides to moldinj; processies. and resultin, laminate properties, are shown in Tablei 2-5 and 2-6. respectively. 241.4.1

Opcm Mold Hand Layup

simoothed onto the exposed surfac of the layup in order to provide a better finish. 2. Vacuum bag. A film (usually poly,,inyl alcohoIlI er nylon, is phlced over the surface of the part and

scahLd at the edges, or the entire mold is pk.ccd in a bag. A vacuum k then drawn. resuittlg in the applica.ion of atmospheric pressure to thr"lamfinate. I:ven This method consists simply of placing the rethis re~ativ,:'; low piessurc (15 psi) conside•Fabl) intquird number of plies of reinforcement and resin proves th- laminat,: quaiity bh reducing entrapped over a single mold surface, and rubbing or rolling out air and resil|-ricin areas. the air. Curing then is accomplished by one oi the 3. Pressure b:i. In this case a rubber film (o!'Ten "following processes: C t l T n i wconioured to 117c part shape) is placed ovz-i dit- l, tip 1I.Contact molding. The laminale is allowed to -and thŽ mold is scaied snth a pressure plate. Air or cure without the application of pressure. usually at stvam pressure ol up to about 100 psu then is applied room temperature. Heat can be applied to accelerate to thl. cavitN. the cure, but the contact process usually is employed 4. Autoclase. In this variation: of the pressuwc t'tg for large parts and/or short pruduction runs. and process, the ent~re assembly (mold, layup, rubber "~'.•,both heat and pressure 11ay be impracticabic. A stripfilm) is placed in a steam autocime 'nd cLured, norpabla film, such ab cellophane, sometimes is malhY at about 50-100 i'. 2-14

1 '

706-202

___________________________________AMCP

TABLF 2-6 (,FNI-IT.A1. PROPER} IFS OBTAINABLIA IN (&ASS RHFN1FOR( I) I'l.AS'I(*S* POLYESTER

EPOXY

PHENOLIC

j(-LASS fMM. FPL FCR11l

GLASS MATT

rROPF RiY

OH SHEET MOL iINi3

CLASS CLOl H

GLASS MAT

GLASS CLOTH

SP~lCI l'CfRAVI] Y

1.315 -10 2.30

1*012.0

1.810O ?.C

1.910O 2.0

1.70 TO 1.95

CLASS CON EN1,'. BY W H TENSUL STRUNCJ H,

251TO15

460T057

65 TO 70

41 TI 1

15000G 10 25,000

30,000T0 ?0,0C0 14,OOOTO 30,0%0 ?0,000TO 60,000 4,.000 To 60.000

COMPRESSIVE STR LNLI 11, 1

15.o000vO 50.000

?5,OOOTO 50,000 130,000!0 38,00T130,00010 70,030 17,000 TO 40,000

FLEXURAL SI R[ N67. H,

25,000 10140,000

146.0 00 7TO 90,000 20.0 0T1 26.000 70,30010 100,0 00 10,0 00T . 95,0001

i475

6.5X0,5 il. NO1ICHED HAR? ot -!1 in .O[ NOTCH)l VWA[ER ARSORP N ,i Z411i Lhil. HICKNESS, '

81

TO030

05

-

BURNING RAIL VOL AýITIiTT 50. RH AND 73" F,owm. 'cnOi

U1.1i10i.6 uU -

-

.Q

-.

3(10 T-3 15

-

-

10-

04

11

120 TO 180

60 10 120

fU

33010O 500 ____S--*--

.8x 1 0 3B

125 10 140

iIlU. _

____I___-__

O--1EXTHIGUIS11ING

'IS-

.;~U. U.UjI,

.:

-

3-N TO 3

lol

8 TO035

1 TO 26

8 TO 15 rrnnrt

o-n-

I

COI IE~ WNUOUS

ARC RESISTANCE,

O

330 TO 500

I-~..'-4 .XC _

5

___l"

100 TO 110

_

_

350 TO 500

-F-NONE XI 71

1

20 TO 150

'CC1MP!LFD 1Y RF INFORCED PLASTICS, COM~POSITES DIV.SOCIE3Y OF THE PLASiIV. IN('USTRY INC 2-4.1.4.2 Spi'nyup Ili this nethod. continuous roving is chopped into I1-to 2-In.lengths and blown, into a spraying Ntrearrn of resin and catalyst (hat is directed ugainst the mold (A fast roorn-tcrnilirature-setting polyestef generally Is used.) 1fhe mix~ture is land rolled to reduce air and icvec the surface. The resulting part is similar i11 con%.tructioii !o a chopped fib~er mnat hand lay up.

peratures to 350*0 F commnonly arc used. Prepreg. fixbvics and tapes usually are spr-ciied for aircraft applications requiring maxinium strength-to-weight ratios. I owcver, fo'r cornpkA) shapes and volume production. choppeci glass preforms (held together, like mat. b) a sinalI arniunt oý resin binder) frequently -ire used. 2-4.1.5 Surface Finishes

Althugh(hi C-1s i vey . uficentforlare cm-

-or man) applications, a smooth surface, free from aii pockiets and exposed glass fibers, is required. Such a surface may be specified in order to improve weathering characteristics, material-handling capabilities, human contact applications, or. 6riiplv, appearance. There are three different methods used to obtan a smooth, resin-rich laminate surface: 1. Veil mats. Thes.- consist of loose, nonwovf-n mats ot glass or synthetic fibers. THickness may range

poricots. it has limited applicastion in aircraft constructiooi due to pour Uniformity o! thickness and h-o-wightratos.outdoor lw relaivey s~ 2-4.1.41.3 %latched Die Molding

Whenever -'4)sely controlled thicknecsses are re-

quircd. two miold halves ar±. necessary Matched dit mrolding also is practical for high- volunic production even where high-quality surfaces and close tolerances are not required. Pressures of up to 300 psi and tern-

1W

70fi

inav be customn rnoldicd b) procedures sintilar to those dcsc:rihcd in par. 2-4.1 for low pressure, closeddie lamina.te, Industrial lamnizates are used for components of simple geomectry rcqUI.ing internmediate strength, lightweigh~t, and nonmetallic characteristics. In hell. copter construction, they frequently arc used for wear surfaces, such as oil conduits and lpuleIys for control cables anad iii electrical circuit boards. industrial laminates can be made with a number of' chemical, thermial, and electrical properties by varying the ty'pe and ratio of resins and reinforcemerits. Those combinations that presently are avsail~tblc coniinercially art; sho~ i in Table 2-7. In cach case, the laminates are manuifictured by stacking Lip. sheets of the irnpregnawed reinforcement (or by vrapping, in the cas;: of' itbes or rods) and curing tHicn undcr hecat arnd pressure. Vcry high pressure rainging from about 200) to 2500 psi used, resUlting wnhigh-quality, void-free parts. Fromi the biisic: moinbinations shown in Table 2-7 niore than 70 standard grajdes of laminates are derived. of these, 32 grade,, are classified by the

fToni 0.001 to 0.030 in. They are so loosely consiructed that resin content in the veil are.; is thbout 85 by weight. 2. Gel coati. This technique conlsists of spr'i,coating the mold surface with 0.0i0- to 0.020-in. layer of thixotropic (nonsaggiiig) resin, which is allowed to sct ayig rio uto theglas rinfocemnt. Resiiicint resins usually are used so as to provi~ie a compromise between scratch resistance ::nd impaL-t strength. Most gel coats are polyesters. bu! th,.methd an e usd aso wth poxes.mechanical. 3. Thermoplastic fiims. This method consistIs of laminating a film or sheet of weazhcr-rcsist::iit and/or decorative plastic, such as polyin)! fluoride or acrylic. to the GRP surface. This technique should be applicable to a vai jety of GRP processing methods with both polv-estci and epioxy rc:,ins. but it has not been used widely in the past. Re-cently. however, .i process involving vacuum-forniing of thermoplastic sheccis -which then are reinforced by %praying the bacck side with chopped glass and polyester resill has found wide acceptance. especially in the nianufacture of large parts (up to 300 ft-). FABRIC

-are

-A.MiNA US..

-

I

-....

A.cs

!at

Industrial laminates, also cýalled high-pressure laminates, are reinforced p~astics that are maimifactuied in standard. simple shape% such as sliccts. rods, and tubes. F-abrication of prits For materials generally is accomplf~ished h) standaird

(NI NI.\). Descriptioois of ihe 1NWMA grades and thenr applicatiois are contained in Vol. 46. Moderni

mtetalorking operations, sicias cutting. drilling.

po,- design. (icuicral characteristics lesiltinc' Iront the selection of the various reiniforcemeints and resins art: descrihed sibsequciltl).

punch~ing, and machining. In conrarst1

tO

'th

Lcdpi.asloaetepretesf

thcse laiminaites. The designer should consider grades. app~lication. aiid producihility prior to final corn-

Molding it)

the desired shape as discussed in par. 2-4 1- Where prooneicon quantities \5arrant inold costs., pairts aulso

I ABLF 2-7 F

RE SIN I YFF

REINFORCLE?'N1

PH tjOIL

FCR%*At D[[

Sh1.LI

Iý h

luf

HAI~YPE

C it-

I L0-I

I

CC)MTN FAWiC AS[BES 10 PAIq R

I

K:

~ ~

'1

I

0,1

S E. HE KO

iii. I

týcý

S:,K01,11 4-

C[

%tIL

I I TL112

I:OLlI

.

~

1~

-

ASBESTOS FAhkICI

NYLON FABFiC ..

1III

AND MAT

'FROM1 MiODRE N PL AST H Eii'C

L i zz

1

*

MATE RIALIS i-.S,'iPr., UJ.iLJ F L~[ ';I'. i.....

2-16

iIF

.II.'NS

NHI

-

AMCP 7W202

-

-.. ) .2-4.2.1

Reinforc4.ment Sdelltion

ParaLI'.el) low fur glass laminates, and c.ost is high.

laminates art paper, cotton. nylon, glass, and iisbcstos. Attribute%of these materials are: 1. Paper. I hie leas~t expensive, and adequate I-or many purposes. Kraft paper has rclatively long fibers and is thc strongest type. Alpha ciellulosc offers improvtd e~ectrical properties. machinability. arid uniforminiy, while rag paper laminates have the lowest water absorptiort and intermediate strength. 2. Cotton. Better impact arnd compressive .strengths than paper, and most grades are only slighily more costly than paper. Electrical characteristics. howevcr, generally are not as good. 1 he heavier fabrics have the be:st mechanicial properties, while the fine weaves have good miachinability. 3. Nylon. Low moisture absorption anid excellent impact strength and electrical properties, as well a,. good resistance to chemicals and abrasion. l-owevei. nylon laminates have relatively poor creep resistance at elevated temperatures and are coniparatively expenisive. 4. Glass. Highest mechanical .stre~igths by lar.

5. NKllainiiirc Excellent arc resistance at moderate cost. NMclhanicul priopertits, and heat, nlame, and -cheminia;l resibtancc qualities, also arc good.

Thcse materials also have superior electrical propertics~ and he-,m reskistnce. Cost is relatively hiigh. 5. Asbestosi. Used in the form of paper, gnaL, and

2-4.2.41 Speeifirstions lit addition to the NEMA Standards (Pub. No. I-I 1-19,65). thei foliossing M~ilitary and Fedcral Spccifi

~/ fabri-c. These laminates have excellent heat, flamne, chemical, and abrasion resistance. Costs range front low%foi those sith a paper base to high for a fabric base. The designer should select the type of miaterial best suited for the application. considering interface recIuirinlents mind felmt~bihty,. maintainmbility, pro'Jucibilitq, and surv.ivability.

2-4.2.2

Rvsir Setewicin

The rcsins ase6 in the manufacture of industrial ipihrdi.,. 1 Jui: c. .1-o!lyestcf. S!!(voome, and

-.

kRoomn temipcature nwcchaniCAl properties are con;-

The common reinfsoricenents weda in lrigh-pFC~Surc

meLAniinv. The characteristics of each arc: I. ieot he;~li idl~use Te b lr.ihe

resins are iriexptensise and have adequate reloperties (inccharotcal, electrical, therniul, and chenical) for inanN dcsian apiiin.irmlnrced 2.hpoxý tUsed csp"ejial v% here tcgh rcsistam~c to chemnicals and moisture is rcquired. Mechanical properties andi dinicn.,.ional stability also arc superior,. 3. 'olye~tei . Less coimion. but use~d 1icrnechainjCIIAJr1ld electrical a m.tn.espeeialls %%, e flame rcsist1tiiee is a1rcqu!(cnicirt. 4- SilcOiUC. Used p~rima1ri ld,wt glass fahm ic \Ahere iicelt resi!m-!icc ito S001~ i-- rc'jimrcd. Arc: esmmalirCc is1 v\Menlemt and 11(1i1roiv arbsiorpt ion is loss. 1 lie \c r lIs mA ivssipamt lol laciot ol the-se rc~smitv at higrh ItieqUefl~cs i, utiliWc rm radMi and r-±dio ins~ulator%.

:~

~~

2-4.2.3 Speckif Types In .ddition to the materials listed previously, there arc two special typq-s of industrial laminates that deserve mention: 1. Postforming grades. Made from resins that. altltough thermoset, will soften eteough at elevated temperatures to allow the material to be molded into intricate s!ýtpes. Sptcial paper or fabric reinforcenieit also is used, permitting considerable stretchilig

2. Clad laminates. Clad, on one or both surfaces. with a %ariety of materials, including aluminum, coppu r. stainless steel, silver, magnesium. and various rubbeis. "lhe corticr-clad laminates (gerierally glass/epoxy) are used widely as printed circuit boards.

cajions arc applicablc to fabric larninates.: I. ILP-S09, for sheets, rods, andi tubes of various resins arid reinforcements 2. MIL-P-79, for rods anu tubc-ý of paper/phenolie, cotton/phCF.olrC, and glass/melamiinc 3. L-P.5 13. for paper/phenoltc sheet 4. MIlt-P-I 5035. for cotton/'phenolic sheet S. MIL.-P-W824.

for cotton/phenolic sheet for

Aatcr- or greasc-lubricated beairings b. MIL.P-I5031. for glass/mclamnine sheetc 7. M: i -P-l15047, for nylon /phenolic sheet. 2-4.3

FILAMENT COMPOSITION

Thi., paragraph is concerned primarily with high-

pciforinance composites. consistinig of plastics rewith nonw-ovn filarnints of glass. boron. 11nd high-modulus graphite. Because the fibers areV itonisso-efl and, usually, kintwisted. they can be packed to high lither loatdings. The fibers (-an be Oriented .mloiig the axes of stress in p-rcrorticod to deSAgri requirements. allowing efficient ut iti~ation of the outistaniing properties of this type of rcinforcement. Whic~i the spccilic strength tiensile strengthv-to-densit) ratio) anid

specific modulus (Younig' modulus-

to-dcnsity ratio) of tire metal alloys (alumninum, steel. ialil

corumoril) used in aircraft construction are

cornilpared, it isýshow ii that they are necarly equal to 7~ to 9 X 10'!n. arid 100 to 110 X 10, in., respwcively.

I c

-,.

-

.1

W-Ak

Although S-glass offers a substantial improvement in botO quantities, niore important is the comparalively recent introduction of the exotic fibers. with boron and &;rapbitebeing of primary commercial interest. Unidirectional composites made from these fibers have specific moduli in the 600 to $00X 10'N in. range. Thc specific strength of boron composites is comparable to that of glass. while graphite cornposites are somewhat lower in this property.

2-4.3.1 Types of Relmorcemevit A summary of the properties of the previously mentioned filaments that are used in reinforced cornposites ~s contained in Table 2-8. The derivations and characteristics of these fibers arc discussed subsequantly. TABLE 2-9 TYPICAL VALUES OF PHYSICAL AND MECHANICAL CHARACTERISTICS O3F REINFORCEMENT FIBERS Lsw

MATERIA

Ius I -GLASS

0O92

S-CLASS

0.0oý

MEIA DEST

97--

t-LS OD9

'NTPGTN)0.093 GRAMlPTI0.6

GRPIE400 (PA~,

*into

EC11

Tj0J

5ID 50

-11

7.4 92

85

14 13

5-2,

50)

5.4

60.0

650

300

4.9

50.0

820

Ui±.

40.0..

-

--

fYE

.~. 635

MIHTS-904

24.3.1.1 F-glass This glass was developed originally for its superioc electrical properties. Glass roving is manufactured by drawing the molten glass thiough resist ancc-h.-ated platinum bashings at about 2400*F. From 51 to 408 (usually 204) filaments are gathered into a single strand, coat2d with a binder, and wound onto a drum at approximatcly 10,000 fprn. The coating bonds the filarents a strand, protects them from abrading each othcr, and also serves as a coupling agent to improve the resin-glass bond. For use with epoxy resins, an 801 sizing usually is specified. Requirements for Fglass roving are. contained in M IL-R-60346 under the Typc I classification. Standard continuous roving uses ECG 135 strands - where E designates the glass composition, C indicates continuouis filamenzrts, and G dcsignatcs a filamerit diameter of 0.00037 in. - resulting in 13,500 yd of strand per lb. ECG 67.5 (408 G filaments per strsnd) and ECK 37 (408 K filardunts of 0.00)052 in. 2-18

n

N

243.1.2 S-glass This composition, sometimes called S(994), was developed under Air Forct; contract for its highstrength properties. S-glass is available in the some forms (roving, tape, and prepeg) as is E-glass. The standard roin~g decsig'iation in this case is SCG 150,

niaigta there are 5.000 yd f strand per l

10 0'~vs

5.A

5s

diameter per strand) ravings also arc avuilabic. A roving package is made by winding a number of strands (or ends) under each tension onto a cylinder. The number of ends ranges from 8 to 120. with 60 being the most common quantity. Staniard packages range from 7 to 35 lb nominal wveight. E-glass rovings are available widely, both dry and preimpregnatec! with a variety of resin systems. Prepreg tapes of unidirectional filaments up to 48 in. wide, having a nominal cured thickness or -ither 0.0075 in. or 0.0 10 in., also arc available. These can also be purchased in tuo-ply bidirectionai (0 deg. 90 deg) or three-ply isotropic (-60 deg, 0 deg, +60 deg) forms.

due to the lower specific gravity of S-Slass- The major d0est46krrent to its wid,, use has been its cost. which is about i 5 oimtes that of E-siess. A cuumiiwiciu giadc.

containing most of the S-glass properties at

s.ew, lower cost, has been introduced. Another development is 970 S-glass. which has

20% greater moduius and ultimate strength than SThe chemical compositions of various glass reinforcements are presented in Table 2-9. IITS-901 and are the epoxy-compatible sizings for Sglass, while 470 sizing is used with S-12 rovings. Sglass rciving requircinecits also are contained in M ILR-60346 under the Type IlI classification.

2-4.3.1.3 boron Filaments These products currently arc made by vapor deposition of boron on very fine tungsten wire. Work is under way to develop boron filaments on glas-1 or graphite substrates in order to reduce cost and total density substantially. In order to make handling practicable, the material usually is supplied in tollimated prepreg tapes that are one filament tihick and up to 3 in. vvidc. A Military Specification on boron filamen~t preprcg is MIL-B-83.169. 2-4-3. 4.4 Craphlte A wide v'ailety of filamentary carbon p~roducts is produced by pyrolysis of organic fiber!. These prod4ucts may be divided into two broad categories: low-niodulus and high-modulus materials. Lo%% modulus carbon and graphite are iused freqjuently in

-

AMCP 706-202

NOMINAL COMPOSITIONOYCLIASS

TYPE

SiO., A12 03 MgO0

E-~GLASS

54.3

15.2

4.1

S AND S-2-GLASS

64.3

24.8

10.3

970-S-GLASS

-modulus

*

graphite crystals. Material developed in the U S usecs ra!vyQn fibers and hasq an irre~ular trpopcorn shape) 3~=. cros sctin. Matti ial developed in Engiand i, pyrolyzed from a polyacrylonitrile (PAN) precursor having a circular cross section. In either case, the average filament diameter is 0.0003 in. The Brilish PAN material is made in untwisted tows of 10,J00 filaments, and is available in continuctis lengths. The rayon-derived, high-modulus graphite used in the U S is made in continuous leng~ths from 2-ply yarns having 720 filaments per ply and 1.5 or 4 twists per in., depending upon th~e manufacturcr. The greatest development activity in highpromanc. flie s lfI'-,su~e uponi graphite. Dupimtn the small filaments, it can be formed around radii cts *sinall as 0.05 in., a major advantage over boron fiber. It also is expected that the greatest potential for cost reduction and product improvement lies with graphite. Evidence of both was displayed recently in the comi.:etrcial announcement of a 75 X 1IV psi modulus fiber at $400 per 1b and a 30 X 10, psi fiber at SCO per lb. Laboratory quantities of 100 x 106 psi modulus fiber have betri produced. Because new products and new manufacturers fre(luently enter the field, the data ini Table 2-8 include only those products with which a significant amount of experience exists. G(raphite fiber can be pr,-duced in the same variety of forms as can glass Fiber. Thus, in addition to yarn and to%, fabric, mat, and chapped fiber can be supplied. As with boron, however, the most practicable form for most applications is unidirectional prepreg

z

-

-

CiD

BD 3

7.

8.0

-

-

62.0 19.0 9.69.4 1--

woven forin, which is produced directly from rayon fabric at a fraction of the cost of high-modsulus. graphite. Hlowever, these products. used primarily for -high-tempercr-mure insulation and ablation. have no known applications in helicopter construction. *High-modulus graphite fibers are produc:ed in a three- or Foor-step heating process. During tme final step. graphitization, the fibers are held in tension, thereby imparting a high degree of orilentation to the

-

GeD

tape. A Mlilitarý Specification on high-modulus graphite fiber preprcg is MIL-G-83410. Among the most serious disadvantages of graphite coinposites are poor abrasion and impact resistance. 1 hiss. surface protection frequently is necessary. Allso impeding the exploitation of this material until -cr% recently has been its low interlaminar shear strozngth due to poor resin-fiber bonding. However, surface treatment% have been developed that result in %hear strengths above 10.000 psi.

2-4.3.2

R hiu, While all of the resins discussed in par. 2-4.1 have been used ir filament winding, epoxies are used almost exclusively for aircraft applications at normal operating temperatures. Where nonwovcn, high-performance reinforcement is ustd, the bcst available resin system also should be chosen since the difference in resin cost represents a very small pcrcentage of the total part cost. Phenolic and polyimide resins are used only where very-high-temiperature operation is specified.

24.3.3

z

Manufacturing Processes

Structures of nonwoven reinforced plastics ma) bc formed by filament winding, tape wrapping. automatic tape layup, or hand layup. Filament winding can he performed with glass rovings. gr;aphite yarns. and boron singlk filaments. This process is practicab-le for %cylindersand tanks with high hoop stresses. however, it is limited to hollow structures with convex surfaces. Normally, filament winding is accomplished by rotating the part on its axis as on a lathe. Parts also have been wou-id by revolving the spool of reinforcemeni around the fixed pail. (ieneral~y, prepreg is used, but wet winding also. is practiced. In the latter case, the reinforcement travels through a bath of high-viscosity (at ambient temnperature) resin system that is healed in order to lower the viscosity for efficient wetting of the reinforcement. When the impregnated reinforcement is coolet'

2-19

K-

AMCP 706-j22an

'...

- . .

to ambient temperature, the high viscosity is icattained. Latent curing agents must be used in order to obtain a rcasonablc pot-life for the heated resin sys-

various rcinforccnitnts in such p-roportionsan orienltations as arc required in order so obtain #Imost any intermediatte properties. The possible

tcm. T41K winding is similar to filament winding, cx-

effects of differenit thernini expansion coefficients must be considered, howvver.

inforcement (generally 1/8 in. wide) are wound. A recent advancice in fabrication technologyi unr ically controlled tape-laying machitie capable of applying prcpreg tape (hicated, if desired) at a controlled rate and pressure, and shearing it at the desired ls-ngth and angle. Still another machine applies reinforcement in three dimensions by weaving fibers perpendicular to the normal laminate. This, of' course, greatly increases interlaminar properties, which usually arc limited to the capabilities of the

construction of the spar envelope, skins. trailing edge, etc.. of rotor blades. Design studies have suggestcd the use of boron andhgmous g rpien these same areas, as well as in rotor hubs, swash-plates, drive scissors, transmission housings. drive shafts, airframe stiffeners, and entire fuselage sections. Boron hardware development presently is more advanced than .hat of graphite because the material was introduiced earlier. However, graphite coinposites are expected to be useful iii many of the same

resin,

applications.

Hand layup is still the miost widely used method where winding is not practicable. This process is no different from conventional layup of glass mat and fabric, except that the fibers are nonwoven and oriented, and generally are preionpregnated with resin. Filament-wound parts usually are cured under wrapping tension pressure only. allthough thicy mnay be autoclavedi or vacuum-bagged. Parts that are laid up (rather than wound) may be cured by an appropriate method as described in par. 2-4.1, such as pressure 'bag, vacuum bag, autoclave, or matched die molding.

Considerable design and physical property infortonicnaneinMLH)K?.PrI ndn Re.4

j

r4

2-4.4

HONEYCOMB AND SANDWICH CONSTRUCTION Sandwich constructio~n, ass shown in Fiog. I- is a composite structure comprising a combination of allternating. dissimilar, simple or composite materials, assembled and fixed in relation to each other so as to obtain a specific structural advantage. They are made of three or more laminations of widely dissimilar materials that can be considered homogeneous when bonded together. The layers include the facings, the bonding agent, and the core. The primary functions of'the core arc (1) to separate the outer layers so as to obtain a high bending stiffness, (2) to oupport these outer layers (the facings) in order to prevent elastic instability when they are highly stresised, and (3) to carry shear loads.

24.3.4 Applications Nonwoven, oriented filament composites are in order wherever maximum strength and/or stiffnessto-weight ratios in specific directions are desired, Thyar otuo~! ratc~~ wl-e nroyisrquired. Typical properties of thes components are shown in Table 2-10. 1 is entirely possible to mix the

TABL 2-1K TYPICAL UNIDIRECTIONAL COMPOSITE PFROPERTlE6 IJASFl) ON COMMERCIAL PREPRE~S

CFoLEXURAL

PLY.-c

FIBER IENSIL E 060511 IF1Xl :A[I0ltSSIVI COI4EN. I PLNG-l, MODULE,STRENCOI1, MOVLIU. STRLIIGIH to Ps ki %______ VOL k,si 1 i' V. ps i k,s 10ps O 61j 109 71. 200.0 7,0 99

4

L -GLAS.i

0.0015

S--GLA5S

0.0015

63.5

BORON

0.0052

50.0

220

0.1

186-232 21.6-3M.9

230.0

1.9

120

245.0

20)

443 460

82)0

SAV-I L SPICIF IC S0 1RENGTH. N ooiUS, OCIH IGI in106.q 6.01Z9 Z.54 I0

9.00

0.0122

s

bIn.r

117.020D 16

3.04 2.45-3.05

12o 390-420

HMG.90

0.0m0

51.01

120

25.5

116.0

25.n

00

7.95

0.054F

2.22

47i

t:

THORNEL -50

0.3011

53.0

104

292)

116.0

24.1-,

-

1.40

0.0536

1.93

465-

:,

NORGANITE-I

0.0130

43.0

-',

821

0.0104

-

-

0.0130

52.5~

0.0552

-

-

S2-'

--

MORCANiff-Il

12 -

___-___

163.5

"'FROM3UCOMPANY'S SCOTCHPLY TECH4NICAL0DATA SHEETS FOR EPOXY PREPREG$;I1009-26 2-20

010 SHEAR

1.2.96

RESIN ONGLASS ARDRESIN OORONQ 1AND0 GRAPHITE.

5.

AMP706-d20

-

A

BONDING MATERIAL HONEYCOMB CORE

2A

IA Fiue21

adwc

tutr

Proerl deignd sndwch onsruFigurhas.mandwinch Stue atreat1tiehavr Rf5.

advantages; high strength-to-weight dnd stiffness-toweight iatios are the most predominant. Secondary advantages include fatigue resi ýiancc, impact resistance, and aerodynamic tfacizricy. A comparison of minimu-I-wvight design for

~ IA.

Honeycomb sandwich is the lighitcst possible material that carn be used to achieve an optimum stiffness-to-weight ratio. A com~parison of various materials (Fig. 2-2), based upon an equivalent deflection. suggests a 30% weight advantage when comn2-21

'

AMCP 706-202

3600

~limited,

oneycom b core const i uct ion represents by far the

______________________I _________

MATERAL j HONEYCOMB SAND~IICH 0.058 NESTD 1 EAMS0.05 NESTD " ' EANS0.08 STEEL ANGLES 0.058 MAGNESIUM PLATE 0.058 ALUMIUM PATE0058 STEEL PLATE 0.058 GLASS REINFORCED 0.058 PLASTIC LAIAE

mnost cases, no more than 3/16 in. Among the materials available for sandm~ich zipplication are conventional honeycomb. foamis, and balsa wood. Advantages of balsa wood and f~m are and their use usuaily is due essentially it) a limited physical characteristic requirement rather than to an overall property consideration. Balsa wood is used predominantly in flooring applications, where the need for continuous support is provided by the fibers. Employment of foam cores in a sandwicht construction is.essentially, a cost consideration. Both b -alsa wood and foami may produce adverse effects. and also may limit the environmental capahilities of' the construction.

WEGT1b 7.79 10.6 1.86 25.90 26.00 420 68.60 83.40

Figure 2-2. Weight Comparison of Materials for Equal~~Doinsuch advantage when compared with a flat aluminum plate.Fiberglas Optimum fatigue resistance is a byproduct of sandwich application. The increase in flexural and shear rigidities of the construction, at no increase in mass. provides for an increase in the fundamental modes of ecttotohigher octaves. In addition, the attach-

most efficient utilization of parent material. Conentional honeycomb cores, as illustrated in Fig. 2-4. are cs-witially hexagonal in shape and are manufctured from almost any material that can be made into a foil thickness. Properties of' hun~ycomb cores can he predicted accuratldy, based upon the con. figuration and the parent material properties. The erits of one type over another are related to the properties of the foil material; the relative increase in efficiency is related directly to the increase in the prpryo1 h oe i one~ycombh core material can be made from metals such stainless steel,ihrsn and titnu rfoas aluminum, iega mrgae --

as nylon-phenolic and polyimfide. Other types of core material include those made from Kraft paper and Dupont's Nomex* nylon-fiber-treated materials. core material provides radar transprncadatssadilti.Ithsowiecrc cntnsadalwls agn.Katpprcr cosatanitlwostngt.KftPrroe mtra saalbei ayvreis n sue 200

uAl21

ment of the core at the facing provides visco-elastic

3mi

damping that prevents amplification of resonance. it ca180 is quite possible to design a sandwich structure for an 3 .1 . AT ifintRe life under cycling loads, provided that the 163 (HALAC maximum loading is no more than 35% of the ultiU~ 160\zmate capability of the construction. A comparison of ~ 20 hr AT conventional sandwich structures (Fig. 2-3) in a sonic environment indicates that the sandwich can operate

*

in excess of 500 hr at approximately 160 dB, (decibel) while skin-stiffened structure of the same weight will fail at less than 200 hr under 130 dB. Aerodynamic efficiency of a sandwich structure is a consequence of the continuous, uniform support of the core materials. This characteristic is accepted widely in both aerospace and aircraft applications. Vertical support of the material in the sandwich construction is limited in span to the cell size, which is,in 2-22

-65

,3(*VADTI

T20 -

F I'VD

It -

TP%-5

-

TUCUl -

162 (1

~%%4

____

C) 140.

460 hr AT 167 OB ___

SRCTR SKIN-STFFENED_______R

120 0

100

200 300 400 -500 TIMIE, lir Fiue23CoprtvSncFage Fgr2-.CmatieSncaigeResistance of Conventional and Sandwich Structures *Registered Trademark

..

AMCP 706-202 when cost is u factor and/or thermal conductivity is of concern. Dupont Nornex nylon-fihbertreatcd core .matcri;.l, though recently developed, has thermal resistance and the properties required for aircraft flooring applications, Employment of a honeycomb core material in a construction is an exact technique. Physical characteristics of the construction must be investigated thoroughly, and rclatcd to available core properties, piior to the firming oi the design. In addition to the structural requiremen;s, the environmental operaring conditions must be explored, Common honeycomb types (Fig. 2-4) include the conventional hexagonal shape, a rectangular flexible core, and the reinforced and square cell shapes. The rectangular core is, essentially, an over-expanded hexagonal core. The flexible core is a configuration departure in that it inctL"Jcs a free sine wave that allows the core material to assume compound curvature at no sacrifice in the mechanical propcrties of the

an additional flat sheet in the center uf the hexagonal cell so a%to favor a mechanical advantage in a specitic orthotropic direction. The square ce:ll core is a consequence of manufacturing case, and is employed primarily where resistance wclding techniques. are rcquired in order zo develop the core material. Although a predominant use of honeycomb core material is for constant thicknesses (flat, single and compound curvature applications), it also is used for such components as airfoil sections. The mechanical properties of the core material in a sandwich construction also must be considered. The core, whether isotropic or orthotropic, may be considered as a continuum spacer for the membranes (the facings). Typical properties of balsa wood cores arc presented in Figs. 2-5 and 2-6. Figs. 2-7, 2-8, and 2-9 illustrate typical properties of hexagonal aluminum core material. Several different alloys are presented, Table 2-11 is a presentation of the propcrties of typical rigid foams.

foil material. Flexible core, unlike anticlastic hexagonal core. does exhibit characteristics of a syn-

The term sandwich construction describes the close attachment between face and core material in this

elastic material. Reinforced hmxagonal core employs

type of structure. Should this attachment be weak, or

RECTANGULAR

I

r

HEXAGON

FLEX-CORE

SQUARE

REINFORCED HEXAGON Fi,

2.4. Commoa Honeycomb Configurations

"VP

2-23

AMCP 706-202 absent, the construction is no longer a saldwich. Attachmcnt of the core to the facings is necessary, and must bc of sufficient strength to develop the full mechanical properties of the sandwich construction. For example, if the construction is loaded to its limit, then failure is expected to appear either in the facings,

or in the core, or simultaneously in both. However, it cannot appear in the attachment between core and facings. It is most importdnt for the designer to investigate the properties of the bonding agent so as to assure compliance with these requirements. Adhesives of various types and propelties currently are available to satisfy every sandwich requirement. Table 2-12 contains a partial listing of

•.2500

common adhesives currently in use. Laboratory shear BALSA WOOD

II

-

- 2000

I

t 1-

,presented

, 0 Imaterial

:000 ,,

S500

2

o

4

6

8

bond strengths at room temperature of aluminum-toaluminum bonds with varicus types of adhesives arc in Table 2-13. Useful temperature range and strength properties of structural adhesives after exposure are listed in Table 2-14. process of applying adhesive to facing or core must not be ignored. For the adhesive to be

•

-The

10

efficient, it must be applied to joining surfaces that are free from oxides and contaminants, and its appli-

14 1G

12

DENSITY, lb ft3

cation must take place under controiied en-

Figure 2-5. Properties of Balsa Wood - Compresic

800 8

Strength ,s Density

-

I-

600600 .ALSA

WOOD

400400

-1-

•=o• 00

_ _

0

(,

5056 AND 2024

4

2

6

20

0 .-

•

8 10

12

14

16

52

/200

-"

DENSITY, lb It'

-

0 Figure 2-6. Properties of Balsa Wood -- "1 Strength vs DensitFl

2

4

6

8

HONEYCOMB DENSITY,lb.It 3

Shear

10

Figure 2-8. Typical "L" Shear Strength

=-2000 1IGO

-0

S150o

''-

-

10.,

Co 80

U 500 C

,I.ib52, 40

L

505b. 2024

40

0 I,.-HONEYCOMB

2

4

6

8

10

DENS!IY, 1b It'

Figure 2-7. Typical Stabilized (omprev%.ie Strength 2-24

U

2

4

6

8

HONEYCOMB DENSITY, 1b f3

Figure 2-9. Typical "U,"Shear Modulu%

10

1cM

OFRIGID OAMS TABLE 2-11. PROPERTIES OF RIGID FOAMS'.

DENSITY

DENSITY,

COMPRESSIVE STRENGTH.

SHEAR STRENGTH

Ib/ft3

psi

psi

MIN

MAX

MIN

MAX

-

CO 2 BLOWN URETHANE

1.0-I.1 65.0

FREON BLOWN FR(,, t1.5

M1N

MAX

POLYSTYRENE MOLDED

2,100

0.21

1.00

600

450

152 4.0

1525 15D0

1,500 100

10.20 11.00

200 65

0.11 0.11

0.37 0.16

350 350

250 250

1.3

4.5

10.0

140

15.0

95

024

0.33

115,

0.5

10.0

8.0

200

13.0

90

0.77

175

46. 4510.0

120

0.24

38.0

5

6,000

0.65

15.0

25.0

600

3,000

__ _-___ S

'L

IC

ON

_--__ E

_

_

.. .0.'11 2.0

8.0

13

110

I10

251

LHEAT E ACTIVATED POWDER

12.0

. ROOM EMP LIQUID

3.0

PHENOLIC.

LOW DENSITY MEDIUM DENSITY

.. 1

.

.....

2 .0_

HIGH DENSfTY

"vironmental conditions. The elapsed time betwcen "preparatory cleaning for bonding and the appli-

'"

cation of adhesive must be held to a minimum. Procss control during application and throughout the bonding of the construction is vital for the devciopmcnt o" the spcciificd properiy for the sandwich Adhesive manufacturer recommendations must be adhered to methodically. Design considerations for sandwich structural compon nts are somewhat similar to those for homogencous material. The main difference is the inclusion of the effects of the core material. The basic design concept requires the spacing of strong. thin facings far apart in order to achieve a high stiffnessto-wcight ratio. The lightweight core material having this propekty also will provide the required resistance to shear and the strength to stabilize the facings to their required configuration. Sandwich is analogous to an I-beam, the flanges carry direct compression and tension loads in a similar manner as do the facings of the sandwich, and the web carries the shear loads as does the core material. The departure fiom typical procedures for sandwich structural elcments is the inclusion of effects for shear propertie:ý "-on deflection. buckling, and stress. Because the

S.

-5" 40 o

0

024 j

.80

.60

0..ii 540 .850

35.. ... 360

650

-; , 2

14.0

I)

0.20

250

10

30.0

020

250

1,100

90.0

030

200

.

.

.

..

115

'.v0

.

.

-

500

0.30

325 -

.LUm3 ur.ru U II.,,,j

0 -~o... 360 38.0

I

..

10.15

PRE-FORMED PACK-IN -PLACE

MAX-

18,000

SELF E".PANDED

EPOXY

M

10.20

1.5

EXTRUDED

.

-

hr-ft (Ff/In.), SHORT I FULL MIN MAX TERM TIME

-

"

MAX TEMP,OF THtRMAL CONOUCTIVITY

022

..

I.30 '

-.

facings arc used to carry loads in a sandwich, prevention of local failure under edgewise, direct, or fiatwisc bending loads is as nccessaiy as is prevention of local crippling of stringers in the design of sheetstringer construction Struiurai instabiiitiy of a sandwich construction can manifest itself in a number of different modes. Various possibiliti ,Iustrated in Fig. 2-10. Intcrccllular buckling (face dimpling) is a localized mode of instability that occurs when the facings are very thin and the cell size is relatively large. This effect can cause failure by propagating across adjacent cells, thus inducing face wrinkling. Face wrinkling is a localized mode of instability that exhibits itself in the form of short wave length in the facing, it is not confined to individual cells of Lcllular t)pc cores, and is associated with a transverse straining of the core material. A final failure from wrinkling usually will result either from crushing of thc core, tensile rupture of the core, or tensile rupture of the core-to-facing bond. If proper care is cxcicised in selction of the adhesive %ystem. the tensile bond strength will exceed both tfl tensile and compressive strengths of the core failure. Shear crimping often is referred to as a local mode

2-25

"' '

.

"

TABLE X 12. COMMON ADHESIVES IN CURRENT USE AOIISIVETYPETYPICAL AHSVYL TRADE DESIG144TION Ar 30

NITRILL PHENOLIC

EPOXY PHE NOLtC

ADHSIE TPE AHSVTYE

3MCOMPANY

INARMCO

METLBON0D402

ViNYL PIIENILIC

TABLE 2-14. USEFUL TEMPERIATURE RANGE AND STRENGTH PROPERTIES OF SYSTEMS MANUFACItIRERO___ STRUCTUJRAL ADHESIVES'%

FM47

AMERICAN CYANAMID

At 31 MELOLONO 105

3MCOMPANY NARUCO

AEROBOND 422

ADHCSIVE ENGINEERING

HT 424 I HYLOC 422

IAGE*

VIY

ME MELBONfl I 5471

NARMCO

IAF iG 3M COMPANY FM123 AMERICAN CYANAMID HYSOL HYSO L 9601 PLASTILOCK 717 BF GOODRICH IRELIAL1OND 711&393 1 RELIABLE mrc IHP 103 HEXCLL

MODIFIED EPOXY 250CUR

METLBOND 328 MODI~iEA EPx F 120.ALL 31.0SCA.,.

GOOO TOEXCELLENTr

30-1 7W0 6

00IO

FAIR TO GOME

PHIENOLIC

225

PHENOLIC

400

2004~9W

-67

130-W00

~

0 UNUOOIFIEO

HECL

HYSOL

PLEL STRENGTH

________

AMERICAN CYANAkIID HYSOL

HýYLO3C 901 B-3

TYPICAL VALUES LAP__________

PHENOICRLE

HP36IXE UNMOO'FIED kPOXY

USEFUL TEMP

PHENOLIC

$00

1___0__1____

13-00

POOR TOMEDIUM POOR TOMtDIUW

m0m

_____

PEOI

,-

____

MODIF IEDEPOXY 250

1540-M

20C'E MOIIEOPOYD 3W0 CURE -

InIl 0W30 5020

IEPOXY POLYAMIEJE

IO 306

GOO

________

GOO33 O_________ 20-_____31_

-OIMD

NARMCO

POO

______

12UREPSE

3MCOMPANY

CURL

t;*otfl~:~.-.~.-.

Lu

i

.......

instability for which the buckic wave length is ver short duc to a low transverse shear modulus or the EC?216 3MCOMPANY core.. The phenomenon of shear crimping o1curs EPOXY I'CLYAMIOF. HP 31( HEXCEI. quite suddenly, and usually causes the core to fail in I CO~ 81y5O ~shear. General instabihiy for configurations having POLYAMICEFno34 AMERICAN CYANMtDfnoeninCELxcept at the boundaries ~951 involves overall bending of the composite wall HPI~.1 301P3132AHECE coupled with transverse shear deformations. Whereas MOOFIIUREMAESFM3 AMERICAt; CYANAMID phenomena. general instability is of a widespread COR SPIRNG AT 41320 31.CEOLPANYutrellrbcln n rnln r oai RELIABOND 398-420 RELIABLE MFG 41 04IEXCEL

(

-

CORESPIICIG A 320 *.Ki

~AnHFSIVFS

TABL 90 SHER BND TABEN .18OSHEAR

A

3MCOMANYnature.

REtLARONO 310B

BOD M

RELIABLE MFC

-13. C H OF DH1I3E

NT

FlDEIE

ADHESIVE TYPED STHEART ADHEIVETYPE STRNGTH NITRIE PHEpLIC*

sufficient core zhcar riidity. The basic design p-rinciples of sandwich construe-

ion can be summart'Aed as follows: I. Sandwich facingi s/mal be at least thick enough

towithstand design streins under design loads.

VINYL PIIENOLIC

4200

EPOXY PHENOLIC

34030

will not occur under dsign loads.

UNMOIFID EPXY MODIFIED0 EPOXY-250 CURE

4500

MODIIEDEPOY -30 CRE 300of EPOXY~~~ 50o ____________of

POLYIMIDE

3300

*AVERAGE VALUES AT ROOM TEMPERATURE.

* ITEST

SPECIMENS ALUMINUM TO ALUMINUM, LAP JOINTS.

2-26

mknsofyi-

2. The core shell be thick enough and have sufflcient shear rigidity and strength so that overall sandwich buckling, exo~ive deflecion, and shear failure

______350

*

Premature general buckling normally is ,ucjbanuTiraiuwn

3. The core shal have a high enough modulus of elasticity, and the sandwich great enough flAtwine tensile and compressive strengtths, so that wrinkling either facing will not occur under design loads. 4.LAIE For cellular honeycomb corms where dimpling the facings isnot permissible, the cell size spacing shWl be small enough so that dimpling of either wall irto the core spaces will not occur under design loads.

In~ addition, selection of materials, methods of sandwich assembly, and material property used for

~

I4

AMCP 706-202 design shall bc compatible with the expected environment where the sandwi.h is to be used. For example. facing-to-corc attachment ihall have sufficient flatwisc tensile and shear strength to develop the required sandwich strength in the expected environment. Included as cnvironnlert arc cllct%of temperature, %ater and moisture, corrosive a!-mosphere and fluids, fatilue. creep. and any condition that may affect material properties. Additional characteristics - such as themmal conductivity, dimensional stability, and dcctrical continuity of sandwich material -- should be considered in arriving at an effective design for the intended task.

2-4.5

ARMOR MATF:RIAI•S

I herc arc available a variet) of armor matcrial, and matncrial combinations that can be used for pas,ive protection of helicopter%. Armor types, vsith appropriate Military Specification references and a relative comparison of cost, availability, machinability. weldability. formability, and multiple-hit capabilit). are summarized in Table 2-15. Table 2-15 also lists the areal densities, and provides a comparison of strength. hardness, shock and vibration, and resistance to corrosion for the various typce of armor

"

FACING FACING

HONEYCOMB CORE

CORE

t t tt

ttt (B)SHEAR CRIMPING

(A)GENERAL BUCKLING

4444

444

(E) CORE CRUSHING

(C) FACE DIMPLING i

D (D)

!

-

,--

SEPARATION ,-

ttt t

FROM

CORE

t t

Figu'e 2-10. Modes of Failure of Sandwich Composite Under Edgewie Loads 2-27

na

-

- .

___.

I

X I

II

-

.

..

I2QQ

II~g

~~~~~~~2xI2

.

.. . .. ..-

Q0C

Q,

.

2

2

2

4

~

0

X I

0"

-12

Y

I

00

~

I

IT

2

1

-

'R

___

____m__7m

001'1 0 0

Ii

a

z2 z

42z

I Z- Z

00

0

rI

ý E

z~ Z~

0

0

8..~ _____

.8

z

2

0 8

Iz 02

=1

D

&0 FF4

U

3~~~24 5.IE

F-

C.

~

4. .2~ .~

-w.)

jS-

0D

0,

Sh:

2-28.

NOLV--Jd

SI~

VN

13X

AMCP 706-202 TABLE 2-16. FABRICATION DATA FOR LIGHTWEiGHT ARMOR MATERIALS

0250THICKPLATESIZE

4MAY

335

K-T SILICC.T4 CARBIDE PLAST IC COMPOSITE

PLASTIC COMaPOSIht

PLASTIC CMPUS'TE 1AONOITEIiC TILE 28,,. DIAG5X10l

UONOLITHIC TIEL MONWLTMIC TILE l.T~. 4XWIPANELS EA6,,,.. 7 1AiR..

HAVE BEENMADE

PANEL

I HICK

1

VERYSIVALLRADIUS IN 130THDIREC'IONS INANNEALED CONO

Im. DEEPDISHES HAVLBU~N ExlvPLOSIVELYPOIRMI P

I

NC( NODEXPERIF

O AVAILABLE

-

AI

SHAPES MAY CURVED

SLIGHTIMPROVE-

EON-,

TC~LREENRE

FLATPANELS

TOGI5CEUTIRED

7TRCHCU17TING

PRONELDING, CEDRE REUIED

STAINAUSTENITIC LES'STEEL IMIC OR SU6%ERG(C ARC) OR LOA PERRIIIC HVTTROGýIE

GAS ORHPLATIA ARC_________

FRCTIGTECHNIQUES

ANNEALFO CONDITIONt HARD[EX HEAT TREATABL IELECTRCTE IMC GIVESBALLISTIC JCINTHEAT -TREATED

.- DIAMDO I DOLS~

BE NELDEB M

HHOSCAN'T

COCNEITION -STAINLESS NITH 700EPREHEAT

FJCTRODDES V

WEQIR 1', ECxCARBID10E i9ITHSPECIALCARTIISHTORCHME AREINC TAIED1 OQ IHIAI AID O2HHA FR HEAI TIP.,OPOT LEUSENJ MAY

FDRILLING

.RI L:AGf'OSSILE

USINGMASCINARY Iy 8s yjOT

WITHSPECIAL CERAMICDRILLS

YES

f-

RE'WINAIN

IISERTS THREADED

VARIOUS ýYPES or

FOI 7A~WN 3TRiCTLIAE

, IN FIBERZAO TYPE PER-i~-IPHEKALSUPPORTIINSERTS

BRAKETETRY A4iDTHREAODE IN8tIIINS

ITO

FOR ATACHIENTTHROUHBDLINGFIANGE -

IN BOLT71 COPESIok COWICI.-EEATIGIIIS

-

NOTIMPORTAILT I

-~n

TLS

-

0ELVINS,DP I""HIIIECYANICAL JOINTS

PAWE. JOINWKý NITmOES

-

.ýM

ALEtIT C1ONii

Z

mawlAu~s. Tabk 2-16 su mrzsfbiaindt tht a

be, cosdm

o

ror instalia'iion on heli.

Aotes

2-4.5.1 AvILk14 Meakuil Maw&t that cana N. cunsadereG for Ue dn armor design. aimi t~.cif iropertics. includc 1. Aluminum alloy. ProdlicrA ku splash from bulk.t impacts thar., other armor materials in comnn~ vv, is excx:%*vnaiy effectivc ngiainw~ yawed ard higto-obliquity impwcU. and is nonmagnelic. 2. Titanium. Nonnaagsscic arid ralaistant to sea w'a~er corrost*n. 3. Homogvcou' steel. Rotoid fromn a starl alloy with the toughnes and pucct of elongation naxsswy xo achwor a good rzintence to bodth punc-

Must BE 7F PRAMIC1 RELIEVEDor CON PR[55IQFN5UNDER BOILTý ME EAD WITH THROUGH CRMCI 130.TING

CONTINUOUS Pj ASTIC B'.1KING.PANELS WAY RE APPLIED T'1omo A PFRAMEUSING METALDOUBLER

CAZ WE OfACHIEVEDC Pv RELOIN'.WITH CAh' OE A-HIE. L, B~t EfLOIN,. U141ES3 IAR[Lh EtECTA3CE IN IIRV EtDING UNLESS

I'

PANELS UPTO02.-..

2TGi.n.SMALL OUVATURE XT;AES8T ALED INGON :.C THICKNESS END IN TN ANNE APO CON,I ,ION

IATIKES

EFE ROL OPCLSATLL

2fx6x%,...UPTO 0001I6

9,00"H..

P01,5

EXTrNTO~r.OMuu4 6AUE

TAL6,HAR, NrSS UORTS1CL

fIMAX

1

I1I's01HICK

,!) RDUS Fra'P CAL 30 APIIIIEA1T

1

ETTREA"E, TE

qIGii-I-ARDINIES SSTEEL

ueadsaln.I

ONE

TO BACKRIG MATERIAL SIZEAND Of REOUIRED SHAPE

WIHVARYING DiGREES VKA CE

sdtsignatedmantc

ale" austeilk aftoi am not mwAgnt. If thes steels are coldworked, bowevu, tbcy become mnagfatic. 5. Hard faced steel armor pl,;te. Comnposed of a h'ird surface overlaying a softer IMck~ing matcrial of tougher steel. It is somewhat moire effective. on a weight basis. against solid shot than isface-harwkwA armor plate, and can be fabricated, by spoial tithniques, to a curve. It is magnei. 6. BalI6tic nylon. Providaes czoluen protectiot. from fragnaesols and tum~bled projuctika. Balistic nylon pads or quilts can be owasýJre for replaemont of insulation and soundi-attenuatirg b~les W e~ad/rtaIcn The ballisti level of nylo 2-29

AMCP 706-202 with fastcnc~s and/or attachments should be etablished and/or verified by gunfire tests for each confguraion.at 7. Ceramic. Built up of various materials, each intended to perform a particular furiclion in defeating the projectile. for example, a glass-fiber-rcinforced plastic to absorb the energy of impact, faced with a layer of ceramic tie (aluminum oxiide, A 1,0,; silicon carbide, SiC; boron carbide. B4C. titanium dliboride. TiB., etc.) to shatter the projectile. On a weight basis, some of these composite, compare favorably with standard steel armor plate for stopping solid shot. However. they have poor capabilities for stopping mnultiple hits, and produce many secondary fragments when struck. Ceramic is the bulkiest of the matcrials listed here. and usually is the most expensive. S. Ceramic-faced. The ceramic facing may be applied before or after the armor metal has been shaped or formed. 9. Transparenit. Composed of glass or clear organic polymers. either alone or in combinal ion (M IL6-5485. MIL-A-7168. MIL-A-46108). In general, ceramic armor exhibits the lowest weigpht per unit area for protection against armorpiercin~g ammunition 'cal .30 and .59)). Metallic armior exhibits substantially better multihit capability, although the probability of a small panel of aircraft armoi taking a multiple bullet hit from a high-firingrate Sun is remote. Metaflic armor for aircrew seats may become~ competitive on a weight basis when the arirror is used simultaneously as support or structure. figUred

*

2-4.5.2 Deslga For desgiv strcnith and rigidity requirements, refer to NIIL-A-88&0 and AMCP 706-170.

2-5

ADHESIVES AND SEA LANTS

2-5.1 BONDING AGENTS There are literally hundieds of proprietary adhesivc formulations suitable for various aircraft bonding applicationps. Some of these may be used in bonding a wide variety of materials, while ot~crs are usable only for highly specialized purposes. Adhesives generally are categorized under the two broad classificaions of structural and nonstructural types. 2-5.1.1 Structural A&Wnves Thi, category of materials is used for bonding primary structures that arc subject to iarge loads.'rypical ultimate band shear streirgths are several thousand psi. Structural adhesives usually arc formulatc~d fromt thermosetting resins that, when mixed with a suitable curing agent, react to form an infusible and 2-30

insoluble solid. Depending upon the type of curing agent, the conversion may occur within a few minutes roomi temperature, or. at the othcr extreme, it may reqluire heating up to about 3500 F to effect a cure within a reasonable time. The latter type of material, due to its low reactivity at low temperatures, can be premnixed and stored (often, under refrigeration) as a one-comnponent system until used. Nearly all applications of structural adhesives requirc lixturing in order to hold the components being bonded in contact during cure. this is because at some point during the cure cycle the adhesive goes through ai fluid flow stage. Most structural bonds arc made with tape or film adhesives. These are usually from 0.005 to 0.015 in. thick, and may be unsupported or supported on thin, open-weave fabrics (carricirs) of glass. !,yon. or other fibers. Filmn adhesives have two important advantages: 1. Uniformity. Variations .n thickness and cornposition are minimal. Because both shea, and peel strengths are sensitive to bondline thickness, control of this variable is desirable. Although bondline thick-1, ness olso is affected by curing pressure variations, film adhesives. - particulzrly those having flow restricted by carriers and/or high-melt viscosities can reduce thickness variations appreciably. In addition, film adhesives eliminate the weighing errors and inadequate mixing that are possible with twopart liquid adhesives. Quality control checks can be mnade on each roll of film before production parts are bonded; this is not practicable to perform on each batch of most liquid adhosives due to limited potlives. 2. Ease of assembly. Film adhesives are available in a wide -ange of tacks. varying from dry to very sticky. Complicated parts can be assembled simply by cutting the film to the shape of the desired bondline and laying it on the first suwface. The second surface theni is placed in position, and is herld by the adhesive tack until bonding pressure can be applied. Films that are not tacky at room temperature are tacked readily by momentary contact with a hot iron at sirategic locations. Parts of many layers may be laid up in this manner and bonded at one time. Adhesive waste also is mi~nimnized when films are used because there is no exmrss material left to set up in the mixing container. Most film adhesives require curing temperature. of 250*-3509F. and. thercircr. have long shelf lives. Cold storage usually is advised, however, although some types are stable for many weeks at room tempetature. A few types are available that curm at lower itemperatuics, including room temperature; these must be stored at temperatures well below 0*F.

I~

(

CtMCP 706-202 The other common phy stal form for structural adhesives is the two-part liquid niixturn. These mnatrials consist of two componcats that icact. when mixed, to form a them mosetting solid. Us~ually. tlhey are 10017 nonvolatile. Many cure at roomn ticmpe.'ature in a few hoars or days: other-, requir' hcat to curc Frequently, they atre in the form of high-viscosity pastes containing inert fillers und/o; thixotrofir. agents. In contrast te most film adhesivi.s. however. 'iesc: uncurcd pastes usually becomie fluid whet, heated. Pot-lives, like cure ti.edepend upon the rate of chemical reactivity, which is influenced greatly by temperature. Thus, adhesives that cure rapidly at room temperature may isave only a few minutes of pot-lf.fe while those requiring hightemperature cures have pot-lives varying from hours to months. Less common structural adhesive forms i-iclude onc-component pasle:; and powders, and all icquire elevated-temperature cures. Essentially all structural adhesives of interest for hczlicopir~r applications are based upon either epoxy or phenolic thermosetting resins. Because these rn~utpritik ire bruitt,. ~inhernith thri' iav;i~ly aire-,i~dt with elastortiers or thermoplastic resins in order to improve peel strengths. Polyuretharie adhesives also show promise, as they can be forinulated with both stre-ngth and flexibility. To date, however, they have not been used widely ini structural aircraft applications, Epoxies atz- the most versatile and widely used structural adheaiives. They have excellent adhe.:ion. low creep, low shrinkage during cure, and 100%~ nonvolatility. The liquid or paste typres have either low peel properties or '.empc.-t-urc resistance, and

'~fiexibic

J

arc

gvb

iadi*PAoto.

ritM,ificmuasautnio

(t`01 ainnprovcmvs-z

of these particclar characteriso'Ths) than are the film types. The latter can be modified with tough thermoplastics such as rylo-, or polyvinyl aczetal resins, Primers (low-viscosity solutions of adhesive dissolved in solvents) arc availabki for use in conjunction with modified cpoxy-filrti adhe~ives. Their primary functiorn is to prote.. prepared metal eurface-s rrom contaminatiop and oxidetion since epoxy films have adequate wctti~ig and adhesive characteristics without primers, Phensolic adhesive, 'iscd in the aircraft industry always are modificd &~ an clastainer or another resin Although they can be pruduced in liuid form, they now are used predominantly as films. Vinyl (polyvinyl formal or butyrill) phenolic adbesives wcre thc armt materials usei '9- aircraft metal boiiding. Rubber-phcnol.c evtisAi'ves include those modified With neoprene or nitrile rubber, the latter presently

)

heing the most widely used typt: of ela.ýtomcrphenolic structural adhesive. Lpoxy-phcrnolics arc used primarily because )f their outstanding temperature re!.istance. Being very i gid. they have good shear strength and creep resistance but poor peel and impact properties. 11ccaus,; of the poor %wruingarid flow characterisvcs of the clastomer-phenolic film%, a cojiing of liquid primter on the substrates usually is, advised. A% with all phenolic condensation rcactions, gases arecevolvcd during cure, necessitating relativel) high bonding pressures The actual pressure required to contain these volatiles is a function of thc temperiture: rise rate, 100 psi is a typical recommendation when the bondline is heated rapidly. All of the adhesive types discussed previously may he used for mectal-to-metal bonding. The selection will depend upon the relative importance of such factors as shear strength, peel strength, temperature resistance, chemical resistance, fatigue and creep properties, fabrication method, and cost. Generally, a modified epoxy or niitrile-phenolic film adhesive is chosen for primary structural applicatior.s, while at paste-type epoxy and simple contact tooling may suffice for secondarv structures with 1c, ;-critical requirenient%. The aequirenients fut sevefal cLa.,scs of structural adhesives are covered completely in 'IMM-A132 and M MM-A- 134. Although there is some overlapping. M MM-A- 132 is conceetned mainly with film adhesives while M MM-A- 134 generally has less stringent icquircments which are met by the liquid- and paste-type epoxies. Cured, reirnforced-plastic composites can be bonded to themselves or to metals with the same adhesive's and icchniqucs used foi bonding mnetals. In addition, adhesive prepregs can be used either for an !6

u.*p or ;.- L -odn"*-wer

ii,

.-

,-

--

e*-.-

a conventional r: i rforred-plastic layup and th'e substrate. These imstc~ials consist of a structural grade of reinforcement impregnated at B highi resin content with a resin formuiatior. having good ;dhcsion qualities. Reinforced plastics can be bonded to metals and other substrates by employing betweeti the substrate and layup a layer of conventional film adhesive that is cured simultaneously with the laminate. 1hi-; pirocredure is adv:'ntageous in that it precludes any mismatch of mz-ýtiqsg surfaces, a problem that always exists to somei extent with preformcd parts. While this technique has been found effective with a number of adhiesi',c and laminating resin combinations, such materials must be selected carefully for compatibility with both chtmical reactions and curing temperatures and pressures. Most of the epoxy adhesives also are suitabiz. for bonding facings to honeycomb core in applications 2-311

Ir

S

¶

whern good flow and wetting ability, and low curing pressure ate reluired. Some of the phenolic-bmued adhesives also may be wsed for sandwich construction, although most ar not reownmmnded for

through solvent evaporation rather than by chumicai cure, and therefore do not require temperatur or presure for curing. Because initial taCt often is adequitz to hold in position the parts being bonded,

this purpose due to poor lilleting action and the evo-

even clamping fixtures frequently are unnecemary. On

lution of votatiles during cure. When phenolic adbesivs arn used in sandwich bo-rdini, the core either is perforated or pesu is reased just prior to reaching the fral cure tempemture. MIL-A-25463 coains requirnents for adhlesive for bonding ssndwic•. It defines two casee Cila I for facingto-core bonding only; and Chus 2, for bonding facing to core and iQWerts, edge attachments, etc. Beause mos adhlesive suitable for sandwich comntnu2on also can be used for rmtal-to-metul bondin narly all samdwich adhesives are qualified to both MIL-A25463, Class 2. and to MMM-A-132. The adhesive puepdescribed previously also mnbe used in

the other hand. because these adhesives rmain thermoplastic, they lack the temperature and chemical resistance of tlh thermosetting tactural adhcsive. Where arnewhat stronger or more temperatureand chemical-rea•itmnt bonds are required, semistructural adhesives, such as the two-part epoxia and urethanes, may be uwd with room-temperature curing. Cements based up-au a solution of the polymer being bonded are L -ad frequently for bonding noncrystalline thermoplastics such as acylis, celluloWs, polycarbonates, polystyreves (including ABS),. and vinyls to thonselvea. The dissolved polymer gives

fabricating sandwich panel with reinforced-plaitic

body to the cement, while the solvent softens the ad-

facing. Out disadvantage of this procedure, how.. ever, is thkt a relatively porous laminate is obtaiad due to the lack of labminating pressure between cUl dub* Of SCVUwith Typical bond urergths obtainabie from several of the common types of adhesives ar given in Tab 217.

hecrnds. effcting a weld or bond when the solvent evapo-ates and the plastic rhardens. Transparnt acrylic patics also may be bonded

-5.1.2 Nemssmlra Adlhbes Thxse adhesives are used primarily to bond interior acasaor~s made cf a variety of materials, ineluding plastics, rubbers, metals, and fabrks. Because a joint failure would nor bc cawarorhic in thes ap-

plications, consideration of the highft- possble adhesve strengths is not paramount; and other fooon-, sich as cost and csanvetiemt. cMA be givcn tquaa matention.

"The adhesivte-. prefrred in thew applications

g'lrnelly are

baud "n

solutions or dispersions of

various elassomers and thcnnoplsuics. They set up TYPICAL PWPRtTiES

twnrftmwt

evllhoacnn.

t'nnt

A

clasioae-tesed adhestives are preferred; they offer oot adheson to mmmy matrals ad boetpr I -8

diffcnsat thermal cz iuon cosrAcieps. Elastoeneric

adhmves may be dn.ohlv in a suitab organic solvent or dispersed in water; tackifying reams, aetioxi-

darns, pca

~stiia and naforcia filles am uNWa

TABLE 2-17. o COMMONLY USED STRUCTURAL ADWESI•Y• T I-PEE LP

CHEMICAL TYPE

PIHYSICAL FORM

CURE TEMP."F

MODIFIE EPOXY 14YLOX-EPOXY

FILM FILM

250 350

5100 6100

EPPOXY-FHENOLIC

SUPPORTED FILM

NITRILE-PHENOLIC

SHEAR STRENGTH. psi -67' 75" 1u" 250"

5700 6500

290o 3400

1000 2200

350

3200

3500

3300

2900

NEOPRENE-PHENOLIC

F iL

SUPPORTED FILM

350

410

4200

2400

1800

EPOXY 'GEN PURPOSE)

2-PARr PASTE

15-200

EPOXY(HIGHTEMPMAODi EPOXY'HIGH PEELMOD,

I-PART PASTE 2-PART PASTE

25C 75-200

3500 I900

1500 3000 2000 2000 20D0 2500

1100

800 25W0 3000 400

OALUMINUM ADHERENDS TESTEC PER MUM-A-i3Z AT THE INDICATEDTEMPERATURES, 0

20 60 10

2 2 2

-

30 100

25 90

40

20

3 2 25

ALUUINUM CORE AND VACINGS TESTED PER MIL-A-25463 AT THE INDICATED TEMPERATURES.

S~2-32

SANDWICH

STRENGTH, III in. -67" 175" !80"

PFEL

iui.-Ib "31.WIDTH

-67' 75I

isO,

9, 35

60 170 33

35 45 31

NA

NA

NA

4

NA

NA

NA

2 2

NA NA

NA NA

NA NA

5S,

Il17

.

nOmi'~VI-.

mcthacrjlatc monomer ano acatalyst. whsctu have excellent strength and transparency. MIL-A4576 definft three typls of two-part ncrylic adhesives, type P contains solvent and is covered in MIL-P45425. Types 11 and lI arc withou solvent andmaybe u A for bonding plasics as covered in both MIL-P-5425 and MIL-P-184. For bonding of diimilar maerials, f•lxibk libmU Or fabrics, rubbers, or other such materials,

69mstrag,sjt; Skae~. 4-m W

ftj

36 ]z

(

,

*

" components of the formulation. MMM-A-1617 covers requirements for adhesives based upon natural rubber., neoprene, and nitrile rubber. Adhesives based upon natural or reclaimed rubber are suitable for bonding such items as rubber and fabrics to metals in applications where oil and fuel resistance is not a problem. Neoprene- and nitrile-based adhesives generally have greater peel strengths in the same applications, as well as good resistance to oils and fuels. The neoprene type usually is best for bonding neoprene and most other rubbers and rigid plastics, and has the best heat resistance. Nitrile rubber adhesives are preferred for bonding nitrite rubber, vinyls, and other flexible plastics.

)

Silicone rubbers should be bonded to themselves or to other substrates with silicone adhesives, such as those described in MIL-A-46106 or MIL-A-25457. No heat or pressure is required. Contact adhesives are a special type of elastomerbased adhesive having high immediate strength upon contact of the two coated adherends, but they do not permit any repositioning. They are covered by MMM-A-130. Other speciai-purpose adhesive specifications inude MMM-A-121. MMM-A-122. MMM-A-19, MIL-A-24179, and MIL-A-21366. 2-5.1.3 Pr•'silmg Operalihs Process and inspection requirements for structural adhesive bonding are contained in MIL-A-9067. Factors to be considered include type of surface preparation, control limits and methods of surface treatment. solutions. clean-room layup area requirements, prefitteng of parts, adhesive storage controls. handlin- of cleaned parts. application of primer and adhesive, tooling concepts. temperature and pressure controls, secondary bonding of subassemblies, rework, and destructive and nondestructive verification testing. Equal in importance to the selection of an optimum adhesive system is the selection of the best surface preparation for the adhesives and adhe,.ends being used. Some suggestions are given in MIL-A9067. Other recommended sources are ASTM No. D 2561 for metals. ASTM No. D 2093 for plastic surfaces, and Ref. 6. With some metals, such as aluminum. the surface treatment is practically universal: while with other metals, such as stain3css steel, it is advisable to evaluate different treatments with each combination of alloy, condition, and adhesive. Significant batch-to-batch variations in a given type of alloy may be noted. For most reinforced plastics, a wet-sanding treatment is recommended to obtain a water-break-free surface. When nonstructural ad-

AMCP 706-202

hesives are used in bonding, .i careful solvent wiping and/or sanding treatment wi:! suffice for many materials. 2-5.1.4 Desgn of Beoded Stnrctures Adhesive joints should be designed so that t~ey are stressed in the direction of maximum strength. Thus, the adhesive should be placed in shear while minimizing peel and cleavage stresses. Maximum bond area and uniform thickness should be provided for, and stress concentrations should be avoided where possible. Scarfing and bevelling are two methods that sometimes can be used to reduce the cleavage-stress concentrations at the edges of lap joints. Test methods for sandwich constructions are described in MIL-STD-401. while numerous other test methods for adhesives are contained in FTMS No. 175.

2-5

SEALING COMPOUNDS There is a degree ipoverlapping between sealants and adhesives: most sealants must adhere in order to be effective, while'an adhesive generally seals the joint that it bonds. In addition, many sealants are formulated from the same basic polymers that are used in adhesive compositions. Sealants are related particularly to the elastomeric adhesives, and many of the qualitative comparisons made in the previous paragraph apply to sealants as well as adhesives. In order to form trowelable pastes, sealants are formulated with hirher viscosity and lower tack than are the elastomeric adhesives. Lower-viscosity sealants also are available and are suitable for dipping, brushing, and even spraying. These materials. n•,wever, are classified more properly as coatings. Commercial sealants are manufactured from a variety of polymers, including polysulfide, urethane. silicone, neoprene, acrylic, butyl rubber, chlorosulfonated polyethylene, and polymercaptan. In addition to the base elastomer, a typical sealant formulation may include curing agents, accelerators. plasticizens, antioxidant-, solvent thinners, and inorganic fillers or reinforcing agents. Sealants may be one- or two-ccmponent types. All of the ,tter cure into tough. thermoset elastomers. The one-component sealants are subdivided into three categories: nonhardening putties that remain permanently s.ft: solvent-release types that become s.inihard through evaporation of a volatile ingredient. and types that cure by reaction with atmospheric moisture. Pruperties of the latter, after curing. are similar to those of the cured two-part sealants. (One further form of "sealant" is the cured elastoineric tape or extrusion. Because these nm.st be 2-33

AMIP 706-202

-

held in place mechanicall~y, they moreF properly might he called gaskets.) All effective sealants inust have a high ultimate elongation and a low modulus in order to acexpansion and contraction of the joint rcommodate being scaled. Most commercial sealants have these qualities. They vary widely, however, in their degree *of recovery, ranging from near 0% recovery (or 100% plastic flow) for a permanenitly soft putty to nearly 100% recovery for a cross-linked (cured) elartoamer. Thisproprtyis important because a lowy-recovery sealant. once compressed, must accoRmmodate subsequeni joint expansion entirely by its elongation, or it will fail. A compressed, high-recovery sealant will return, as the joint expands. to its original dimension before it begins to elongate in tension. Of the various chemical types of sealants. only polysullides. urethancs, and silicones are currently of I inmportaince in the aircrikft industry. Thcsc arc all high-recovery elastorners when cured. Polysulidecs are most commonly used in helicopters. where they act as both scalants and aerodivniasnc fairnng comvoundi. They have excellent adcharacteristics and resisianic to suicnis anad fucls, weathering and Aging, and icmrpcraiures up to 250*3. MIL-S-7124 and MIL-S-8802 describe the two-part elastomeric scaling compounds with increasiregly severe requirements for adhesion and resistance to temperature and fuels. MIL-S-87t4 compounds are formulated purposely with very low adhesion for such nses as fuel tanks access doors. A *grade for scaling electrical components is desciibed in NIlL-S85 16. Osme-part. noncuring. polysullfide puttics also are available. A mater-i.dl of this type is de* fiined by M!-Q-1!3 UR .1W ,it isitntfisr w ling ofitcal instrumenits, but is aseful for various purposes. Silicone sealants have autstanding environmental * resistance because they are unaffected, relatively, by * temmeatures ranging fromn cryogenic. to more than SW5*)F.~ and by moisture, ozone. and -iltraviolet radtat ion. However, because they arc the most expensive %ealant%. they are used4 only where these ex-cellent propcrties are requirod. Some: types also have very -htsion

*

gotod electrical characteristics, andi arm used to seal electrical systerfs. MIL-S-23586 covers siliconc scalants for electrical applications. and MIL-A-46106

describes a general-purpose, room-temperatugecuring adhesive-sealanit for both mechanical and electrical requirements. As ordinary silicones have relalively poor fuel and oil resistance fluora-silicone sealants should be ausd where these properties arc required. Both one- and two-component maierials mein common use. The former cure by absorption of atniosphem mc humidity, and. therefore, cure very slowly 2-34

in confined aircas or in thick sections. Primers usuaillv are recommennded to permit maximum adhesion to) m1cwKis Although of totally different chemical compost-Llion, polyurcthane sealants have many similtrities to the silicones. Both two-component and onecomponent moisture-curing types are common Primers (often silicone- based) are recommended, but. in this case, primarily fur retention of adhesion in humid or water-immersion situations. Thes escahnt exhibit complete recovery after extended outdoor exposure. They also are useful in cryogenic applicationr, where they are surpassed only by the silicones, and in electrical appl"cations. Polyurethancs also have excellent oil resistAnce, and grcater abrasion resistance than any other sealants- Osie problem is loss of adhesion upon exposure to ultraviolet light. An area where scaling compounds frequently are used is in edge-*ealing of honeycomb sandwich pancis. because joint expansion and contraction are not major considerations in this instance, relatively rigid sealants usually are employed. These are essentially thc same materials as the epoxy (and occasionally urcthane) pasic avhiicivci discuassed prev~iously-. cxcept that microballoon (hollow microsphcres of glass ( or plastic) fillers frequently are used to produce a 4 lightweight. closed-cell structure. Sandwich panels * also may be sealed with an edge w~apping of Fibergias prCsprCg. Viscous sealants mxy be applied with a variety of equipment, ranging from a putty knife to a cornpletely automatic mixer-dispenser system. Fluid sealV ants (coatings) may be brushed or sprayed. One.compawenst sealants suapplied in czartidges can bc apniod from miolan

ornl Ofir-oterated

*ums.. or these

seslants can be diwspesd directy from pails or drumis by air-powered Or hydraulic pumping equiptrient. Twvo-psar wsearlas can be. weighed and mixed by hand or by mcteing-mixing equipmient that dis.-el penses the compongnts according to prelet ratios. Frozen cartridges of pranixed sealant akso arc available4 commercially; thewe must bc stored at -40*F until just prior W~ use.

24 PAINTS AND FINISHES "IP

NTAD O IW RG 10 AITAN CO ItSOG II MIL-F-7179 prescribes in cletail the manner in which the external and internal surfaces of at khecopter arc to be finished. Uther helpful documcni~s arc TB 746-931-2. MIL-STD-171 (MCR), and AMCP 706-100. Helicopters require a Type I protection. ime., protection Against severe deteriorative conditions, F-or

241

)

most .;.jrfaces. this involve.; one coal of wash primer (MIL-C-85l4). one coat of primer (MIL-P-23377). and tv 3 top costs of attopcoat rot examiple, TT-E516 or M IL-C-8 1773. Preparation of the surface for painting will differ with the type of m, tal and with the surface (external or internal). For this handbook, exterior surface arc defined as all visible surfaces of an end-item that is housed within the helicopter and all visible surfaces of the helicopter. including all portions of the system that arc exposed to thc airstream. Interior surfaces are the nonvisible surfaces of an end-item that is housed within the fuselage of the craft. Prior to painting, aluminum surfaces usually are finished with Anodize MIL-A-8625 or Alodinc 1200 (N4IL-C-5541). and malnesium with Dow 17 or HAE (MIlL-M-45202). Nonstainless steels are phosphate-treated (MIL-P16232), stainless steels arc passivated (QJQ-P-35), and Fibe.'glas surfaces are sanded and cleaned with naphtha (TT-N-95). The first coat of paint applied is the wash pri mer. The term designates a specific mate-ial the( comd-

~hinr- the

-

nmrmrwflc nr shn inhihitiwi twacreth

.-sr.

craft arc the nitrocellulosec and acryiL-nitroluvllulosc lacqucrs. which contain a wide range of pigmcntatior. They art preferred becamusc of the case in iemoving thew. with solvents when it is necessary to change camouflage or color schemes or when repainting is required. They also are applied with a spray in volatile solvents (MIL-L-19537). TT-L-S 16 desceibes another suitable top coat. and one that as meets air-pollution regulations. This coating is a styrenated phthalic alkvd resin combined withi the necessary amounts of driers and volatile solvents. Thc mixture contad'ns 50* resin solids. inc~luding small peircenitagom of antioxidants, wetting agents. and stabilizers. A wide range of coloring pigments isV available, and these arc present in amounts of 24-45%. of the total solid conicrnt. There arc special paint formulations for camouflage. battery compartments, hilgh-temperature areas, walkways, and antiglare applications. Rain-crrosionresistant coatings (MIL-C.7439) are used on the Icading edges of the rotor and on radomes. There arc sixciail formulations for high visibility, and paints forC.f lettering and marking. Rubber, both natural and syn*hei,

...,,

surfwms suh as. ;!L-.

n

metal conditioaer. w~th those of the convcntio;:J-anticorrosive primer. The essential coDmponent)s D: wash primrrs are phosphoric acid. chromate merit, and polyvinyl butyral resin. Wash primers can be formulated that are effective equally over iron. steel, aluminum. treated! magnesium, copper, zinc, and a wide variety of other metals. The advantages of wash primers include ease of application and rapid drying, useful iange of tzmperature and humidity, application to a variety of metal, effectiveness in prc-

plastic windows, arc not painted. Particular attention must be directed to assemblies in which dissimilar metals arc joined. It generally is required that each of the mating surfaces shall be finished with the minimum number of coats required for interior surfaces. Where magnesium is one of the metals to be jcined to a dissimilar metal, tei metals shali be separated by MIL-T-23 142 tape or MIL-S8802 sealant. The tape shall extend not less than 0.25 in. beyond the joint edge in order to prevent mois-

venting underfilm corrosion, and good adhesion as a %peasmc suuuu SU"rI W&kla. 11 11U1 1eU4iUGLY

tur-s from bridging between the dissimilar metals. All naotes and couniersinas that attachin; pans pass through should be prinied, and all joining bolts, screw%, and inserts should be wet-primed when inserted. Preftrably. all stecl nuts, bolts, screws, washers, and pins should be cadmium-plaited. 46 PCA IW E

used wnsh primer is that defaned in NIIL-C-8514. a

smooth-finish, spray-type, pretreatment coating furr'ished in two part,~: resin component and acid comnP;ýwint. The materials must be mixcaJ prior to use. The piameir, which must cumform to MIL.P-23377. is used oam the wash pritm. It iscompatible with t usual =cYlic-aitrocelaloW Laqutu top camt. as well ass with the alkyd top coats (TT-E-516) sand urethane (MIL-C-81'.73). The two-cnsnpoximt, epoxy-polyaamide systern has high dhanaicu and solvent resintanc OWd Unusal wmhrbi.It is Wsjy-ajWIpled This specri~atic. also provides for an addkiicua dlams or materials wshitalc for usew~and ir-pollartion regulatiom. The aivailabiliy of daumas or costauiua Ai wpo~itgkg Smir rqpktimup is beomia# increasimigly iunponsat.. a=d this flactor shoul be kept in mind by the dein*=. Tb, top coatsw mofst osm pscifind for Azmy air-

)

42SPCA

3

FNSE

In addition to the finishing of surfaces with organic coatings as described in the previous paragraph, Ithere are a number of special finishes for metal which serve to provide the desired protection without further applicaiion of organic coastiags, or that arc used to provide a suitable base for the application of organic finishes, Many of these processes involve the devefopment of a durable crrosion-resistant. oxide layer on the surface of the metal. Although the development of this surface o.%de film may or may not involve the use of an clecuical curient. the chemical effect is similar and the proess is called anodizngs. 2-1)

-.

'

4 p

r

~ *'.'l .

I

there art many different finishes. For alan some or whic ame used to provide a bws for paint amd some of which provide protiection withont furthe painting. The processes all involve chromates noan eaidiAia iesgedimat, and rnll have proprietary compositioes. The performamot of these methods of tutanamn is governed by MIL..C-5541. The reagents say ba applie by sprying. dipping. or swabbing. generally. the metal is dippe in a sequence of baths ad inuus that cwrme dean, rine-rained oxide uniformy ditrbiated ovea the surface - with no coarse grains and so untrwated areas. Class I A treat. me"muim wihsawd exposure to sult spray for 166 hr and ane unpainte or follwad by wash prime and pimin tmtmeats. Class 3 coatings are simiLar to Chmu IA coatings, except that the deectrica reseisumie is low, Formagesim. hemarctwo primary anodizin~g manesium there ar te Do7tramn Foreamns (MIL.M-4S202) and the ote meho is the HAE treatment (MIL-lW43202). both of which involve elecroltican iziag f te mtalsuracein rde to buildolyui an fairlytikae of ahcmptlesrs=i ourder to a~

buid

~d

~

hic u

a lyerof aily

a..

''~

coplx

to provide hydro3gcn embnittlcment relief. Still another class of finishes frequently employed is the flame-sp.rayed type. In this technique, metals. silicon dioxide, titanium dioxide, alumina, or other ,m-A=rand -mnnr

aumium

is

'~

Ferrous metals that arc to be painted aregiea phosphate coating in accordance with MIL-P-16232. The~se coatings are of two types: Type M. which has a phos.phate base. and Type Z. which has a zinc phosphatc base. Type M coatings arc more resistant to alkalinc environments than arc Type Z coatings. When they arc applied properly, the reaction form?, a mixed-metal phosphate coating on the surface of the firrous metal that is dose-grained, Fine, and Wre of powder and course grains and that Ws¶the surface well. Thc treated surface is more res"5stat to corr-')3ion and provides a firm basc upon which to apply wash prime and prime coaings. In all of the foregoing treatments, the metals employed, especially the ferrous mectals, aro subject to the absorption of hydrogen from the solutions, and the hydrogen serves to embrittle: the metal. It is desirable to promote the diffusion of hydrogen from the metal by heating tt 210a_225*IF for 8 hr in order

fa,

-.. -..-.

n

ý... -

oxyacetylenec flame

~~ria or .Ai imln a - - --- -,;o.

nittnam nr -*. r

either as strands or powder

-

outdoor exposure without further coutin~s. The prIewhere they are vaporized and deposited on any subfetred method involves pretreating. primting, and art strate that wili condense and hold them. By this epoxy-polyamide finish. means, similar or dissimilar metals cam be applicd to metallic or nonmetallic surfaces. Ceramic materials Part made of co. rosion-raistrat sfteel arc passivated in oider to deveiop their corrosion-resistant can be applied in order to provide abrasive surfa=es qualities. This preceess serves to remove the "activewear-resistant surfaces, or flamec-resistant coatings. centers on the surface and to leave a thin, durable, M11L *674 covers the flame-spraying orf metals. transparent layer of oxide that prevents further corWo., a metal surfaces can be built up and subseauc oroxipaivhebyac ioesng the mea.parssinatn aqeos puii~s mtc.hOned uslapicain hade enai be thefam-stns rcomlshved or oimmierstaking the mertal ainatin iqeosquetc.:l mahned usflapicain orde toeepai thftpisons ?$AUSUu

t

nW

~

urnu

01Aumn

W

M.#inous20MUM

charomate. The temperaure of immersion varies from. 70' to I 55F, depeniding upon the alloy involved and the intended operatiag tempeature. Prior to treat ment, it isessential to wash parts carefully inan alkaline solution in order to remove all of the particles Owr iron that may have acewiwlataid ea the surface, as these woulod develop rua staie. durting the tieatnh.&it. The paswvation orocus maistatied in QQ-F.35. Ferrous surface that ame not to be painted usually are treated w~ith black oxide. The insulting depos*it is a hard. durable. oidin surfac. dma isattracive and is so. cwhat res"ISta to coroiomn aOW to wear. The process is applicable to both womstaolese ami simailes steels, it invoives imnaersiM the previously cleaned part in an alkaline. or alkafine-chromtate, oxidazing solution, followed by warmi and then cold rinses. a fInal chromaic aciod dip, and drying in

spraying oi toaimur

Iwater

worm air. The proý is defined in MIL-C-63924. 2-36

cauxvic contingi upunthe1 mflt5i

exhaust skirts or*je enginies for oxidation pl-3tection. 2-403 PLATINGi Another method of applyi-ig attractive, durable and abrasion- and corrosion-resistant coatings tiG metals and plasticsias meta plating. The platings of most interest in lieheopter dwmii are copper, nickel, chromium, and cadmium. Extupt for electrical cornpoms whecre electrical conductivity is impcrtant. copper plating is used only to provide a bae for thc wear-resistant nickel and chramijim plating.. Chromium and itickel plating. am used to provide hard and wear-resistani suirfaws for suiad object as SWa fasteners, strap holders, handles, knobs. seat arms, instrument parts. and other itemi where painting would no be satisfactory or economical. Cadmaium and zinc plating are employed almost exclusively to

provide galvanic protection Mgainse cxonvon. Cadmiuna is fthpreferred ecoming for ferous meal iAem

-

"

k2

such as nuts, bolt, screws, inserts, Lod pins used in assembly. particularly where dissimilar metals arc

izing plastics, and. becausc the danger of hydrogen

employed. A treatment of m( tal plating can be found in Rcf.

embrittlcmcnt is nqI1gi.', it also i- used for the plating of high-strength stmi: parts for high-stress ap-

7. In this pr

s. & ur methods used extensively: chemical reduction or elt-ctroless, vacuum vapor deposiluin. and molten metal dip. In electrolytic platin&, the item to be plated is cleaned so as to provide an oil- and dirt-free surfic, and then is connected as the cathode hn an ecoctrolytic coll. The anode is mtde of the plating material, and when an electrical current is passed through the

plications. It is the preferred method of cadmiumplating high-strcngth bolts and nuts and other fasteners, Fnd MIL-C-8837 detail. the requirements for this application. Galvanized steel products arc made by dipping the cleaned, preheated steel in molten zinc. Cadmiumplated parts also are made in this manner, and earlier tin coatings were applied by the dip process. How-

electrolyte the netal isdeposited upon the surface of

ever, the dip coating of steel with cadmium and tin

the itn being plated. In somi c4ses, the item first is coated with a thin layer of coppep, which adhers

has been superseded by the more economical and more prccisely controlled electrolytic procesm.

Selectrolytic.

readily to the base meta and forms a finn surace to which the platirg metal (either nickel or chromium)

U

•-

and tenacious coating. It is used frequently for metal-

MIL.T-l0727 covens elecroplatingt and hot dipping of tin.

can attach firmly. Ther are a great many proprieIn both the electrolytic and electroles processe, gtuy Oe•aolyuc sciutior, formulations and processes. hydrogen embrittlement is a ,agnificant hazard. DifConsiderabl, skill is required to ob.in a uniform, fusion of the hyIdrogen into the metI under the elecfine grain and brih coat, and much care must be trolytic forces is gSr-'ter than in the case of elctaoawcied to asure cleanlisms avoidac of poisom less deposition. The danger increases with the pn enUY of 41rms. biisenug ad cracking. an strengah and moduius of0t1 iatd materiai. 1-01u it avoidanc of ccarWraieed platings. Appliable is necessary to program a hjdrogen embrittle•mt t'CFedera Specificatis am (JQ-N-2l%, -Q.C32O, lief heating cycle in order to piomote the difftinn of Ja QQ.-P-416. hydroge: from tie basic metal. The optimum t,:mc LkElctr s or chemical-redutin plating depends and tempcrature will depend upon the nature of the upon the gena'mtioi of activated atomws of th1; metal coating material ant of the base metal, as well as to be deposited aiacemt to the wafAvaa•a eal kurupon the scheduling requirememts for the part.

fSX upo whicbh tht olatiMg is V.be m4xiAwJ. Thre are proprietary disa*

mCpo74a,• lsie5 u

lion presses for maya sia~n

rodtc-

used in pktling, and

prqimated or nonimpimqsate, and made from mnasy

the nicel re.1coig •se•v to make the surface more wear-resiwua mand kwa impact-misitive. MIL-C-

of the advanced plastic materials. The tapes asy or maly not have adsive on out or both side

2674 covem the p4Asat of eleciroless nickel.

(peswAre-sentsitive tapes). One application fo tapes is in the maiking of heli-

cop(t.

at so"

M IL-P-36477 may be usd isinlis of point for al rea.

Deca. confonming to the mquirments of

Plun4 wa*l, and are related st as to o,4tni a uii. form cmiosa- Tim son of the plating "4 may be

teusbal and internal mnArkings within the sim limaits spcimd. They may be prasaare4-asve. adhsive-

a hot war or a mohtea pool of t1w mN.a. Th metal is healed ekricaly W a taeperamtun a& which it veporimir, foui thLblsrfac. U ,tr dzh± & ,,t of an eketrosatak held aqp"ied betweenw iwtl swurce ard

backed, and scord and are aplied over pre•,AMly finished surfac. Antislip tale a&re umid on walknays, seps, mmd similar artea wULhiglrvn L" ate m• rapioyed iiiWn)C waica Wnia. Cautio. siwWbcid hausd

the iftsu biti plaed, the atoms. of the m"ga beCots se~amedsad sawatracted! to thme sekrfta. T'his

Pro-m jwodsii

wi ewapuicety bright, coheret.

,

I ..

K

., 244 TArt-I Tapes of varying ompoeitL'on. texture. thwknkos and width arc used in a varidey of ways in helicoptr. design. Fairic tapes may he woven or mmnwoven. ia-

vacuum dckibmr. 'hM i•,m to be plae arc racked of the

L

vamuum plating process and from the molten ametl

Such as nsickel-pltns of magnesium sdaafams whene

from th sorm

1)

drogen embrittlemnt eli•f. Embrittleasnct from the dip process is minimal.

cput•am dAat

14-V

Generally. the platinqpecificenion wil require hy-

also for Sot metl *&s plastc busc uateral,. Them• is less dwaola of hy•#d-,o evibnittlicnent with this prOOM. The' etolmas m9thod is unitablC for healird y;na or for us ift the field or shop wher Lhratoay or proce line facilities arm twn abailable. ft aho is ef~ctive for ==ne of the more diffcult jobs,

Va-uum de*,iton plating is conducsed in a

.

to ilatuiv tisef antisip tNpe edga arn rnot epusmii to ai"ow thatcant eta" the a to gaol. Thm - of 3-37

d FOD, eoý. They point up iVhe importance of careful delineation, in the quality assuranc provai~is1 of the A PU model 4ppv-ification, oý d.-Aign and tesi ronclitions that simulate the helicopier eflvironmiria. The vibration environment is primiary. The profile

of amplitude and froquencies at the APU installation must be defined. The amnplitude and ficquencies of the AlFU-gencratcd vibraiion also should be spc~ifled to anticipate airframec structural probleinis, APU air inlet iemptrature limits arc specified, and the APU installatioan ust be desigtned to prevent recirculttion of main tngiiac or A?t3 *txhau-ýt gas. Carteful aocntion must be 7.iven to thi.t prokilena because ai is impractical ii' attemrpt to limit heiicop ter operfatiori in undesirable wind directions. Ir.1ut duct locatioav. either must be rtmote fianm exhaust outlets, or sdfety shutdowii sen~oms must be pro-

vided. Similar arguments apply co~cerning FOD. stird, or dust inglestioin. Genetally, higi inlet doct locat iors are pi cferable bec-ause concen!vtion i. dust in huvem is straxified vertically. IVany APU strtic-. problems ca!. be traced to fuel aysteen components. St.mc of these re- 61t ffrom helicopter fied system contamination. Addit~onal trrphasis should be giv.-n to filir~tion, and to Pk'C. oenur of main ýRnk contamination by proper fuelhandlicip methods. A major sou ce :)f reliability problems arise& fromi

-'

706-202 maintenan;ce reqluireenirts cith.cr scheduled or unschcduled. For example, more contamination is in. troduccd into oil systcms (causing exccisive wcar and early bearing failure) through frequent oil level checks, oil addilions, and oil changes than through seals arid vents during normal running. APU design should stress minimum scheduled riainitcinacie, throwaway filters and components, scaled systems. and automatic controls requiring no adjustment. This approach not only will increase reliability, but also will decrease not life-cycle AF~U costs. M,5 SAFIETY PROVISIONS Good APU safety design must inciude provisions to prevent a failure frorn causing helicopter diemage, and. if possible, to permit mission completion in event of a failure. Thus. the APU installiption shall be designed so that fire, APU rotor failure, and crash damage arc contatined within the APU compartment. APUJ-rotor constainment is an important safely cnieain n a ehnldi eea as Strulturc can be designed to withstand and hole a tni-hub burst at overspeed trip condlition4 (a.ssurniing a fuel control failure). but this causes an une~e'irablc weight penalty. Alternatively, rotor integrity car, bc. d-mcnstrated by in.....J.p.c-.-.-lizc tests to micarure stress !evels under operating coitditions. Furtkarr. systems cani be arranged to gv'Arantee that blade failures oct.-ir first (e.g., stress groovc.9, but that smaller mass bladr fisiltirc cart be contained within the casing structure.

Fuel and ignition sour-ces s1haý' bc: separated by m~eans of the compartment deiign. On'ý philosophy is to put the entire A PU into a fireproof ci~nip-irtmcnl. A seooaid philosophy sodi-s to prevent f1ie by confiniing 1w.) sourucs and by s.-gii.gating the hot scction %v'itha bulkhead. The APU coritrv~As arnd oil sumn should b,: housvo in fireproof contair'crs. Electrical itibsy-terr ignition sourcce., should be roijted or housed away hornm bel lines. APU inlet air sh~ovld be dlucied from outs'd. thc helicopter to prevetit rocirctrlation in case ol cornuartinent fire. [ire detectors and fire wxinguishing cquipwmemt shall be. u~vd to pr'olcc: against firc within the compartioien (see par. 32.4ý , ll APU fuel system romponents shall te crashworthy. Fuel line-s shall be made o: flexible hose with steel-'hraited outer fheath, with the mtn.'rum ni-niber of .ouplinps. At bulkheads, the hosce should bc run through unc'it, using fr*mngibic hose stabilizer fittings. Whcn lines So through w firewall, selfscaling, breakaway couplings shall be used. All line supports should be frangible. Lines shalt' be 2CLI30% lo~igce. than necessary to aicomroadate structural dispiacements.. Routing shall be aong the heavy basic 3-21

AMCP 706-202

-tPreferred

*

structure. but away front electrical comp,.mnents (unless electrical systems art shrouded). D~rain 'incs for -.onbustor. fuel pump. gearbox, vents. etc.. shall be connected with frangible fasicrners. and made of low-strength matcrials. design calls for enigine-mounted fuel bogst pumps with suction fuel supply. In the event that tank-mounted boost pumps are required duc to fuel subsystem configuration, they are acceptable if meunted with frangible attachments. Electrical lead wires must be 20-30% longer than necessary. and shrouded to mirimize crash danmage. Self-scalintE, bicakaway couplings shall be used at all c )nnections. Filters and valves shall withstand 30g aoads applied iii any direction. Electrically actuated valves can-hec bulkhead-mounted, with wiring on one side and the valve and fuel lines on the opposite side. APU oil tanks and coolers also shall withstand 30g loads. and 'must be mounted away fte-icn impact areas. They shall be located within ttue compartment. but away from hot sections and inlet air ducting, to prevent ingestion of spilled oil. Oil filters shall be integial with the APU. Ratteries and clecoical accessories shall be located high enough in the fusciage to remcve thanm frt'ii possib:, fluid spill~gc areaý. They vA~yll ttt comnrtrt-

3-22

mntcnaliscd with lnexible fire resistarnt pancls. Extra %."rc length is needed and s~hall be supported; wilh frangible connections. The basic structure .thall %ith. stand 30)-g loads applied in isny direction. REFERENCES I . C. R. Bryan and F. F. Fleming. Some InternalFlow Characteristics ci Several Axisvnimetricol NA CA I-Series Nose Air buttis at Zero Flight Speed. NACA RML54E 19A. July 1954. 2. K~uchemtaiin and Weber. Aer-odynamnics of Propulsiion. McGraw-Hill Book Co.. NY. 1960. 3. J. Seddon,. Air Intakes Ja~r Aircraft Gas Turbines. Journal Report. Royal Acronaoticall Society. October 1952. 4. SA E Aerospace Appnlied Thzrrnodynamics Afan*a'l. Society of Automnotiv. Enginccrs. NY. October 1969. 5. T. Himka and R. D. Sekmple. Enginc/Tranx.mission/.4irframne Advanced Jn~rg ration Technique~s. TR 75-16, USAANIRDL, Fort Eustis. VA. May 1975. 6. N. 0. Johnson. Crash worthy Fuel System Design Cr-iteria and Analyvsis, TR 71-8. USAA I'Al DL. Fort Eustis, VA, March 1971.

AMCPX70202

CHAPTER 4

TRANSMISSION AND DIIVE SUBSYSTEM LI4T OF SYMBOLS A

C -capacity of bearing far lire of ii! cycles with 90% probability of survival..' - case convection cooling cotfricient, AP0 - load rinclnation factor (helical get), dimensionlest

- liner (steel)O01, in. speCuic reat of oil, iiitul ln-t F1 b geSar diameter, in. d D actor - macrial buring), dimnsionlss D~c- bolt circle diameter, in.Dp)- stud pitch dianieiter, in. 0,, - anvolute base circle diameter, in. DI - inside diameter, in. D~in - spline minor diameter, in. D, - ouarsroo diameter, in. DP - oitsch diamecter. in. D, - pitchoodiameter, in. D -=Major diametfer of spline, in. D2 - outside diameter of spline tooth mnember, in. d - pinion pitch diametcr, in. d - track of breked wheels, ft d -' light alloy &Wcion OD, in. E - modulus of elasticity (Young's modilus), c

e C_

£

)

- frequency of backward tiavelling wave, Hz race curvature, %of ball diameterr

baxing fo ccolu-iner

B1)- life at which 10% of a bearing population fail, cy,.Ies or hr b Hertzian contact band semiwidth, in. b -bea~ing 013, in. Cc. CA

fa

psi

processing factor (bearings), dimensionless ED energy dissipation rate, Btu/in.-min El combined modulus of elasticity, psi r -pitch plane misalignment, in./in. F - flow. rate, gpm F - face width of gear tooth. in. F - lubricztion factor (hearing%). dimensionless FA, -bicakaway blip for"e. lb F, = effective face width, in. F., - average effectivC face width, in. f - coefficient of friction, dimcniionicss -

DESIGN

G

EHD matcrial paramecter, dimensionless - speed effects factor (bearings), dimensionIns - lengthwist tooth stiffness constant, pjs

h hE

misalignment, factor (bearings), dimensionloss ftoil film thickness, psin. - EHD oil film thickness. #sin.

G G (is H Ii

-

1

specific weight of oil, lb/gal

7 ý

of inertia, slug-ft' Imoment i geometric shamc factor, dimensioniexsst-l

K K K, Km, K, K,

K KI

K, k k k L L

-Hertz stress index, ftpsiniols -strcss concentrationfatrdiesols nrifatdmesols - life factor, dimicnsionl~ss - misalignmricrt factor, dimensionless - overload factor, dimensionless - reliability factor, dimensionless - dysicze factor, a dimensionless Mtemertue factor, dimensionless -tenmperlatur factor, dinensionless - conversion constant - con~tact line in~lination factor, dimnensionless - geom~etry factor, dimensionless - gear face width, in. - design life or scheduled removal time (TBO), hr

LA

=

Lrj L2

-

L10

-

M Mf Inl Mir,

on, Opt"

i

-,-

Adjusted life, hr

gear center distance, in.

- life for 2% failure of a bearing popuiatlon.

hr life for 10% failure of a bcaring population, hr -mechanical advantage, dimensionless - moment, in.-b - prortle contact ratio, dimensonless

.

V Y

= gear rutio. dimensionless

=Con~itc! ratio factor. dimensionless modified contact ratio (spir~l beve; g5or), dimensionless

-

/*

4-1

number of teeth (gear or spline) or bolts or sttds

N

TA

- ambient air temperature (cold condition)

-t number orteeth on gear

T,

- critical temperaturm, F

- number or teeth on piniot

7".

-

T s

- circular tooth thickness, in. - initial temperature, °F

N

- number

NI NP

n n RN nap

PL PLC Pc P ?d P4

number of discrete values number or radial nodes - rotational speed, rpm - critical speed, rpm normal operating rotational speed, rpm - pinion rotational speed, rpm - load, lb power loss, hp - power lors to oil cooler, hp -power loss to oil cooler (cold conditiwi), hp

Pf P, P, P,

base pitch, in. - diamet'al pitch, ir.' - transverse diametral pitch (measured at large end of bevel gear), in.-' - friction power loss, hp - pt ver inuut to transmission, hp - power loss, % of transmitted - mean transverse diametral pitch (bevel

P, P, P p

-

T. U VI V2 VT VT V, W W

w Y Y

-

Y4

W Wd W,

Sgear), in.-

oil pump los,, hp fastuner tension !oading. lb gear windage power loss, hp pump discherge pressure, psig

- average external surface temperature, "F - average external surface temperature (cold condition), OF -- EH I speed parameter, dimensionless - rolling velocity of faster of two bodies in contact, it./sec or fps - rolling velocity of slower of two bodies in contact, in./sec or fUe. - total rolling velocity (V, + V2) of two bodies in contact, ln./wec or fps (Vi + V2)/2. in./sec or fps - slig "4 vldoc;ty, in,/sec or fps - load, lb - helicopter weight carried on braked wheele, lb - EHD load parameter, dimensionless - dynamic load, lb - failure load, lb gvat touih luad, lb -- effective gear tooth load (K. W, + W).), lb - running load, lb/in. - modified Lewis form factor (spur gear), di-

-

SPM

OF

W

-

circular pitch, in. -

- modified Lewis form factor (helical gear), dimensionless

Qs Q

torque, lb-ft or lb-in. brake torque. lb-ft skid torque, lb-ft - flange torque capacity. lb-in.

2

- modified Lewis form factor (bevel gear). dimensionless - total transverse length of line of action, in.

0.

-

stud torque. lb-in.

7

=

R R

-

mean transverse pitch radius, in. reliability (for I-hr mission), dimensionless

Z,'

-

R,

-

distance from pitch circle to point of load

a

- pressure viscosity coefficient, in3/lb

a

r r

application, in. - radius of curvature of g-,ar tooth, in. -radius of curvature of pinion tooth, in.

- linear coefficient of thermal expansion. in./in.-*F - contact anghe, deg

-

.nrcbability of survival, dimensionless

S S.,

rms surface finish. pin. allowable endurance limit stresi., psi

SA,

-

S, S,.f S S4 •S, S, S,, T

- compressive (Hertz) stress, psi - compressive (Hertz) stress at failure, psi - hoop stress, psi - bursting stress, psi shear stress, psi - tensile stress, psi - torsional shear stress, psi - tempcrature, °F

TI

-

4-2

bearing stress, psi

ambient air temiperature, °F

4

mensionless

normal circular pitch (helical gear), in.

i

_11

.

.- ,,,,, a

rn fam~ ,.,,

,e,,onsn,.,

modified scoring geometry factor kspur gear), dimensionless

- fraction of theoreticai contact (splines), di-

mensionless - incraent, as A I, deg F or A(D,,/ 2 ), in. 6

-

&• a 4

-

mean CLA surface roughness, pin.

incremental lrowth, in./(in.-*F) - deflection, compressioo, or protrusion, in. - efficiency, dimensionless or %.subscripts c ana f, p. and i denote coarse %'idfine pitch, pump, and transmission, respectively - ratio of oil film thickness to surface roughness, dimensionless -

failure rate, hr-'

C

.

a)dynsric viscosity,Ib.-sc/in.1

FS

made lighter, more efficicnt, art d ss costly if it werc

Poino)-is ratio, dimensionless v ..heict] tooth kcad line inclination, deg EHO parpmeter, dintcnsionleus 3 -denshy. sltrg/ ftl or lb/in. a -standard deviation, dimensionless gear tooth pressure aingle, deg - ormal pressure angle (helical gear), deg 0, - ransverse operating pressure angle, deg 0 - gear tooth helix or spiral angle, deg - angular acceleration or devele.-ation cf rotor, rad/icc' W rotational speed, Hz

41INTRODUCTFION 4-.1 GENERAL Thc proper use of this chopter as an aid in the

achievement of satisfactoiy transmission and drive

syscmdetildesgnreqirs aclarunderstanding of asi Te cntcitsrefectthe past in coceps. stvcal generae anhe suggdested analfor tio u-nrdw..rhac sinr te~chnanurs

9

not necesary to conrider interfaut effects. Hiowevecr. the design optimization technique-,. addresscd in this chapter are limited to the components of the powvcr transmission subsystrms without consideration of i be possible ovcrridirig effects peculirar to a given aircraft configuration. The designer must be aware of the significanice of thc to*.al Army environment. Major subsystem comnponeiits such as gearboxes, driveshafts. or hanger bearing assemblies probably will be subjected L,) rough treatment daring shipping, handling, arid removal from or installation on the helicopter. The consistent use of sophisticated or special tools and torque wrenches simply will not occur, even though specified by the designer. Extremes in temperature. humidity, sunli-ght, precipitation, and sand and dirt cotmainwllcurdinbohfgtopac :onamnaion wqill ocur urig bth nigh:topera. tion and lengthy periods of outdoor parking. Pres. tawtr sr laigcupct-crilyn

mal shock. Exposure to such hostile envionments will cu eetdyfrln eid ftmbtms not compromise the mission availability of the hhelecpter. Improper maintenance, tool drops, and rvreo mrprisalto hr osbeas willrs o r. Compro nernstaldesion wustre toslern ando

desiner ustpracicehis kil fro a ovin or The use of geared transmission systems predates irecorded history. A relatively sophisticaled differential gear-drive system was employed in the Chinese

forgiving of such treatment wherever practicable.r

*C.Ln...&I

~

~ ~D.C 10

%An

i

and geared drives still represent the most efficient mthod of power transmission. This chapter is intended to encourage rather than to subvert new and unconventional approaches to old problems. The sole limitation upon incorporation of the unconventional in helikopter drive systems is that the ba:sic rules of nature (la~ws of mechanical physics) are relatively inviolable and should be treated w'ith respect, There is no perfect or unique design solution to a given power transm~ission requiremenrt, and all known designs have been compromised by the individual requirements of the aircraft into which they must be integratid. Optimization of a design cannot e viewed within the context of the power tiansmisints alone; thorough trade-off studies must such factors as suspension, layotit. airframe ,]support structure, rctor control systems, aircraft weight[ and balanc, space limitations, and locations of eniginecs. Almost any known drive system could be

[

\consideir

K'.-.

helicopter and its drive subsystem; thcs:: cleaning solutions may be more than 100'F hotter or colder

tebasic selections of gear, bearing, and shafting repiresent the present state-of-the-art, leve, te tehnoogyas dscused oesnot exclude the future in that areas of uncertainty and limitations of knowledge are emphasized wherever they appear applicable. The mechanical drive system 'tauratiens

`

.

0. 4-1.2

REUIEMNT RQIEET

General requirements arc applicable to all drive-1 systcm configurations. There also are specific requirements that vary according to th~e aircrafl configuration and intended mission (par. 7-1.1, AMCP 70)6-201) and general rc. 1uiremcnts peculiar to particular configurations or arrangements of engines and rotOrs (par. 4-1.3). 4-1.2.1 General Re~quiremnents Ccrtain requirements are common to all Army helicopters regardless of configuration or intended usage. The desired level of attainment of th~ese requirements and their relative importance generally arc specified in the appropriate prime itern development specir-cation (PIDS). Because ne transmission and drive system represents a significant portion of the total complexity and cost of the helicopter. these common requirements must be considered during 4-3

A

P

•-detail

design. Such require ments -- without conof their relative importance - include perreliability, maintainability, and surviva-

output, with smaller amiounts of torque being carried. as the distance from the output incmueses. The wecight of any gear reduction stage is proportienal to the second or third power of the torque. Therefore. itorth-

-- • 4-1.2.1.1 Perfoima|e' The contribution of the drive system to heclicoptecr

while weight savings can be made by using drives in the order ranked -- concentric drives near thc output, parallel-axis types at intermecdiate or conmbining

performance can be defined in various ways. lHowever. the folluwing factors predominate: 1. Weight 2. Efficiency Size 4. Noise le' el.

stages, and intcrsecting-axis types nearest the engine if drive direction changes arc required. There are occasional instances where these rules may not apply: e.g., high reduction ratio may be undesirable at secondary-lcvcl power outputs, such as tail rotor or auxiliary propellers. because of the long distance from the helic:optcr CG. Althou~gh thc total drive subsystemn weight may be less with a low'cr tur-

Ssideration Sfoimancc, bility.

S3.

-0t&-

-.

4-1.2.1.1.1

Subsysitem Weight

Weight of the transmission and drive system is minimized by attention to compact size toge~her and

with the use of high-strenlgth materials for dynamic

qie being carried by shafting, it may becoine difficufqt to obtain a satisfactory CG location due to the larger nmoment of the extra weight of the higher ratio final reduction stage. The specter- weight,%of current helicopter main gearboxes in range from 0.30 to 0.50 Ih/hp for red uctio n r a tio o f 15 :1 to 7 4 :1 ( F ig. 4 -1) . T h e n ext uc •:auc U ||uuubtic uly w il . ... this lrndlcx dr "oin.. .... the 0Z2 level. 4-IZl. TnmsonEfcey(

components that make maximum use of the material properties available within the limits of pr-rmissibic failurc; rates and reliability requirements Superior . = . . oLc.~ . . .. . - -A.. a it . . . o .l ,M:. ~ ~ ~~q ..... low -d ensity m aterial for forged or cast housings also necessary. The latter is import-.nt because gear:'%•are l•_-• ~ ~~box housings comr.prise from 20 to 60% of total transmission weights in current designs.4-.112TrnmsinEfcey The requirement for efficient power transmission is large gear ratios per stage generally SExcessively of such importance that gear types other than those weight. The pinion size is determined primarily •'••.•add Slisted in pai. 4-1.2.1.1.1 seldom are considered -- ? by the torque. transmitted and is relatively ind¢-• for application to the re-Ain po%%cr drive sc~iou.;y gear of the the weight but star ratio; o1' the Spendent member increases roughlyus the square of the ratio. Large ratios per stage usually also are inefficient. mConsidering the sib of support structure and bearoor ings, as well as of the gears thmselves, the gear types themselvgs by increasing weight and power loss worder 1.(for approximately equal gtaonratios) in the follow-cri.t ingThmanner: Concentric drives: epicyclic or planetary devices falr4a4 n ciblt eurmnsSpro Star - ~uiiycotrlinpr~~sngad reversing 2. b.Parallcl-axis drives: rotation. spur, helicalo *Ii'!and horring-f

h-

ris

pt prlxd gabxsi ag rm03 =ducionrtoom51t sg n itrct-a 4:t

o05

bh

o

u

e

r H.4I.Ten

bone

3. flterslwing-axis drives: spiral bevel and hypoid.reurd This ordering reflects the addition to tor-thcs the intended speed change per effects, stageW ofin additional

/,/

"

c-rules*D

.que vector translations and rotations. The con-

S|centric

axis; vhe does nota translation alter tht torque dri-'€ .paralll drive requires of output vector dre with respect to itruto and the intprsrcting-axis driveo-ellers, introduces a new coordinate tfrough tht rotation of the output vtttor with ocspect to input. The selection of drive types should follow the given ranking, beginning at the ouput drive. Also, the plargst reductions should bw•ith kn cloto to the final

Q,

on ""--AH'IG

b

of t

TActEOFF PO*FCR lopt 10" F jure 4-1. Helicopter Main Gearbox Weightos Takeoff Power

" -

J

.:

-

i..

,

..

wc~-I 7*2 " example, with a 1% in-. "train for many reasons. For crease in power loss the life cycle cost for an assuned fleet of 1000 medium helicopters would be increased by $100,000 per helicopter by the extra fuel necessary to perform a constant mission. This is based upon a

6000-hr life, a specific fuel consumption (SFC) of 4 0.65 'b/hp-hr, and a fuel cost of $0.016/lb. Firthcr, the average helicopter lift capability ranges from 5 to 10 lb/hp depending upon rotor disk load.ng and 3I3 C. operational variables. Therefore, a medium helicopter of 2000 hp suffers a useful load reduction of fiom 100 to 200 lb with a 1%additional power loss. Thus, when comparing an alternative gear system of 1%lower efficiency. the basic gearbox weight would need to be reduced by more than 100 lb to compensate for the power loss. The reverse drive efficiency of a gearbox also must I 303 2o00 1000 be watched carefully, as excessive use of recess action gearing can create a problem in autorotation. For inENGINE INPUI POWLR.hP stance, Ref. 2 dcscribes high-ratio recess action Figure 4-2. Power Loss to Heat vs Input systems that operate at high efficiency as speed rePowIm - Typlcl Twin-eMgime-driven Gearbox ducers but become virtually self-locking when operated as speed increasers as in autorotation. A helicopterautootatonaldesent of 15,OUU tlb200 grossftminis weight making an •singits as those shown in Fig. 4-2, as possessing a given efautorotational descent at 2000 ft/mmi is using its ficien4CY apart from a specific power rating. Tise transavailable potential energy at the rate of about 900 hp. mission loacs shown result in efficiencies as. follows: Drive system windage. flat pitch tail rotor drag, and a In l ownr. Efficiency, few minimum accessory loads require appro.6imately Input ower, Efficiency.% 100 hp. A 95% reverse drive efficiency leaves about hp 7 795 hp, a reasonable level to sustain the prescribed 500 74 descent. However, sudden yaw control requirements 500 93.4 conceivably could boost the revirsc drive require1500 96.9 iment momentarily to 200 hp. In such a crse, either 2000 97.4 the rate of S,1.:__:. . .descent . ... would . ..increase . ... 4sharply r . .hor ..some ; 2500 97.8 Iw111i.t,1c crll.gy outllw 'I 'borrowed-frli.the maii.98.0 3000 rotor, with aby slight reduction of rotorvelocity. speed Howbcng Th0windage losses are influenced 98.0 strongly by oil S~~compensated an increase in descent copesaedbyaninreseindecet elrcty HThe ever, if the reverse drive efficiency were 50%, the tail rotor and accessory load would extract 400 hp from the main rotor, requiring almost 50% increase in rthe mfadesent r torrqusing an an almost50 rtr in seeds, rate of descent to sustain safe mnain rotor specls. The power losses of a typical high-speed twinengine-drive main gearb x operating at constant speed might vary as shown in Fig. 4-2. At full speed there is 25-hp windage loss with zero power transmission, and then the power loss due to friction is added as the power transmitted is increased until a total loss of 60 hp is reached for the full twin-engine power input. The slightly downward inflected curve shape is rather typical for most modern, heavily loaded, high hardness gears and antifriction bearing systems, although in som- instances a virtually straight line may be observed. Clearly, it is improper to speak of a ScarboA having loss characteristics, such

viscosity, the amount of oil supplied to the various gear meshes and bearings, and the oil scavenging characteristics of the transmission. Small gear-tohousing clearances, poor drainage paths, and excessive oiling should be avoided. Good estimations of gear windage losses F. may be obtined from Eq. 4-1 (Ref. 3)

P. -

10,, 100)( 10

hp

(4-1) )

where D - gear diamater, in. L - gear face width, in. n rotational speed, rpm This equation represents an application of basic propcller theory (Ref. 4) and is based upon air density at 4-5

"

.

'i

AMCP 706-202 standard sea level conditions which i, 0.00238 slug/ftV. However, for W.IL-L-7809 cii at normal operating temperaturci p -1.748 slug/fil. Consequently. if thi. hel~coptce designer can estimatr. or experienirtally determine the oiliness of the -. nsmission atmosphere, an average p may be employed. Ref. 4 saggests 1, )4.25:1 nir-oil ratio in which case Eq.. 4-1 can bte exprs"cd as / P,. - 2.18(

DI

,

hp

(4-2)

10"-F

Methods of gcai mcsh oiling also affect windagc lossesi 1)iffcrnces of virtually 200% hmve been reported to exist between in-mesh and out-of-mesh oiling for a large. iinglc-reductiorm gear set (Ref. 5). The torquc or load-sensitive centribution of the transmisision to power loss is duc almost entirely, to the rolling/sliding lo!nd-carrying bearing and gear cdcmcnts. Good first-order approximation, for antitriction bearing tosses it mioderate speeds arc given in Refy. 6 and with a little greater precision in Re .f. 7. However. where very accurate decterminiations arc required or where high-speed applications exist, the prediction of these losses requires an understanding of the rrr complex factors involved (Ref. 8). For optimizat ion studie~s and examination of the effects of external dcflections and asymmetricAl loading, there is no substituic for a good digital computer program of the baIc equations such as is presented in Ref. 9. In general, the significant power losses in bearings uc~tut in regions of appriciable .ýontact zone slip. uýSIC re the most efficient for most applications. Radidiy loatded ball bearings are lower in efficiency. whl b, aular %cntact (thrust loaded) ball bearings C.'N64~ sivi~iti.,i..rty greater friction losses. When the 1; latar aft: used in verry high-speed operations. their r'-iwem 11-)3characteristics dcttrioicite sharply as ex* cc~silve :ctifug.11 and g.-yroscopic forces affect ball ýUICeM-tiCl. SItandurd tapcrc.1 :iollr bearings (conical tolling ticincints) ire the 1kiht zfficicent of thesc fout. typres. ia~becaosc cf the hea.'ily loaded cone rib that'.ý in id~contact whiit ;he largt cod CS the i,ccnical roPliti. A. ti,- type os' anpuhat coniact cylirt. drical rci.er (Ref. ;01 rmi), offir hdvami.ites when fuloy dcvetoped. In Nhs ty~ve, ?ý)c co*ne , b lo~'d is theorcti,_-Oy reducA;J. bjh iincreascq s-lidiril' rttsuIs V~ nCm or Wt~h of t'ic -jic-rellcr co-aaci~s- Expcticivre Is '.nsuff!.icnt f,ýr L-valuel,,i'ol &Itz. relative iicat utnrnAtioil. bitt irdicatio, -A are O~a! it wit! pyrov' it,, b::

.'

A-.re

effincrt, than iht. staaCurd tupcretJ zolicr znd

that itA !osses will approach those of the arigular co.)tact bitll bearing. The plain journal hearintg is unsuitable for heavily loaded so liansmission uipplications (see par. 4-2.4.2), although it may have satisfactory application in lightly loaded accxs~ory uses; the power lou with this type of bearing is at least twice as great as for any of the previously mentioned types (Ref. 11). The hybrid1 boost bearirg (hydrostatic plain thru-st bearing in series combi~nation with the stationary or rotatifib ring of an angular contact ball thrust bearing) has been evaluated experimlentally (Rcf. 12), but no helicopter expecicnovv is known. The experimental work indticated a friction torque for the hybrid ocaring of roughly twice that for the simple ball thtust beaiing. The accuracy of the calculated losses for bearings is dcpendcnt upon proper installation and application dcsign, Excessive preload, due to insti-llation and/or the~rmally induced. c.:n camily result in doubling thc friction torque arid also will have detrimental effects upon fatigue We and reliability. The power los~s in gcaiing is ti~iiie an involved and cnrvr~ICiit neirte~efcwih ranking, listed earlier in tihs paragraph hold true fmn basic efficiency also. There are reawinable ranges of gear ratio for which specific types if grer drives arc bcst suited; when thi:~ upper limits are exceectce. the resultant power loss is apt to iincrease to a level where superior overall efficicm~y may be obtained by reventing to two stages of lowcr ratio. Tie suggested optimrnu ratio ranges are: 1. Conecrntric - epicyclic, 3:1 to 5:1 2. Parallel axes - straight spuir, A1to 2.5:1 -single

hclicals, 1:1 to 3;:1

I:A to 10:1 3. lnterecting axes - spiral bevel. 1:1 to 3.5.1. A nuintibr of assumptionas are included in these suggested ranges. The lower limit for the epicyclic assumes a simple u(oWirevcrsing) design. The planet gears bmcco-ne excessively small so that the sun driver ten&~ to uct like a specd-incrrasing drive *ith a long. inclficicrnt a.-c of a oproach. At these low ratios the epicyclic system becomes planet-bearing-capacity limited in iltht it is difficalt to fit sufficient~y large bearings to carry the nececssary load if the gear teeth are stressed to satisfactory levels. The upper limit for the cpicyclic represents a reasonably designed, four-planet idler systemt whose weak point is tht; tcndenzy for pitting of the sun gear. Be-caust the dasigni is sun-pinion-diameiter limited (to acc--ptabic Hertziart stre~ss levels) and die planet gears art; i'~ur tiiwý.5 the diameter of the sun, the. system is rapitdy becoi-aing inefficient from a weight stand-

do

06M'

_AMCP

point. Th: carrier structure, ring gear, and ring Sear housing (ir used) are excessivci,, heavy. If used in a final drivc stage, the low speed of the plane bearings x rotor speed) is not conducive to the formation of adequate oil film thicknesses for good performiance even though th~eve is virt.:ally unlimited space for high-capacity planet bearing designs. 'The upper limit of th siryl helicals is based upon grcatr than 15 deg in order to minimi7e the thrust

the subrtgimne or microclastohydrodynamics (Refs.. 13 and 14). Fig. 4-3 ieprescrits the salient cnarficteristics and interdependent variables influencing f tbroughout the rainge of specific film thicknesse. The ratioN)isintroduced to give physical meaning to thesw regimes in terms of the roughness, or surface finish of the activo tooth profiles according to

comonet o ýh tothloading. The extended upper limit given for herrirngbonce designs is based upon the use of high helix angles (4'AM 35 deg); thus. mayxim'im

where

I(9/4

Iadvantaf

h

h

6

olfl hcnspn -. mean ceiflerline average (CLA) surface

r istaken of the thrust component cancella-

tion that permits atlainnient of very high face contact ratios which in turn permit some reduction in the profile contact ratio with an attendant reduction in sliding velocities. The upper limit for the spiral be-vcl gear ratio reflects the use of approximately a 90-deg intersection of axes and is based upon increased losses due to excessive sliding velocities in both the art of approach and the recess. Tbc lower limit f'or spiral bevels would be applicable for overhun-gmounata

7UF.X

mEltrIaIIUII

Ar

4A.

aLA3IIa%,ULIaxcs

full-straddle-mounted.,9O-deg-axis systems cannot be ,) accomplished below an approximate ratio of 1 4:1 for 9'0-cleg axes. In all cases the lubrication is limited to low-vis-

* *loss

cosity synthetic turbine oils with ro unusual additives. The specific. efficiencies obtained in gear meshes are basically consiecred tu be represented by analogy to classic physical roncepts. The friction power of P, of sliding bodies in contact is given by Eq. 4-3. (4-3) P1 WV~f/550. hp

roughness, pin. Region III of Fig. 4-3 chn be ntglected fMr helicopter transmission components. The entire rcgion is defined by classical hydredynamics; and the properties of lubricant viscosity, xliding velocity. and load interact to build a supporting lubricant film) that complcte!y separates the load-cart-ying mechanisms. The observed friction is primarily dieptrident upon the viscosity of the supporting film. Region 11 actually extends (on a submi~croscopic scale) into Region 1,but the true importance of th region is that it represents a transtional phase that "nly has become defined with enigineering s-* niCi "P.n.e within the past decade. The pressure distributions within the loaded gear surfaces are considered basically Hertzian, but the Irilm thickness is deptrndent uposi the additional faýtcrs of elasticity of metals and the property of greatly increased lubricant viscosity under the H-ertzian conjunction pressures, The observed film thickness is known to increase with increasing entrainment or sum of rolling

where W

V,

load. lb

CO3

f - coefficient of friction, dimensionless Minimization of power loss would simply seem to re-

FIAL SIALE

manner with thi intensity of load [expressed as a

*

~~~and factors descriptive of the lubrication regimc~ and the lubricant itself. Lubricant regimes sit classified loosely (from thin

~i-lm separation to thick) as boundary,. elastol'ydrodynamic (ENID), and hydrodynamic.

FIRST

TA5(

-

-PLIN

-

*-8ISPFESNPU1--

LAN

quite attention to minimization of i andf. However, the apparent quantity f varies in a very complex

comprcssive (Hertz) stress Se], the slidinr velocity,

HTYDODYNAMIC

LENDCR

BOUNOAR

sliding vclocity, ft/see

OilRO1GEARS__ BARINS.. IQ~L

MAIN______ ROO

-

5-A '--

I

RAICSIIGV

I CRAC

AI'PA1111T~IFANCEPA IfCREASING VISCOSITY.

Q AYSAOBEATR!INC

lubricant

but the following approximation is uselul. In this of definition, the boundary regime includeE

-~syst.-m

-4-7

BEARING$

FIM5.5

IIS

C-HNS

AI

Figure 441. Lubricatlon Re~gimes

H'

I

i.

AMCP 706-2D2 velocities VT(VI. V, + V2, where V,, and V. are the velocities of the bodies in contact) of the loaded bodies, the lubricant viscosity at thc conjunction inlet, and the pressure viscosity coefficient of the lubricant, and to decrease slighily with increasin~g toad and V,. f he largest values of this region represent full separation of the loaded bodies, while the lower values permit some mieial-to-metal contact of the asperities of roughness peaks. The most widely accepted cxpres~ion in use today for EDH film thickness hE (Rcf. 15) is

HE tZIANORV)

LIISRICANb; L CONTACT

PRILSSURi DISTRIBUTIONI

V

/

..

LURh1

.*--

.-

-

.-

~~ V, - VELOCITY Of BODYTI

hE -

2.65

5

31,pn

/..0X

(4-5)

-

or HCOY 2 ELO0CITY

Figure 4-4. Elastic Body Contact Pressure Distribution and Interface Contour where tl'c three El-D dimensionless parameters are G =x'(materials)

u - 10F

(speed)

W

(load)

=

W~~

O t

and in. E'-combined modu'us of clasticity, R = mean transverse pitch radius, in. PT= mean rolling velocity (VI - V2j)/2I, in./%"c W = running load, lb/in. a =pressure viscosity coefficient, in.'/lb A, =dynamic viscosity, lb-sec/in.3 A physical senise for hE is shown in Fig. 4-4. Eq. 4-5I,~ is isothermal, in that it does not treat the effects of material heating in the con~anction, but is believed to hb rosacnngkv agpeutgt. I,*% f^ V /

-I

*

-

j

V

%,o1-,eS , f ,,t1t,ast

--

The observed friction in the EJ-W regime is prim'rivduc to viscous shear 4J the lubricant in the high-pressure field of the conjunction, Much experimental data exists to relate friction values to certain dimensionless parameters. Most take the form shown in Fig. 4-5 (Ref. 15). Such relationst iips hold for constant values of surface roughness and lay, and for specific lubricant types. The motimportant conclusion from these data is simply that friction is relative~ly low - on the order of O.u2 to 9).04 - for components operating in Region II. A mre detailed analysis that considers the thermal aspects of 131-I1) solutions as applied to simple involute gears may be found in Ref. 16. Region I, defined as boundary layer lubrication, represents conditions that predominate in the lowerspeed com,'oncnts of helicopter gearboxes. In this 4-8

3

O *

.v.k

(d. v V, V.

Figare 4-5. Friction Parameters-

Coefficieni vs EHD Region 1andi k

region, f may be influenced significantly by interaction of asperities in the rubbing load-carrying elemeonts, be they gears or bearings. The thinnest of films represemited in this region may be monolayers of lubricant products that either adsorb or adhere to the extcrior molecular s~irface of the metal. The variables influencing friction include the chemical composition and the interaction of the metal and lubricant combination and the roughness, lay, and texture of the surfaces with respect to their rubbing directions. It is at the lower speeds that very noticeable differences exist in observed friction between the arc of approach and the arc of recess of involute gearing. Fig. 4-6 depicts a very low speed mcasuirement of this pl'.rnomenon involving a spur gear set of minimum atiAinable profile contact ratio (CR) (Ref. 17). The

Although the cxpenimcnts cized were. conducted in bcwundary lubrication ccnditicns that yielded much higher f-values for teeth of coarse pitch. it is int twresting to note thiat the f-valuc obtained for gears of finter pitch was in the range of values expoctod for El-D lubrication. Thi3 iliustiatcs th-t iinportvaice of using gewring of relatively tine Vi'tch to obtain maximum efficiency. In addition, it should be not'nd that the apparent differences in friction betwoens the awvs

very low CR is employed to study the extremes of these approach ard recess portion effects without introducing the data confusion that normally would ocin the zones of double tooth pair contact. Fig. 4-6 *cur also illustrates the driematic improvement in f that results when the contract ratio is increased. The torque Q and pitch diameter D., being held constanit, the higher contact Yatia was achieved by changing the diametral pitch Pd aind the number of teeth N.

-0CR

-~

CR

a

10

Q 130 lb-in.

Torque

:~i9.11i

16

_N

-0.14

1.74

4

0.1d1 2.

1.03

2I,

_

_

_

_

p._ _

5.85f

5.55

rpm

~~1-

40

_

.

.-) CD,

~0.06 0.04_____

0.02

-12

4 0 -4 -8 ANGLE OF ENGAGEMENT , deg

8

NOTE: AVERAGE f 7 0.081,

12

10.034

Figure 4-.Angle of Engagement 4-9

F

*...i II

of approach and recess art masked completely by the avraing effect found in the zones of double tooth pair cont&ct. The basic tiends of friction change in approach and rcs action arc still valid i:i lubrication Region II. Fig. 4-7 (Ref. 18) represents experimental data taken at corasiderably highti values of ourface sliding velocity, load, and lubricant viscosity. Alth..ugh the use of a dig'tal compute: program to examine the many instantaneous contact conditions that occur b. a pair of gear toeth rolls through mesh is the more precise meitod of catulating efficiec:cy and studyiog detail design variations, c fexrek-i -t - -may be obtained by using average values for f. The percent power loss P, of a gear mesh is expressed (Rcf. 19) as: p,

L X 100 96, Mfeatures f

(4-6)

where M = mechanical advantage, dimensionless Values of M for various combinations of pinion and gear tooth members are listed in Ref. 19. For the ,4) o Fi . q-6, ahe •i ia: lar ( 4 u, ia~h ' , . 4.6

P

X 100

1.76%

(4-7)

II~

and for the fine pitch gear set (P4

10) the P, is:

2-.034_ Y0100 - 0.40%

/' P,

(4-8)

8.5 Their corresponding efficiencies V are then simply ij 100;- P, and we find. 9 = 98.24% (4-9) - 99.60% where the subscripts c and f indicate coarse and fine pitch, rexpectively. The frictional differencer noted for approach and raxwzone of involute action are characterized, with respeci to the driving member, by the rolling and sliding contact motions being in opposite directions to one another in approach but ia the same direction during ro.ms motion. The scensitivity of friction to the lay and textural of the mating members in lubrication Region I is s.jown clearly in Fig. 4-8 (Ref. 20). These data represent the results of experiments conducted on a gearedisk test machine with 3.0 in. diameter, 14.0 in. crown radiua,, case carburized and ground, con0 93a c . s'-smab!e e•iec ode vac . n steel disks. The circular ground data were taken with disks ha%ag a circumferential finish of 8 pin. and an axial finish of 16 gin., while the cina-ground ditks

[0..

4O~

~

-ROSS

" •

GKOUN9

300D --

0.%

3.'

''%--

0

500

lDO• 1500 SLIDING VELOCITY V..

2000

ZSO

±

Emm

--

--

SLIOINCVELrClI Y V,, -M.k,.€•

Figure 44. Effed of Surface Texture and Figui 4-10

4-7. Coefficent of Frkilton vs Sliding Velocity

Lay on Friction nd ScffSag Beh• vr

"

V.

*limit

-~

bad a drciinforweti ftaub of 16 pin. and va axial finis of Spin. Conm'queud.ty both types had identicaO reduced finWis number &-values and benice vii tualuly identical A-valies, but the cross-ground data news was RcEO463 irn both type., and the cross-pound disks weas prepared using grinding techniiques nor-naqly used for spur gear Waoth manufacture. A constant ratio of IV/VT - 0.556 wias represented, sai the lubrication was jet auppNWe MIL-L-7806 at 190*F. Theretome a particular point on agear mesh wher V, 3S.6% of Tr is represntmd on this ftzgmat u a linearly increaed gea apecd (rpm) by moving from left to rig1bt -ýith inremasing Y1,. Unfortunately, then. data do no reflact a oonstant load, but rathar the baa6 as delrined by scuffims. The designer shuld be aware that udrfxd lubricant condition there is little he can do to control Sat low speeds asid from refining the surface finish, Additiona-i powe: lowa sources of a transmission systemn include the accessories anid the oit punip. Accesory power requirements are fixd by the individual helicopter requirements. the exact typc of accmoasy 5nvolv')d. and the demand or duty cycle memiPaAMtmnfib

mmrv u

,nwewr

vAmnnrap.

ments igdiscussed in Chapters 1 and 9. Oil pump loss P,, is estimtted adoquately by the aimple equation: Fpk '7 (41) PPhp qP where oil flow rate, gpm discharge pressure at p;Anip outlet, psig converion constant for units, 5.83 X 10-4 efficiency of pnmp (generally from 0.5 to 0.9), dimensionless For a 20.gpm system with -jregulated discharge of 60 psi, the pump outlet pressure would be 120 paig under typical cofiditions. For an assumed pump efficiency of 0.5, the loss would be: (20)(120 X 0-')result (5.3 (2)(10.) (59 1-) 2.78 hip (4-11) (0.5)through F

-

p kr q. -

4.1.2.1.1.3 Size Compact gearbox size is important in the achievement of low subsyst;= weight bemause Ibc housing or casinj that enclome the dynamic components contributes a significant proporticni of the total systam However, compaction should not be emphasized to the point of causing excessive oil churn and windage loome to the detriment of efficiency. Ref. 21

)weight.

suggests that cide clearances of 0.5 in. or pester between Seats and cauing wall* reult in negligible loom due to oil churn. The required clearance[ between casing wail and Sear outside diametpr inarc of conformity, speed of &ear rotation, oil viacosity, amount of oil etted on the gear mesh. &nW the amount of run-off or drainage ouil in the locattion at question. There are no formulas for calculating satis-A far~ory diametral clearances, but some successful design applications have amployed valu&t of 4-. proximately 0.5 in. for 2000-fpm pitch line velocities and 3.0 in. for 25,0I00fpm vcocitics for 1801 dog of conformity and kinematic viscosities below 10 centistokes. Evtn at these clearances. it firequcntly becomes necessary to provide scrapers or some mean to retard vortex acneration and lo Waized recirculation of the oil. When wet sump systemns are employed, suffic.ient vei-tical sprce must bý provided to keep the operating oil level (iaucluding th4& Avr~iawd or foam lazyei) Weow the gears and beerins

i

j

4.1.2.1.1.4 Noase LereleV The fourth jxcrforrnince criteria of low nao" iceW &--&a2--

~

--

~---

-

noise is 5Isuailj of imporAncc cidy ii, relation to crew, and passenger comfort levelt, whale rotor ttid enoine noise arm the principal contribute,-& to thc auiral dotectability of the helicopter. The funda.mental gearmeshing frequencies. which range from 40 to 22,000 Hz in present-day gearboxes, are the p'impty sources of noise. Refs. 22 and 23 identify thc magnitude of the problem for two helicopters. Tht overall helicopter configuration and the rfaulting number and location of gearboxes dictate the areas nlTfeed by noise; e.g.. a tandem-rowo: helicopter may have its forward transmission located above the crew compartment. resulting in less a'avorable noise conditions in the passenger area, while the reverse may be truer gabxa for a single-rotor configiuration. Gear noise may emanate from the gabxa of forced or resonant vibration of the housing or cases. It then reachca the.-iz-r passengers either direct airborne paths (window&, access panelk, or door seul) or through airframe structural pathways connected to the gearbox mounting system. It is far more efficient and desirable to combat suo& noise at it&source rather than to rely sdely upon the use of insulating and soundproofing coatings or blankets in the crew or passenger compartrients. The latter measures generu~ly add more weifjht than would be needed to make comparable improvement in the problem at its source. Sound insulation also in-I cref sw4 maintenancet man-hours duc to the need fot 44.1

*

1

temoval Of the material during airrrame liaspections.

Also, soundprooflng efforts often are deleatod when torn or ciI-soked mateirits art removed and never ri~eplcod, The uae of elastomeric isolation mounting devices at the gearbox and hanger bearing supports is highly effective in rnduc'ang structural noise. Airborne noise should be minimized by clintisating any housinp or case resonance through use of proper wall thickness, shape, or internal gear or bearing quill attachmcint methods. While it is virtually impossibl* to calculate these conditions with suffkienrt accuracy in the design stage, they are itlatively easly measured during initial component testing, and corrective redesign then can be undertaken. Modifications in the shape or involute pr ofiles so zs to change drastically the fundamental and harmonic noise content have been investigated analytizally (Ref. 23). However, the slight variation in profile requite to achieve theoretical improvements was judged boyond tLn. p~resent rltinufacturing state of the art Soni't methods for approximate analytical prediction of resonant per* formance for relatively simple structural housing * :ha~ L~ dvaniMd an RC.24 It is not certain that the elimination of all interactin& vibratory and resor..rit behavior in the various gear meshes Is entirely beneficia; i.c., some sacrifice in the efficiency of the lower speed (boundaty lubrication regime) meshes may result. The effects of axial lubrication upon the reduction of tooth-meshing friction is reported int Ref. 25. Although the engineering field of gearbox noise generation is imprecise as yet, there exists considerable general knowledge that can be of practical benefit to the designer. For example, it is known that hnigh contact ratio gearing and finer pitch sizes pro. duce. ess noise than their opposite counterparts. Similarly, helical gears ame quieter then ttraight spurs; spiral bevel gear are quieter than ctraight bevel or Zerol gpars because of their greater inherent contact raitios reduced dynamic increments or waste loads, and increased smoothness of operation. In addition, increased gear tooth backlash and clearance can help to maintain subsonic air ejection velocities fromt blgh-sprtd meshing teeth (Ref. 26). Viscous films between tAtionary bearing rings and housingsi can provide suffikient damping to roduce vibratlon and noise propagation. Coulomb or dry fric* tion devices have been successilul in damping resonat modes in Sea rims and webs, as have high hystaresis materialis clad or bonded to shafts and webs. onthe 4-12.1.2 UdhIEfy A complete general discussion of reliability con4-42

cepts is contained in Chapter 12, AMCP 706-201. Thin paragraph, therefore, deals with specifics as related to mechanical power trantsmission ccmponents of the transmission system. Concept definitions and numerical values for quantitative reliability indices generally are speciriod in procurement documents and, with increasing frm quency, in PIDS. For transmission and drive system design there are two types of indices: I. Values for such characteristics as mission reliability, flight safety reliability, and system reliability rot the entire helicopter (us.ually for a gi-een mission and operational enviroanmcnt). Typical values and methods of expression might be, respectively, 0.90 to 0.99 for ont. hour of misL,.ion time. one failure per 20.0(N) flight hours, and 0.70 to 0,80 probability per mision hour of no system failures requiring unscheduled nainteniance. Because the reliability of the helicopter is a composite of the reliabilities of the individual subsystems, individual reliability levels must be assigned %stargets for the design effort. As an example, assumei the Request for Picoposal (RFO) specified values of 6.98 and 0.9999 (one failure per 10,000 hr) ur minkimiu and safivt

criteria, respectively, The Ai-

lowable apportionment for the drive system would be dependent upon the complexity and type of helicopter, but typical values for this apportionment could well be 0.999 and 0.9999. respectively. Techniques for desigri to these requirements will be addressed in par. 4-2. 2. values for suhsy,6tems and component~s for such characteristics as reliability after storage and mean time between removals (MTBR). Typical values are a maximum of 10% degradation in mean time betweenA failures (MTBF) after storage for six months in approved environment or containers and 1IS0 hr MTBR, including both scheduled and nonscheduled removals. This type .-f index is directly applicable to individual subsystems, including the transmission and drive system. The MTBF values subject to the 10% maximum degradation limit lite those specified implicitly or explicitly in Item 1, i.e.. the 0.90 to 0.99 mission reliability carries the recirrocal meaning of a 10 to' 100 hr MTBF, the one failure in 20,000 hr for flight safety is a statement of 20,000 hr MTORE, and the 0. 70 to 0.80 probability of zero system fail urs per hr implies # 3.3 to 5.0 hr MTBF. The MTBF levels corresponding to the drive subsytcm apportionment in Item I are 1000 hr (0.999) and 10,000 hir (0.9999). respectively. It is important that the designer understand that MTBR arid MTBF criteria discussed previously ~ interrelate in a unique fashion with the subsystem design reliability when a finite timne between over-

(

'

(

.

F)

A&IcP 705-=2

hauls (TBO) is selected. On the astiumption that. n~l transmission and drive system failures are of sufficicnt magnitude (and detectability) tn. force a mis-

cancellation effecta, leaving a not MTSIF value dofined with sufficient precision to express design reliability rc-quiremenrts.

to be aborted, and further that no TOO (scheduled removal) reqjuirement is imposed, then it follows that the MTBF owMTBR. The imposed requirement of a 1500hr MTBR would necessitate a failure rate A!• 0.0O004 or a one-hr mission reliability of 0.9933 for the tra 1smission and drive system. The relationships satisfied in the preceding statemenits are:

Yh "ft of a given set of gearbox failure data to teepnniiasmto eel eti hrc tcrestxcsofth asysumtemion reveals. cfteoertaingcar loprtng, a If sufihenl syroum in*q teimeci ofah cetwerotrndwlb bevdasaodfnt erottedwl eosre sann cnsatorireigfiuert.Omoroetisensilive (wearing) compontnts eventually will begin to dominate the failure picture as the end of their tueful life is approached. Generally, those componenits operating in deep boundary layer lubrication regimes W19I be the first to influence the picture. lit such instances the wear will progress to a state wherein conditions become fa'vorable for the occurrence of a failtire mode indigenous to that new set of operating conditions established by the partict'li wear state. Fig. 4-9 gives an example of a well-developed or debugged gearbox with relatively low TRO that satisfies the rancom failure characteristics a:curately dcflned by the exponential distributi.on (Rd. 27). Fig. 4-9 represents a gearbox with a 350-hr

Ision

I h and using Maclaurin's form of Taylor's formula R xp~(k MT

.B

(4-12)

where R =reliability (for a I-hr mission), dirnensionless However, if a 2500-hr TBO level were to be assigned, the I 500-hr MTOR could be satisfied only by a higher reliability number (lower failure ratte). The rclation. . - - --

~bnp

Imuy uti

-A

flf IL-. L.. U..f

rpirbbvd ivi bi.0-ha, ivjjaffuw.

3MB

_A< AAIBR

-

XTaO

-

-

0.00067

___00

--

o.00040 X ý.o

•! 0.00067

-0.00040

) -

sn..

ii

. Cý

fý

a50

I

1,0l

-

I

-1

_

scheduled removal frequency, the resulting MTDR is (4-13)

2500

Therefore, A. !5 E •B~ :5 0.0027

scheduled tria

reliability of 0.9930. However, due to the 350-hr

TO+Xr1 B

ABO+IM

AMBR-£ =

_..--

as.

lic~cthcnc-v TBFmus be~ 304hr riditecor IT# I "A responding ireliabilty R ;± 0.99973. It is paradoxical that such factois as flight safety reliability and the increased cost of overhaul of a badly degraded gearbox (extensive secondary failures) may fix the TBO interval at a level that in turn rm quires a significant increase in the required MTBF to achieve a specified MTBPR. Relationships defined in this maniier are tacitly assumed to fit a simple exponential failure distribution. This not only simplifies the arithmetic involved, but enables the designer to think directly in terms of the inverse relationships between the number cf detail components comprising the sub-. system and their intrinsic failure rate requirements. Although many individual gefirbox componients are better represented by other distributions - e.g., Weibull, gamma, and lognormal - the averaging effect on the subsystem as a whole is such as to invite

204 hr.

Fig. 4.10 from~ Ref. 28 represents the distribution evidenced by a %i~el dc-bugged gearbox with relatively high TBO. In thin caa- the upward inflectiorn or concavity of the data points reveas the strong influence of component wear-out as the TBO level is approached. ISimilarly, Fig, 4-10 represents a gearbox with an I100-hr scheduled TBO, an MTBF of SUP' hr, and a one-hr reliability of 0.9998. Ir~ this cast the scheduled removal frequency results in an MTBR of 904 hr. Obviously, design MTBR and MTBF values sart dependent upon the desiner's knowledge of repro 5entative failure modes and his ability to assign reasonably accurate failure rates. Extensive test and service experience with analogous componewts and subsystem elements is essential in arriving at realistic predictions. There is little published literature to aid the designer, but the selection of components with known lower generic failure rates always should be the objective. It should be recognized that generic failure rates (indicativz of the intrinsic reliability characteristic of any component) cannot truly exist apart from the envirvnment in which the component functions. Some insight into these environmental inlucrnces is given in par. 4-2.1.1. 4-13

&mill

-

0

--

GEAR BOXES SURVEYED -.173 32

-

32

26TB -

OF FAILURES . 34

26 TOTAL TIME

-

-A

-HNUMBER

-

172,431 hr

,f71ti--

10

0-0---1

00

01 0 00

-ii

0 12________

12E 17.67

4T TOA

0

OPERATIONAL EXPOSURE TIME;, hr x 10-1

Number of Failues a Hewns Simv Overh. - MATSF ow5001 r

FjS449.

2r

0

0W8

-'7-

04 121010 2

44

6

OPERATIONAL EXPOSURE TIME, hr x 10f Flgurs 4-10. Number of Failures vs Hours Slae Operatin. WINF - 9W br

based upon th. profile of !he design mission; i.e., for Satisactory failure tate estimates for rolling elea mission profil, of n discrete values of bearing load nment boarings may be derived by seeral wethods. P, each occurring for a perorntue a, of the total one simple but sulficient value is: (41) operational time: IN-

S L

I

5 /~h-I(414

1/3

probability of siurvivl, dimensionless AMC i(4.15) - duuign life or scbeuled renowvatime 100)l (TDo) br where The value for Sis rsd fromthe curve in Fig. 4-11~U corrseponding to doe rtio of dbo design life L to thw0 Although in ma-iy instances the life-load exponent life olo at wbicb i0% of the beauing population will more correctly can betaken as 4. thecube value is rad Tkin#10 value for a gienbern gis determined recommended for determinations of failure generally root for the from the bearing msnufacturW&' data rate A. mean cube AMC load. The applicable AMC load is 4-14

-

(.

XA&P70&202

-

0.6

will have failed by emrplo,41ý'a dispersion exponent, tatii'c of typical h fcopicr geat performance whome

I

4 Ii *1

excellent quality control generally results in lower 6ispersion. Very steep slopes frequently Lubrication Regime 1. The life value L2 for 98% reliabilty may be read from Fig. 4-13. a Weibull ~plot. The resultanit failure rate A is then:

*population

0.4

1

1typify I

(4-16)

X - 0.02/L 2

The mean value may be taken as the 50% or median rank foi such steep slopes without losb of significaid accuracy in using Fig. 4-13.

0.1 9S.9)

99.9

go 98 14 9 9..4 99 PROBABILITY OFSURVIVAL S.

9)

94

90

FSpot 4-11. Pro-babillly of Survital vs L/810 Ratio

-

2D

copter appfications will exhibit pitting as the lifelimiting failure mode (pat. 4-2.2.1). The life-stresa-to-4. lationship for gear teeth is far more complex than for arc ar ighe, more types of ni~ctals and heat treatmctfare prgevaen, and the elastohydrodynamic and cemical effects of the lubricant are known with less precision Available simss-hfif curves more often than not are aacd upon the mean pitting, or spalling, endurance of an unknown statistical sample (Ref. 29). Intensive is underway by many organi7.ations (ASME R esearch Program on the Relationship of Lubric8-

I

____

IS

01I0 PLIGiecls

Figre 4-12. Spalling Life vs Hertz Stress WEIBULL SLOP", 5.5 -1

Kresearch

-

-O

ton and Fatigue in Concentrated Contact, for

70

preparation of exml)that should result in thetrdifi"ehr

60

N'~~~ pit-1i

that consider lubrication regimes, materials and metallurgy, and sliding speeds. The AGMA data of Ref. 29 reflect usc of a stress-lifec exponent of between 9and 10, whereas values of 5 have been reported (Ref. 30) for operation in Lubrication Regime I (Fig. 4-3). However, in the A!scnce of well-documented, statistically-significant, test date, the AGMA data should be teken as represcntativc o~f most gear applications. Use, of the RMC value of the Hertzian or compressive stress in the contac! area to obtain the

1tos-1

4o. ---

~

*j

4-

1

'i

-

e

23 I

10--f.-I

4

-

mean spalling life from Fig. 4-12 is satisfactory, al-r though the quartic mean level has be'n shown to offer excellent correlationi in Lubrication Regime 1.

7)

There are many suitable techniques for reducing

this life to a value of failure rate A.One rather simple method uses standard Weibull paoper to reduce the mean life to the level L 2 at which 2% of population

-

Th0------ T

200

Geat failure rates can be determined similarly from design stress levels. Properly designed gears in heli-

gK

2GEAR! GRADE 411.02 AGMA

300-

-

2-

10'

0 $PALLING LIFE. cyoles

Flpre 4-13. WelIulI Me - Spalifq Life vs Gear Popula"o Rank

0

AWP 7Q2

-

41-1,1.1.3 Maaimalahlty A gcnoral discussion of maintainability may be found in Chapter i, ,k MCP 706-201. This discussion, therefore, treats considerations relating specifically to the transmission and drive system. The basic concept of maintainability often is expressed as a requirement for a specific or maximum number of maintenance man-hours per flight hour (MMH/FH). Army helicopters of a few decades ago exhib~ted values as high as 35 MMH/FH while helicopters in the present Army inventory have values ranging from 0.5 for the OH-58 (Ref. 31) to 6.5 MMH/FH for the CH-54 (Ref. 32). Small helicopters as a rule show better maintainability values than the larger, more complex machines. However, the values for any given size of helicopter may vary by 300% depending upon design variables. Current RFP requirements arc in the range of 5 MMH/FH for medium-sized, twin-engine helicopter for organizational, direct support (DS), and general support (GS) maintenance levels combined. A value so stated must be apportioned in turn (par. 4-1.2.1.2) to the various subsystems to establish individual design goals. Achicvcmunt of sadisfactoui-y -aitai•-ability levels is dependent upon two factors: I. High component reliability (par. 4-1.2.1.2) 2. Ease of maintenance, Mainteinance is generally thought of as comprising two categories: nonscheduled (due to random failure or accident), and scheduled (due to time change of

or clamps. !t also should be emphasized that true leveling of •lte helicopter is seldom achieved for comnponent change at the direct support level. Therefore, when heavy components neceLitate the use of hoisting devices, extra care must be taken in the design of algnment devices and structural cicarancs, so as to reduce maintenance effort. External shaft seals always should be essembled in easily vemovable housings or holders to permit bench changing of the seal element. Squareness, alignment, and cleanliness practices all are critical to the proper performance of a seal and are difficult to achieve when the seal element must be changed in place. The shaft upon which the seal operates also must be easily removabie because good practice requires that the shaft be slipped into the previously installed seal, allowing the use of adequate shaft lead chamfers to minimize the danger of seal damage. It also is desirable to have the shaft engaged with the driving spline or some other guiding device prior to making contact with the seal to prevent excessiv sidc loading of the seal. The attachment of all external components should be such that one man can remove all fasteners and similar items. Two examples of poor design that require unnecessary manpower are: I. Bolt-and-nut fasteners through structure where one man cannot reach wrenches on both elements. Tapped holes on nut plates are proper solutions even though a larger number of cap screws may be required because their rigidity or strength may be lower

wear-out components, interim servicing, and in-

than that of a bolt-out joint.

spections). However, all maintenance. concerned with componer.t change is discussed herein as a group. The generic failure rates of many external com-

2. Components that one man must hold while another installs fasteners. Possible solutions include the use of guide pins, longer pilot flanges, slotted

at,,ch Donents in the drive sytem rre

,-.aran-.,, hnl,.

that !h, cnm-

r-t,,nn

n. ,rt

-r.n..

.. h

". •-

r.;•

portents may be ranked in order of required fre-

tion devices to secure the component temporarily.1

qucncy of removal or adjustment. Components such as hydraulic and electrical accessories, rotor brakes, shaft seals, external hanger bearings, and drive shaft couplings require relatively frequent inspection or maintenance and must be designed for ease of removal and installation. Accessibility is the key criterion Such components must not be located too close to one another, and adequate wrench clear"ances must be provided for standard tools, Subsystem components such as g-arboxes are generally not maintained at the field or direct suppori level and, therefore, they must have simple and accessible attachments with structural clearances adequate to permit easy removal and replacement. The usc of integral guide pins or tapered dowels is rccoinmended in any case where heavy components must be aligned for the installation of mounting screws, bolts,

Components that require a specific orientation to function correctly should be designed so that they can be installed only in that position, if possible. When this is not practicable, as in dhe case of some Government-furnished electrical accessories, decals may be ustd at the pad location to provide insiallation instructions. Examples of one-way components are seal housings with drain fittings, hydiaulic pumps that require line connection fitting orientation, and bearing hangers. Components that require tight-fitting pilot bores or similar devices should be provided with jacking pads for removal. One man can operate two or three jack screws (tightening each one a little at a time), whereas their omission might necessitate the use of two men to pry simultaneously on both sides of a component (and possibly a third man to catch the

4-16

r

_-A

*AMC, - -component *.

706-2012

when it breaks f(we). Proper performance of scheduled maintenance tasks such as inspection and servicing is dependent to some extent upon acsuibility and convenience, Inspections that are convenient and of a go-no-go nature are likely to be performed cin time aid with accuracy; those thut require considerable quantitative judgment may bc missed or interpreted incor-

reedly, For example, the presence of vital fluids in all gtarboxes, transmissions, and other reservoirs should be discernible from ground level without opening of complex cowlings. Min-max oil levels should be used to eliminate the nccd for topping off, and the minimum level should be exactly one or two quarts below the maximum whenever possible to discourage the practice of saving half a quart of oil in an open can. The minimum level should be such as to allow completion of several additional hours of operation at the maximum likely oil consumption rate so as to eliminate the need for adding oil when the level is near minimum, While accurate values of maintenance times for transmission and drive system components are not *

5

S

,v, ,au S•.,,t'.t ,

U.flI,,=FE,.

n,,ll) ,l.,,Wpf..,,

p-ua.aavu

data are helpful in identifying present troubleome areas. Table 4-1 presents maintenance workload factots relative to drive subsystems (Ref. 33). The TABLE 4-1. US ARMY HELICOPTERS TRANSMISSION AND DRIVE SYSTEM ONLY MAINTENANCE WORKLOAD (Ref. 23) -

j

*OR K LOAD RATING

PRO6LEMTITLE U A

VERY I II

HG

TAIL ROTOR DRIVE SHAFT INPUT DORIVESHAFT OHTAIL S ROTOR DRIVE ,SHAFT TAIL GEAROTO

ocoincident IU

I

I 4

,V

LO*4

LOW[

SY NCHRONIZ ING OR IVE

SHAF I

x

nial

with ADS-I l. This plan will include many elements peculiar to the transmissici and drive system. A good program plan requires the active participation of the

x

responsible transmission and drive system design ac-

x X

ROTOR BRAKE SLJPPORT

ASSEMBL V ROTOR DRAKE DISK

ROTOR BRAKE PUCKS

tivity to assure practicable approaches with miniriun penalties in &!ive system performance, weight, and cost.

OIL COOL ER ASSEMBLYV

-

with the reduction of vulnerability. Nor-

practice is to design a complete helicopter sur-

vivability-vulnerability program plan in accordance X

TRANSDUCERS CH 54

"INPUT

techniques applicable to this goal are discussed in

pars. 4-2.1, 4-2.4. and 44.3. Survivability in cases of combat hits is considered

OIL PRESSURE

LC MAIN GEARBOX CARBON SEALS ROTOR DRAKE SEAL AL;Y

4-1.1.4 Survlivabllly Survivability in transmission and drive syste. operation may be defined as the capability to susstain damage without forced landing or mission abort and to continue safe operation for a specified period of time, usually sufficient to rctunn to home base or, as a minimum, to friendly territory. The damage may occur from either internal component failure due to wear, fatigue, or use of a deficient or inferior component; or a hit by hostile forces. The current Army requirements generally define the period of time for safe operation after damage as a minimum of 30 min i at conditions within the maximum power and load envelope, except in the case of total loss of the lubrication subsystem; the acceptable maximum power level for safe operation upon loss of lubrication is generally reduced to that required for sustained flight at the maximum range speed at sea level standard condition. Survivability tollowig internal com.pOont failure can be enhanced through such detail design practices as identification of primary failure modes, and using configurations and arrangements to asure kindly failure modes and to limit f-ilure progression rates. Attention also must be given to the elimination or metardation of secondary failures causad by primary failure debris, and to providing for j.3sitive failure detsection long before a critical condition is reached. Safe operation with this type of damage normally can be achieved for durations of 30 to 100 hr. Design

X

I

INPUT DRIVE SHAFT Ail IG I,'LUT OUILL OIL SEAL ("1 4?I

rankings given reWate primarily to other maintenance factors on the specific helicopter listed, and are not to be interpreted as relating the workload on one bali-. copter model to that of another.

Reduction of helicopter detectability and the defeat of specified ballistic threats are important elements in vulnerability redt tion. With regard to de-

X X

X x

tectability, the primary area of concern in the cas of

the drive system is noise Wpar. 4-1.2.1.1.4). While the mower irquency noise levelsare basically associatd1

t

', ,'. -- t

*

A -

K ¾

'with the rotor and/or tail rotor and propeller. the highiev frequencies are generally attributable to the transmission and drive, and propulsion systems and their accessories, Typical Army requirtments specify a maximum scund pressure level for helicopter hover and fly-by at &specific distance from the flight path. Desge Soals for appropriate frequtrncies and sound prossuore@ are given in Table 4-2. Noise level survey iriquirements are described in Chapter 8, A MCP 706203. Design techniques to secure external as well as internal gearbox noise reduction are discussed in par. 4-1.2.1.1.4. defeat of ballistic threats must be accompliahed for smaller caliber ordnance and damage minimized as much as possible for the larger calibers. Depending u~pon requirements peculiar to the mission, the drive system components must he cap&6k of withstanding a single ball or armor-piercing .,762-mm ballot at 2550 fps, aligned or fully tumbled, striking at any obliquity at any point in the system. *,TM 75-dog solid angk of the upper hemisphere (with -Complete

larger ordnanca for %.hich damage minimization

the collector par in the main gearboc or combin~ing gearbox t-sually arc excluded from the survivability requirement by nature of their functional duplication, provided that: I.- No single projectile can kill all duplicated power paths 2. A single power-path kill cannot cause secon-t dary failure of the duplicated power paths due to firagmentation of the first. These two criteria can be tatisfied by: 1. Physical separation of the drive paths sufficient to reduce the impingement angle within which a single proj~iccr can produce a multiple kill 2. Sufficic. !size and strength of the killed-path component to attenuaste the projeenile velocity below the kill thre~hold for the second path component 3. Use of structure between the paths to confine a fragmented or loose component to its immcdiate locale 4. Use of armor to confine fragments or prevent projectile impact. eguaeu

should be considered is 23-mm high explobive: inisbMaCh otimotn Ilt n f cen~iry(HI) sighet iechnique for reducing vulnerabilty. Manly slihtconfiguratior, changes can increese surviva. he esgne to tchnqueo he avilbleto Thaspeifi Thespeifi avilale tehnque esinerto bility greatly with jut serious compromise of efw~ee the stated requirements include: fcecwiho ot I. RdundncyFor example, case hardened gears with tough, 2. Desigi configuration fracture-resistant core structure have surjplisingly 3. Self-sealing oil sump materials 4. ubrcatin Eergncy cnsidratonsgood toleranct to ballistic damage. Spiral bevtl gears 4. Armor.ec lurctocnieain and planetary gears, used effectively throughout the drive train of rimall and medium helicopters Ure in-

Y .

vulnerable to the 7.62-mm threat. Planetary ring

Aa 0

Redundancy is typified by multiple engine configurations. In thesm configurations all individual drive subsystemn components between the engines and TABLE 4-2. EXTERNAL NOISE LEVEL #X rERNAL NOISE LEVEL

IFREOJENCY. Hz BANDi CUJTEW

44.7-89.2

0.-78 M&f 709 1.410

OVEALLsosive 63 12

oc

1.4110-2.9m

(1000 .. 2000

2AZI5.833

4000)

L~~i 5,2111,222

oo

'NWA

SOUNDOPRESSURE LEVEL, CB PER CEI VE0

85 5 88 86portioned 851

76 72

CETE FREQENCY (DETERMINED EMPIRICALLY)

'ONvIFSSEM psWP

SNCILS

gears may be penetrated so that the planet idler gearsX. cannot mesh at a particular segment, but the remaining Sears pick up the overload necessary to continuc normal power transmission. The melatively high contact re jos and coarser pitch of spiral bevel gears t a facto~s that make them particularly resistant to failure from loss of a single tooth segment. Narrowface spu.- gears (less than 0.5 in.) c;an be a problem, and, therefore, it is desiroblc to use greater face widths in all primary po-wer paths. Experience. with 12 7-ns. 'n ammunition is less exten-

than with the 7.62-mini projetiles, but the same

general observations hold true with a slightly larger scale ol' reference. Glear rimns, webs, and integral

outer-race sections of planetary idlers should be pro-

such that a ricochet entering the mesh will deform, fracture, or crack the gcar teeth rather than A the tooth supporting structure. Extensive observation of main rotor gooarboxes damaged by 7.62-mm /

ball -and arurnr-pimrinS (AP) anmulinition

tv

o

AMMP 7W~202 shown the digsctive capabilities of conventional hel;copter gec~ring to be quite adequate to discharge the "spm bullet insto the oil sump without functional failure or the power transmission systrmn. Integral gear shafts quill shafts, and external intercon~nect or tail rotor drive shafts must be of au1Tficient diameter to withstand edge hits from fully tumbled bullets without failure. In thin-wall alumiinum shafts operating with a minimum of 20% margin on first whirling critical speed, an extc.-nal diameter of 3.0 in. is sufficient to defeat the 7.62-mm threat whil- a 4.0 in. diarantecr is necessary to defeat the 12.7-mm threat. Steel shafti-i may be considerably smaller depending upon the wall thicknms employed. Of course, in event of damage the remainin& portion of th~e shaft must have sufficient strength to transmit the required maximum torquc.If the column buckling torque is 300% or more above this torque, &simple shear-stress calculation of the remaining post-impuct area is sufficient. However, when the buck inS margin islevas, it isgenerally necessary to conduct real or simulated ballistic tests to demonstrate the adequacy of the design. Note that it

is undesirable to increase diameters excessively sincc vulnerability is then increased for fuzed round threats. Most ball and roller bearings are fabricated from through-hardened steels and. consoqoently, usually will fracture through the outer ring when struck by a bullet at nt.ar-ero obliquity. The outnouncling case and liner s~tucture serve to expend sorie of tha kirsetic energy of the bulict; however. cony ontional thickntsse of these structures gsterally are insufficient to prevent fracture of the bearing ring. Orientation of the rolling elements of the bearing at the instant of impact has much to do with the ring fracture mode. A zero obliquity hit between rolling elements fraquently will discharge a double-fractured "pit-slice" ring segment into the bearing, while an nligned hit often produms a single outer rikig fracturc and frequently fractures the roll~ni element so well. When the des.Sn allows, a space between the outer gearbox wall and portion of the housing supportiag the bearing liner ind ring is effective in reducing bearinb damage. P~as space provides a place for the spa?'ling debris from the initial impact to txpand and eject,

DUPLEX BALL BEARINGS

ROLLER BEARINGS

Figur 4-14. Typical Tall Rotor Gearbox

-

Vulnerable 4-19

*uRAP 706-202__

_____

thereby reducing the impact shock on the bearing Ving, Some insight into useful design techniques may be gained front examining three configurations of a sitnpie, spiral bevel gear. 90-deg tail rotor gearbox as used on most helicopters with single main rotors. Fig. 4.14 is a schematic representation of a typical minimum-weight design featuring overhung pinion and pear mountings, designated as Configuration I The pinion and gear are both in overhung mountings supported by duplex bail and cylindrical rollcr bearings. A single 7.62-mm hit on any ont of the four bearings probably would not result in instant functional failure; the boaring would continue to orcrate for some time because the considerable driving torque would bre~k up and eject the relatively frangible rjis.~ng elements and cage of the damaged bearing, Howtver, direct hits on either the pinion cylindrical roller bearing or the duplex ball bearing supporting the geau would soon result in excessive loss Of geAr

mesh position. As the ci fectivc: radial clearance of either of these bcatings increased with the resulting rapid bearing deterioration, the operating backlash of' the gear tecth similarly would increasc while the depth of tooth engagemcitt would decrease correspondingly. The probability of the gear teeth skipping mesht or breaking off upon application of sip nificant yaw control would be gkcat. Hits on thc outboard bearings would yield far lower probability of gear miesh failatrc. Configuration 2 is shown schematically in Fig. 415. The beval gear set it; identical to that of Conifiguration 1. Pairs cof the same types of bearings used in Configuration I are now used to straddle mount both pinion and geAr membtrs. In this configuration the increasing radial clearanctc in any bearing sustaining a hit wifl result in less dectrioration of the gear mesh. with a corresponding decerase in tlte probability of gear fetilurc upon sudden yaw control input. A gear-

box of this configuration designed for tail rotor

CONFIGURATION 2

~ROLLLO BEARINC-

7'

DUPLEX BAL.L BEARINGS

Flgvrt 4-15. Tail lRotor Gearbox 4-20

At

-

7.62 mm Proof

r7

AMCu±706202 steady hover power of morm than 150 hp wo-ald bc jiudged capable of the required 30 min operation subsequent to a 7.62-mm bullet impact. However, the probability or functional failure after receiving a 12.7-mm hU would be quite high unless the beating and Sear components were inordirsati'iy large. Configurption 3 is shown schematically in Fig. 416. This configuration has bee arranged to defeat 12.7-mm threats with far less weight penalty than would be incuticd by oversixing the elements of Configuration 2. The overhung mouanting of Configuration I and the straddle mounting of Configuration 2 have been couinbined in this redundant or composite system. Both pinion and gear members are supported by two conventional cylindrical roller bearings and one duplex ball bearing pair. Emergency thrust shoulders art. incorporated on the shafts adjacent to the integfal roller bearing inner raceways. Sufficient axial cekarance should be provided betweeni thec roller elements and the inner racv thrust shoulders or flanges to preclude contact under normal operation conditions (including extreme cold

I

when the light alloy housings have contracted roeitivt to the steel shafL4. H*A#ever, upon functional failure or either duplex ball bearjug. emergency axial locaition is provided by theae thrust flanges. Wse of a thiec-bearinS systcam p, .-wits total functional losit of any one bearing without seriously compromising the operating parameters of the gear meah, however, bearing alignment becomes more critical. As a result, one bearing of thc three-boaring system must be designed with greater ikiterna! -clearance than normal. The spiral bevel gear set shown in Fig. 4-16 has been enlarged slightly relativen *o the geai set shown in thea prior two configurations to derreace vulnerability to 12.7-mrn hits directly ini the gear elements. While numerous other configurations and types of bearings can be used to accomplish the same objectives, the logic used to providte inherent survivability itrnains unchanged. Simailar principitct should goivern the design of the entire dr-'ve subsystem. Their npplication, of coursc, beco~mes more involved as tae complexity of the gtearbox design increases. All shaft couplings, joints, hanger bearings or

3

ow'CONFIGURATION

D'DU:LEX BALL BEARINGS

W IT)

-12.7

*

7

-

Figure A-1,6. Tall Rotor Gear'sox

.

mrm Proof 4-21

13illow block hoirsiwigs, Warotor and intermcdiatc

4-112.1.44 FAmwgsuy Lubricadorn

lyjain gearbxec, and input/output quaill asiemblies tmust be j ined, retained, or mounted wit~h a sufificiamt numaber of redundant fasteners to preclude

1, ic preferred method of reducingl vulnerability is to assure fail-afe or emergency lmbrication in the event of total lows of thz normal luboicant sup*l This rANNhty must all'iw continued. Wae operatior for 30 min at minimum cruise power at mnission 9TSS

loss of function fiomt a single proje1cil. For well

.4prldattachment points, four fusicners ofttit suf-

`-'ic.oveve. otatins shaft joints ard ocaplings ýoften require six or morm fasteners. Frequent UKC Of -flanges, ribs and abrupt scetion changei; in castings, housings. it 1 simiiar structures providc effcti-vc 'stoppage of %.nckprapagatioia, while enhancing heat rejection to the. atmosphere. Internal ribs in Oil SUMP areas are desirable because or the ponsibility of cracking by hydraulic ram effect in the oil as well as by projectile impac. 4-1.2.11A.3 Sgif-" alig Su*4w Another design tecltirtiqw involves the use of selfsealing materiash in the gearbox oil sumlp area. The Inost efficitrnt material now available is defined as Type It ina MIL-T-5579, Thi* rubberized self-scaling compound originally was developed for fuel cells and can be fabricated to defeat either the: 7.t~2-rnm or the 12.7-mm threat. Another excellent defense material for 7.62-mmn threat is a cast urethane coating approximately 3/8 in. thick. With the latter material, the decin of the sump should be relutively silmple, as in a casting cope where the drag may 1-t witharawn without usc of breakaway core prints The cast coating contracts after pouring aad high residual comprersive stresis rcsults. This prestressed resilient coat then shrinks to close completely the hole lei's by the piercing bullet. The coating is relatively dense (2.0 lb/ft2 for the 7.62-mm throat), and also serves as an excellent beat insulator and noisc and vibration damper. Its density is such that the rurface. area to be coated should be kept to a minium to reduce the sittendant penalty on -sizing of the oil cooling system. Flat shallow oil pans and st-mps often provide the most efricieni configurations. The verious shaft seals must be designed so that a direct bit cannot cause all the oil to leak from a gearbox. This may be accomplished by using acnribbing labyrinth or slinger seals in #eries with the conventioaial contacting face- or lip-type of seat and by limiting the o-; flow rate at the inboard seal fecc to a minimum,.i Where external olcoolers and lines are used, current specifications often require the use of emergency o, :%ut-off valves to divert the oil dircctAly back to the transmission lubrication distributien system in thecevent of a cooler or line hit, thus preventing total loss of oil. One s3uch devicc is defined it. Ref. 34. 4-22

weight. Oil dams. wicks, and other itemns o e tain'n a minansum oil supply in the critical bearing areas are simple techniques to employ. Bell sind roller bearing cages may be fabricated ftom sacrificially wearing. self-lubricating composite ma~terial~s (Ref. 35) such as polyimides, Tefloni-filled Fiberglas matrices, and silver-plated, high-temnperature steel. The US Army Ballistic Research Laboratories (BRL) has clemonstrateo composit etilc gears that wear off on meshing drive gears, thus providing aform or geak. tooth lubrication (Ref. 36). Ref. 37 reports a successful applicaetion of a grease developed specilically for helicopter gear and bearing lubrication. However, the nnomal lubricants (M I .L7808 or MIL-L-23699) serve the cquc!!y important fuinctions of reducing fricticn and cooling. Ref. 38 desribes the reauiremeni that ihcrtnai cmuiiibriurn for the nrw "dtry" running condition be established to achieve 30-mmn safe operation aftet loss of the cooling oil. The equ.ilibrium can be established. only by main-K tami-ij adequate ruaning clearances and backlash in the bearings aa4d gears in the presence oC the therm. l gradients that exist in 2he "dry" condit.,on, with its alteredl frictional hWit sourcei and modified conduction, radiation, and convective hetat rejection vaths. The emergency "friction reducing" lut-ricants can be of value in austa..uing safe operation only in such a case. If a gear or bearing loses running clearance., a rapidly degenerative sequence of events results, in catastrophic failure. Loss of' operating clearance results in abnormally high he-at generation because the gear tceth and bearings operate under interferenc conditions with attendant overloads. Ti heat generation in turn produces an increase in the therinal gradients, resulting in a f'irther increase in overload and interferencc until the bearings sei;Le Or the gear teeth #, so hot the*.1 they undergo plastic failure. Specific metl..od& of preventing such occurrence arc discussed in par. 4-4.4-3. The general recomrmended design proceduic (Ref. 38? is as follows: 1. D'ouign for minimal frictional lesses commensuratc wita available manufactuiing ability. 2. Ca' culatc (frictional !osses for the "dry running" regime. An average friction coefficient 1 0,16 is suggested for the first approximation. Use ihis value with the mean values for sliding velocity, and load in the Psta meshes and bearing contact areas.

~

5

A

. ''

AL"~ 706=2O

*

3. Construct a thermal map with probable steadystate "dry running" tunparature gradients. 4. Redesign &I:gear Otbo bimfing elements to provide some clearance under the mopped gradients, Acdded clearance should be provided at high-rate frictional heating sources to accommodate transient conditions. For example, relieving clearance will not be provided by expansion of the gear cane until the increased heat generated by dry operation has beated ONe can. 5. Use materials with adequate hot hardness and frictiou properties for thermally vulnerable cornponents. 6. Provide self-lubrication cf bearings where possibin. Methods include the use of suitable cage materials or the use of appropriately located wick devimes 7. Re-caiculste bearing lives and Sear strc~aes for the increased clearance conditions occurring during operation in the normal lubrication regimc. AdJust design parameters accordingly; i.e., increase face widfts o:-pitch of glear memibei s, along with bearing capacities, as required. 44.2.143is

single main rotor configuration and are powered by either one or two 1~n..A shaft-driven, single-

In soecasesitmyb approl~riate. to employ armolt oprotect the vunml opnn.Ti design technique is the least.preferted bemause it adlds weight. increases maintenance Ptsk times, and penalizes the full-time payload. In such applications, integratl armor is prcferred over parasitic or bolt-on armor. Not only is the weight penalty slightly less with integral armor, but the pitfalls of increasing payload at the expetise of armor rrpnvai will he eliffinated. For most anplications dual-hasrdhess steel armor will be the moat efficient type to integrato because it can be rolled, welded, bolt-fastened, or integrally cast. Design of armnor installations is discussed iki detail in Chapter 14.

rotor driveshaft. Table 4-3 identifieb certain coniuaincharacteriktics for the single main rotor helicop~ters io cumr-nt Army use. iiiIt should bte viimc that accessary creasc with the sir.ý of thto helicopler. Lighti oli*4Lvation helicopters (LOH's) havo few accessory requ~reinents and 1,asibly no drive ,-dundancy. Ingr. ;nra, these helicopters may be flown safely without hydraulic booet of the flight controls, and the battery suffices for emeigency clectrical supply in the evmit of failure of th.- engine-driven gencrat.)r. Utility helicopters (UH) frequently require redundant hydraulic ptimp and electrical generator drives due to the magnitude of the rotor control loads and the increased clectrical loads attendant upon the larger amounts of inatrumeititation, electronics, and mission-essntial equil'ment. Cargo helicopters (f'H) often must havt auxiliary power unit (APU) for ground operation and checkout of electric-al and hydraulic subsystems. It is cornmon practicc to employ an indepnder~t aomssofy gearbox driven througN over-running clutches fromt both APU arnd main rotor gearbox to permit acprto rm te oe ore csr

4-1.2.2 Dii'. System Ccwsulgurstllm The specific requirementa for the drive system are dictated by the dctailasd configuration layout and vehicle requirements. AMCP 1,06-201 sets forth the evolution of an zpproved preliminaty design con. 11Suration, which will include detailed requirements for transmission subsystem power input and output drives; i.e., the speeds, powers, location, and relative orientation of these driives. Typical configuration requkirements for existing Army helico'flers are discusavd further in the paragraphs that follow. 4-.L MiS~ri Ma6l Rotor Drive tGyseaii The ma~jority of helicopiprs in current use are of the __

__

lifting-roto'r heicopter always employs antitorque reaction and thrust device to counteract main rotor driving torque and to provide yaw control for helicopter maneuverabilityr. A shaft-driven tail rotor tocatod at the aft end of the talboomn and arranged as eithtr a pusher or a tractor propeller is used most com.inonly. The tail rotor shaft is driven throu3h a 90-deg bevel gear set that ir. turn is driven from the main rotor gearbox by a long drivtsliaft or series of coninected driveshaftc. Thea power takeoff froir thc main rotor gearbox for the tail rotor is gear'ed to the main rotor drive downstream of the output sido of the freewheeling or overrunning. clutch iocated between the engine(s) and the main rotor gearbox; this arrangement permits full yaw maneuver capability during auto-.rotation or engine-out operation. Accessory drive requirements vary extensively and are dependant upon the primary vehicle mistion and helicopter size. These drives muy be miounted on the main gearbox or isolated in an accessory gearbox that driven by r shaft ftom the maiL, gaorbox. Secon-

.'qvircrrmnts

4-12.2.2 M1ihftlHU-raost Dulee Systems Multilifting-rotor helicopters have beer designeJ and test" in many configurationas - suchv,4 fere and -4-23

ý

TABLE 4-3. HIELICOPTER DRIVE SUBSYSTEMS

SINGLE MAIN ROTOR

-

TAIL ROTOR AND ACC'Y DRIVE DATA

MAIN ROTOR GEARBOX REDUCTION STAGES S SPEE11 rpm

____M.H. ENIN OTPT PU POWER., SPR hp

SPIRAL BEVEL.

T.R. SP6EO

SPEED,

PLANElARY

6,000 6,180 6,600

312 312 1,400

ACC'Y

2 1 1

NONE 1 2

AH-1 G

2,4 6.0130' 4,301

456 354 324

'R.G .6. RATIO

11

-

NONE NONE NONE

IN1 MED. G..

rm'DRIVES

SINGLE ENGINE 01+OHOH-58 UH-1I-

1

NONE NONE 1:1

0.67:14 2.35:1 2.6:1

1:1 1:1

2.59:1 2.44:1

_____

TW!N ENGINE UH-1N CH-3

6,600 18,966

CH+63

13,600

CHZ4 L

~

1,800

j2,500 j7.560

9,000j 7,S00

A NON4E NONE

1 1

2 1

324 203

2

2

185

2

2

185

4,302 %,030

j3

J

010 3,020

3

,j1.31:1 1.22:1

2.91:1 2.91:1

NOT ES: NE'~JRATOR. GE~ovs SONE ACE 550H Y PAL) LIN MAIN l..iARJibA, 4,20u rpm. FOR AND HYDRAULIC GENERATOR TACHOMETER rpm, 4,200 GEARBOX. ONE ACCESSORY PAD ON MAIN PUMP jN SERIES. 40OR 5 PADS ON MAIN a3EARBOX; 2 OR 3..,200 rpm FOR TACHOMETER GENERATOR AND 1 OR 2 HYDRAULIC PUMPS; 2.6.300 rpm OR 1 EACH 6,600 AND Z,000 rpm, DC GENERATOR, ALTERNATOR, COOLING FAN' DEPENDING ON CONF IGURATION. 2 AC GENERATOR. 8,000 ipm; 3 HYDRAULIC PUMPS; 3,700 rpm, 2 LUBE PUWPS. 2,50O AND 3.7G0 rpm;4. AND TACHOMIETER GENERATOR, 3,900 rpm. ACCESSORY G.B. POWER TAKEOFF, 6,020 rpm; SERVO Pump. 4,600 rpm. TACHUMETER GENERATOR. 4,200 rpyn. Ah2 ACCESSORY GENERATORS, 8,000rpm; 4 4YDRAULIC PUMPS, 2 4,300. 1 EACH 3,700 AND) 3,200 rpm; AUXI L IAR Y SE R 10 PULMP., 3,70U rpm. COMBINING GEARBOX APPROXIMATELY 5:1; 1 SPUR AND 2 HELICAL S'AGES 1 S'nUR AND 1 HELICAL STA.GE.

±ENGINE

aft dispowAu, laterally disposed, coaxial, and quadrilateral main rotor arrangements. All of these layouts

feature counter-rotation of even numbers of main

*

rotors to cancel the torque reactions and hence eliminate the requirement for nonlifting antitoi que device. All multirotor helicopters require rotor syr.chronization, which usually is accomplished by interconnect shafting between the individual main rotor &eaboxes, or by dual-oitput reversing reduction I;LtrinS in the cms of the coaxial confiaguration. In instance whert separate ewgines are located et each miain rotor 6tarbox, the crossshafting supplies power to each rotor for engine-out operation. In an~y instance, the intecmonnect drive is essntial to safety 4-24

of flight, requiring reliability comparable to that of the nizin rotor mast and thrust bearing. The interconnect drivz is located downstream from thc engne free-wheling clutches. The only multiliffin,' -rotor helicopter in current use by the US Army is the tandem-rotor CH-47. This helicopter fcatures twin engines of 2650 hp at 15.160 rpm. The engines are located in outboard nacelles high on either side of the aft third of the fuselage. They drive directly into 94)-deg reduction r,0_'TboXes that drive into a combining gearbox alko with 90-dg shaft angle spiral bevel geaA. T\,%oombining box is an integral part of the interconnect syr..hronizing drivc to the forward and aft rotor gearboxes.11ese

..

*,,

.

.

t

xt

.

-, ,7 -0

AMCP rotor gearboxes each feature a sih0lc spiral bevel and two planetary reduction stages with final output At 230 rpm. Thaccessori¶ arc all located at the aft main rotor gearbox and consist of oil cooler bir.wcr, two ckctrical gciucrators, and two hydraulki pumps. 4-11.23 Compomi HICellOpl Drive Systems Compound helicopters arc those that use cuxiliary propulsion devices other than the main liftng rotors in the forward flight mode. The majority of such designs have featured a single main lifting rotor, an antitorque rotor, and tither turbojet engines or shaftdriven propellers for auxiliary propulsaon. The only compound helicopter to undergo development test or Army use has been the AH-56. It was powered by a single 3450-hp engine driving directly into the main rotor gearbox. A spiral bevel gear stage, a compound planetary, and a simple planetary provide the reduction gearing for the main rotor. A spur takeoff located at the engine input to the main rotor gearbox drove a shaft running along the top of the tailboon,. This shaft drove the pusher propeller at the ead of the tailboom directly; and through a 90-deg shaft angle spiral bevel gear set also drove the antitorque rotor. Accessories were mounted at the main rotor gearbox and consise.d of two hydraulic pumps and a high-speed generator.

"4-1.3 TRANSMISSION DESIGN AND RATING

20

simply to achieve longer life of drive system corponents. Sufficient cycles will be accumulated at the 5-mirn rating during the service life of the drive subsystem to require the same bendiutp fatigue gear design, i.e., infinite life, as would be required for a continuous ratig at the same red-line limit. Although a 5-min drive system rating does not usually impair the operatianal capability of a helicopter with a typical speed-power relationship (Fig. 4-17), current Army specifications include a continuous drive sytem rating. A typical requirement would be a continuous rating of the main transmnission equal either to 120% of the power required to hover out-ofground-effect (HOGE), zero wind, at the density altitude defined by 4000-ft pressure altitude and 950F, or to 100% of the intermediate power rating of the engine(s) at sea level and 95°F. The effects of power ratings upon life, overhaul, and selection of standards are discussed in the paragraphs that follow.

I

-

L

i'

4-13.1 Power/Wfe tuteatleo, The mechanical failares of interest to the drive subsystem designer usually exhibit a definite relationship between life and power. The life-limiting failure mode of primary concern for a developed and serviceable gearbox is pitting or spa,:,iig of the gears and bearinbs (par. 4-2.1). However, the life/power relationship for this mode of failure is not reckoned with easily due to the many tacters that govern the

CHARACTERISTICS All elements, components, and subassemblies of

the transmission and drive system are subject to varying degrees of wear, abuse, fatigue, and other environmental hazards. In many instances, ,tandard components will provide acceptable performance for a given drive system design at a savings in cost, ease specially designed components. However, the designer must have a thorough understanding of the likely failure modes o; standard components (pars. 42.1 and 4-2.2) and the pertinent life-load ow lifeenviromrment reldtionships. It is customary to specify an input torque limit for

['I

1_

-

35')-

-

-

-

_00 C0- T-

I..__

. ." I5D-

, )

-

-

engine(s).

,

k

2

•5 0'•-"

a helicopter main rotor gearbox. Indicated to the pilot by a torquemctcr red-line, this limit may be lower than the sea-level-standard engie~s.._ rating of the Depending upon such factors as helicopter type and design mission, the red-line torque usually is specified as a continuous rating, or, in rare instances, a 5-min rating. A 5-min limit may be specified for •energency operation only. A time limit is imposed because sufficient cooling capacity is not available for extended operation, lubricants may be degraded, or

.41!

-

DESIGNj GROSS T

-MAX.ALT.GR

..

0o

W

.

6

90 120 AIRSPft.. ki

ISO

ISo

210

Figure 4-17. Typical Speed-Power Function 4-25

rdationship. The metal chemistries, heat treatnments, lubricants, loads, specific iliding ratios, velocities,

The Hertz-stress/life relationship varies significantly (Fig. 4-18). Each function shown results from

temperatures, geometric shape,, surface textures and roughnesses gearbox deflections, and lubricant chemistry (including water content, and other contaminants) all influence the life of the surfaces in contact, or more properly, in conjunction. It is not unusual to observe dramatic life differences between two supposedly identical gearboxes when but one of the given variables is changed by manufacturing scatt•r, operating variability, system wear, or en,ironmental factors. In a complex system of gears and mixed bearing types, it is generally acceptable to use Miner's rule of cumulative damage in a simplified form for life prediction. A representative root-mean-cube power ievl is calcu!ated from the assigned mission profiles using Eq. 4-15. The value of compressive, or Hertz, stress S,. corresponding to the RMC power or load is then calculated, and the life determined from an applicable S-N curve,

data representing a particular set of design and operating variables. The wide variance among these functions emphasizes the danger in the use of a Hertz-stress/life function without consideration of the assumptions and test conditions. Because calculated Hertz stress is an exponential function of load, little generality is lost by Rssuming exactness for the RMC life-load relationship and selecting an appropriate classic or modified stress-life function to predict the life of any particular con junction whose variablc3 are most nearly represented by the chosen function. As an example of the selection and application of an RMC power, consider the three-mission profile spectra shown in Fig. 4-19. The UH- IH and AH-IG power histograms were constructed from flight recorder data (Ref. 43). The third histogram was constructed using the fatigue spectrum supplied with a recent Army RFP for a helicopter with a mission role

350

I -AGMA 41i.02 GRADE 2 SPUR AND HEL;CAL GEARS II - GROUND AND CARBURIZED AMS 6265 SPUR GEARS (REF. 39) III - SAME AS II BUT HONED FINISH (REF. 39) IV -ASME DISCS 30% SLIP, CYM 52100, POLISHED (REF. 40) V - GROUND AND PEENED CARBURIZED AMS 6265 SPIRAL BEVEL GEARS (REF. 41) VI - BACHA ROLLERS, LINE CONTACT (REF. 42) 1 --

.__

--.

IV

III

-

V

\

I

250 21N 1V1

""16.

I.

I%

I

U.J CLi

C-2

50L

10~

103

10'

108

lip

100

LIFE, CYCLES( Figure 4-18. Gear Stres vs Life Curves 4-26

0'"1'

AMCP 706-202 ratio. Thus the stresses under the power fow the UHIH and AH-IG, respectively, will be:

30[

(S')iV

"RED LIft,

S~and

"4 .. 2

4

E]i...

. 6

A

1o2

(ScI4H.IG

I SRF

12

1

j.

SPECTRUM RED• LINE

]operating

S~~diction

X Itr and 6.5 X 10' cycles for the Eqs. 4-17 and 4-18, respectively. 11 indicates values of 1.5 X tO' and respectively. Thus, the life predicted the AH-IG (Eq. 4-18) by Curve I is 7.22 times the for Also,4-17) withisCurve I thethat life life predicted 3.0 times predicted for by the Curve UH-IHI1.(Eq.

_

.T

for the AH-IG, while from Curve 11 the ratio is only 1.67. Clearly, it is essential that a stress-life (S-N) curve used represent accurately the design and conditions if a reasonably accurate life preis to be achieved....

,The apparent large increase in life at equal values ¢omparison with cn t r th. snprat k•_rl eapr

S17 •-, SHAFT

=

proximately 2.0 stress levels of However, Curve 9.0 X 10' cycles,

REDLINE

16

161.200 psi (4-17)

(0.75)½ (200,000) - 173,200 psi (4-18) Curve I of Fig. 4-18 indicates predicted lives of ap-

14

AH-IG

4a

-

n

10

0-

(0.65)1(200.000)

-"

SHP, hp

10-

Fi;1par 4.19. S6ft Horsepower Spectra Histograms

similar to the AH-lG but powered with twin advanced technology ergines. The red-line and flight profile powers corresponding to this fatigue spectrum are taken from Fig. 4-17. The RMC powers for the three spectra are: UH-IH, 714 hp; AH-IG, 827 thp; and RFP, 1939 hp; representing 65%, 75%, and 69%, respectively, of the red-line powers for the three helicopters. However, because the sea level standard inermedip.te power ratings of the engines for the three hel.,,opters are 1400 hp, 1400 hp, and 3000 hp, respectively, the RMC powers represent 51%, 59%, and 65%, respectively, of installed engine power. On the assumption of no changes in lubrication state with advanci..g wear, the stress-life functions of Fig. 4-18 predict differences in the expected service lives of the same transmission system used in UH- I H and AHIG helicopters based upon their respective AMC powers. For purposes of comparison, assume that the red-line power corresponds to a maximum stress S, - 200,000 psi in a particular gear mesh. Because the Hertz strest in a spur gear is proportional tc the square root of the load, which in constant speed operation is proportional to the transmitted power, the stress under RMC and red-line power w' be related by the square root of the power

the straight spur gear (Curve V vs Curve lii) can be

explained best by the difference in the assumptions used in the calculation of the contact stresses. The spur gear analysis is based upon a cylindrical contact assumption wherein the ratio of peak to mean compressive stress is 4/i"or 1.27324. The sprial bevel gear analysis is based upon an ellipticu, contact assumption wherein the ratio of peak to mean cornpressive stress is 1.5. Although neither assumption is really valid, the ratio of the peak stress for the bevel gear to that for the spur is 1.178 for equal bearing or conitui arias. Ti11S izin accounts ro .. h..l... the stress difference between the two curves at a selected life of 1.3 X l(Pcycles. Additional gain can be attributed to the shot peening process that was applied to the spiral bevel sets (Curve V) but not to the otherwise comparable spur gears (Curve III). Life Rating 4-1.3.2 Tfausmluo Ovetrk The various gearboxes, driveshaft assemblies, and bearing hangers that comprise the typical drive subsystem of Army helicopters in the past may have had widely differing times between overhaul (TBO). Main rotor gearbox TBO's ranged from 500 to 1200 hr, tail rotor gearbox and bearing hanger TBO's were as high as 1600 hr. and driveshaft assembly and accessory gearbox TBO's ranged from several hundred hours to unlimited intervals based upon conditional overhaul. Specifications for next-generation US Army helicopters call for much higher (3000-4500 hr) MTBR 4-27

'

¶

without dictating TBO valuts. However, using the relationships of par. 4-1.2.1.2, application of a 2000hr TOO requires attainment of a 6002-hc MTBF to achieve the. I 00-hr MTBR (par. 4-1.2.1.2). Although this MTBF concerns only failures of sufficient importance to cause gearbox removal, it canvot be attained easily. The ultimate design goal is conditional removal without scheduled TBO levels. Achievement of this goal requires the use of reliable and thorough diagnostic techniques (par. 4-2.4.2) and failure anodes with low rates of progression so aperatioai can continue at least short-term without compromise of safety of flight. The question of a cost-eflective overhaul thne, one that balances the increased cost of overhaul due to possible extensive secondary dwtage and cornosion against the loss of residual usefuk le, is beyond the scope of this document. A coMt analysis of TOO based upon direct and indirect operating cost -A the drive subsystems of light, medium, and heavy twin-engine helicopter is reported in Ref. 44.

tion) design standards and specifications for gear tooth bending, scoring risk, case hardening practices, and gear precision clasifications amx excellent design starting points. However, experience accumulated through development and field tests will suggest further sophistications and modifications. Many useful standards and specifications are published by the Society of Automotive Engineers (SAE). The smaller size bWaring locknuts and washers are useful, but for larger bearings the SAE parts generally are too heavy. The thread specification series also is ideal for use with special beating or spline locknuti because the series includes sizes cornpEtible with standard bearing bores. The 30-deg pressure angle involute spline and scrraticn standards will suffice for most spline applications and lend themselves well to inspection wvith simple gages. In special instances, where greater precision is icquired to improve load sharing among the teeth or to im.. prove positioning or location accuracy for the mating members, a standard spline can be modified by reducing the involute profile, lead, and spacin" tolerances.

4-1.3.3 Trasssdo Stanxrds and Ratings The use of available standards in detail design is encouraged for many reasons, not the least of ,vhich is cost reduction. Available stindard3 can contribute to lower costs if it becomies umn.•',c ry to prepare special specifications; conduct qualification tests; procure special tooling; and othlZrwise compound procurement, storage, and supply activities. The standards available include military (AN, MS. NAS, AND, Federal Specifications, MIL Standards. and AMS) and commercial (AGMA, AFBMA. SAE. AISI, and ANSI). However, the limitations and ratings of standards must be thoroughly understood to prevent their misapplication, These following are some instances in which it -Is better to select a nonstandard part: i. Excess cost or nonavailability (many published military and commercial standards never have been manufactured) 2. Insufficient strength or inadequate properties (published standards for parts such as studs may not provide the required static and fatigue strength or corrosion rewistanoe) 3. Inadequate quality control for the criticality of the application (many published standards include an inspection sampling frequency that is inadequatt for critical applications), Sonic recommended uses of commercial standards are dissuaed in the paragraphs that follow, AGMA (American Gear Manufacturers Associa-

cordance with the standard AFBMA (Anti-Friction Bearing Manufacturers Association) metric envelope dimensions, using the Aircraft Bearing Engineers Committtc (ABEC) and Roller bcearing" Engineers Committee (RBEC) precision grades. De-. partures from standard envelopes may be necessary for very light series, large bore bearings; but the cornmon bore size, width, and outside diamct-r increments and tolerancs should be retained to facilitate use of standard inspection equipment by the bearing manufacturer. Cylindrical roller diameters and lengths will vary among suppliers and may not follow recommended values. However, individual rollers with one of two crown kadius or drop values are usually available from all aerospace geade suppliers. All ball bearing suppliers furnish balis In 1/32in.-diameter increments and occasional'y in I/64-in. increments. Standard grade tolerances in microinches govern sphericity; e.g.. grade 5 implies 5 pin. sphericity. Many special steels, frequently called "tool-steels", using consumable electrode vacuum remelt technology, are finding increasing use in helicopter gcarbox applications. The chemistries of these steels are identified only by AISI (American Iron and Steel Institute) specifications. It is frequently necessary to add special limits on trace elements and inclusions to thas specifications to make them comparable to some of the commonly used AMS (Aerospace Material Specifications) grades.

VV Y S.11GVGF

4-28

poeibc.e

bearings sflmowi

cIin

Qj

,

ac-

'

-.. PcO 7"S>T7,' and XZPL and C. arc constant by definition, we havc the immediate solution: 01 (4-23) .hp/*F - PLC) (i CC Ts - r; - 5T + T)2. ' A S'S

T~)veloped

Substit~ating Cc into either of the initial hot or cold lots cquations (Eq. 4-21 o;- Eq. 4.22) will y~eld the

total power loss PI lIfthe necessary temperawrws ama measured for each tist condition. the individually calculated values of Cc may '3c averaged and a probable error computed by stan- ard statistical methods. 4-1.4.2.7 Theraid Mapping Tests Time and instrumentation capability permitting, final design modirIC46ions or the proportioning of lubrication distribution, alonj, with necessary adustzient of bearing parameter such as clearance and In-7 ternal. preload. may be accomplishe by thermal mapping. Thermocouples embedded in contact with becaring inner and outer rings and with gca blank nms or tooth fillets, for example, whoulid be used to construct a thermol mapi of the tanumimaaon. Messurement of rotating cousponrif temperatures Me quires the use oif slip rings or similar devia. Thw use of infrared photographs of opersting gearbose. also has been very effective in thermal mapping. Hot spots or excessive thermal gradients as' cause for *orrective design measures.

~

~

~

5V

~dg-

Overpower testing, sometimes referred to as weak point testing or modified stress probe. testing, is intended to yield rapid results to enable the designer to make timely charnges. Th~e purpose of this testing is to producc failures and definec failure modes and failsafe features, not to demonstrate rmliable extenided operation. However, a 100-hr failure-free overpower test at from 100 to 125% of maximum continuous power on two samples certainly would indicate that the gearbox was ready for life substantiation or qualification testing. The maximum recommended overpowei test icvei is1201-130% of normal red-line power, although in some instances 110% is used. For valid test results, the following conditions described in the paragraphs that follow should be satisfied: 1. Lubrication states should remain unchi igcd for the main power path ecu-ponents (Fig. 4-3). EHD film thickness as predicted by thc Dowson equation (Eq. 4-5) is relatively insensitive to load (125% power should reduce h values by about 4% from their 10D% power levels for an isothermal condition); however, because the temperature of the conjunction may incrrase as the 3/4 power of load, which in turn will reduce the viscosity of the typical MIL.L-7808 oil by 28%, and of the h value by 22%, a cautious evaluation is demanded. Excessive deflection must not occur. If debevel gear patterns degenerate excessively, their reduced area, cou.pled with the inci eased it%oth load, could result in doubling unit stresses ht the 4-31

-

r [

AMCP 706-202 overpower levels. The "'small-cutter" and oiher types of spiral bevel gears tend to resist pattern shift with increasing power and are good candidates for successful overpower testing. Iir well designed planetary gear reductions, it is not uncommon to find a 50% increase in uait stress for a 125% overpower test at constant spvcd. 3. The mechanical limitation of ball bearing load path constraints must not be exceeded. There should be sufficient race shoulder height and bearing mounting rigidity to retain the ball path fully at the overpower test condition. 4. Cylindrical roller bearings should have suf-ficient racetocrown) to preclude ent roller lercong crown (orra o prcreade s severe ever end lo an du eton Ssimple Hertzitan deflection. -vrpwe. The increased thermal gradients preent during Sovinrpower testing must not result in excessive bearing preloading or gear misalignment due to housing distortions. Design criteria for successful overpower testing must preclude gear toatli bending fatigue failure, case crushings, or scuffing (scoring) failure modes. Acfencatidtewear without ceromi eon of s t desin function of the gearbox for the specified test intervan is te criterion of success.

4-1.4.4 Other Life and RelabUilty Sibsbmatlatlon Testing A 200-hr qualification test is required by AMCP 706-203, and follows the tests in the preceding paragraphs. Also required are a 50-hr preflight assurance test (PFAT) and a 150-hr "must pass" qualification test in a ground test vehicle (GTV). Beyond these tests, it is frequently desirable to conduct extended bench or GTV tests to assist in the determination of initial TBO levels and to uncover failure modes not detected in previous tests. All testing in these categories is based upon spectrum loading conditions. The selected spectrum should have an RMC power level in excess of the anticipatcd flight spectrum. Because most lubrication system elements (including shaft seals) exhibit failure modes that are insensitive to power levcl, no meaningful accelerated test programs exist for the lubrication system, and its evaluation requires the accumulation of many test hours. Although the majority of lubricat;on system components will have undergone some degree of evaluation in early tests (par. 4-1.4.1), evaluation of their performance in the total system environment must await these extended time or endurance tests. 4-32

4-2 TRANSMISSIONS 4.2.1 FAILURE MODES Many competing failure modes exist simultancously in any mechanical transmission device. The modes rewognizod as dominant are often representative of the life-c' cIe phase in which the observation is made. Recognition, classification, and definition of safe operating limits are fundameptal to successful design. Failure modes may be identified as primary and secondary for ease of analysis. In one study based on component replacement at overhaul for the UH-i and CH-47 gearbox, secondary failures were shown to exceed(Ref. primary failures by at least an order of magnitude 46). Although the majority of design effort is directed toward preventing primary fai! area, the cost of drive subsystem maintenance and overhaul reflects the total of both categories. Therefore, reduction in secondary failure modes is an important objective for future design.

4-2.1.1

Prfry Faili

Modes

ae fmp odnan iderv iaiecausethosoet render a comomnnt unservicenaif because of some .self-genrated conditional occurrence other than normal wear. Cracked, broken, pitted, or spalled eluments that fail while operating at normal loads, speeds, and ,nvironmental conditions are representative of this failure category. There is a reasonable statistical level of occurrence for primary failures, perhaps on the order of 0.5%/ 1000 hr, that typifies the normal dispersion associateed with acceptable and cost-effective design practices. Failure rates in excess of this level aor considered a result of design or manufacturing deficiency. Identification and elimination of components with excessive failure rates arc the objectives of the qualification assurance testing outlined in AMCP 706-203. Properly designed and manufactured drive zystems must not exhibit catastrophic primary failure modes. It is not unreasonable to expect primary modes to be exclusively noncatastrophic. This criterion may be satisfied by inherent redundancy in load paths or load sharing, or by failure prcgression rates that arm commensurate with available built-in failure detection and die.goostic d,. vices. ,.onscientious application of classical structural analysis methods as modified by relevant test and service experience, coupled with adequate quality assurance methods, effectively will eliminate static and bending fatigue failures. However, the surface durability of loaded members such as gear teeth and

9"

,,A

706-202

roili-.4 element bearings is by no means thoroughly understood or easily preuicied. The interaction of the effocts of friction, lubrication, and wear (the modern discipline of Tribology) is the subjoct of intensive rcewm•h (Ref. 47). Drive dvsign is influenced by variables suct, as metals (hardness, microstructure, chemistry. cleanliness. residual stress), finish (routhness, lay, texture), surface treatments or coating, lubricants (base oil, viscosity, additive package), moi3turc and other contaminants, speed, slip, Hertzian stres., contact geometry, friction, and temperature. These variables, separately or in combination, may vary observed life at constant stress by a factor of 500 in conventional helicopter applications. Their combined effects also exhibit slope variations from -5 to - 12 of log.log SN curves. Because it is impossible to consider the

t

quantitative effects of all permutationa of the pertincnr paramnters deoribed in current literature, the significance of relevant test experience cannot be overemphasized. The classical stress-life equations or published S-N data must be viewed only as starting points. Table 4-4 presents useful qualitative influences of qome of the variables affecting S-N chata4,teristi.. There are many combination effects among these variables, but virtually none that result in contradiction of the indicated trends. The presence of relatively high slidc/roll ratioz and thin lubricant films is necessry for the surface pitting life to be sensitive to the additional factors shown in Tablc 4-4. Pitting or spalling generally is considered to be the result of metal fatigue due to cyclic contact stress. Under idealized conditions, the initiation of pitting occurs at a considerable distance below the

TABLE ". LIFE MODIFICATION FACTORS VARIAUt E

INCREASED LIFE

-

-_.', :

SURFACE DURABILITY

REDUCLED LIFE

OUALIFICATIONS

MLIALS

HARDNESS

RA60 -- 63 W-0

RETAINED AUSTLNITE

.

3. Improper attention to joint and fastener toq"ireinuceis. Sufficmet flange thickness must be proevidto dtribute loads uniformly among the proloaded tension fastenr (bolts or studs) susJ on case flange joints. Fastas pmload must fi sullicient to "maintain tension at -651F. preclude strs mvmsls during normal ocaillatory loading, and maintain Vflane contact under tension loading. For ,opeuly designed fangs with compatible fat.cr spacing, a conservative value for fstener tension loading P, for monact-loadod cylindrical joints is given by P

-

M

(4-,)

b

where N - moment. in.-lb Doc - bolt cirl diamneter, in. N - number of bolts or studs (equal ipacing assumed) The yied and ultimate streagths given in MILIGIDIl-5 for tandard AN studs repesnt th , r ab•m ,htead wih rn-r= iAIh.___. -. on have a ofte Sowemcharactenst, more daletios efect upon the instaled arengtsof eaidl. AN aeude than upon lgW sm. Eves with caxius practicable pu Vpendnlaaiy for tapped

boles the combined effects of aper snd squarenew of joints may indwa benteg koad such that 3/&-im. srom studs demoasrate ueauflc filure at 90% of hd"ook minwimum values. The significance of th is probbly bete demonstrated by the fact t"at typil 7/164is. dimeu stdsusumlly demonstrat double the iintu tenile strength of comparable "t

*

rlrm. 1

":

.

-&.-A-

a6 WIM

-:4.

WE&" mitaf.

-.

•,si.'

-4

,

weight pnatie, of 10, and 25%. rspectively. S fltid tapsled holes for s• istalatioes m'um be vented. Whe toerqu must be transferred throuh a flang joins a ouqu capacity Q(4in the abeewe of externlr Unas loads is given by

*whew N bloc ,o iJn d Daking 4-"

a toqpueL bia. - ma4k of smu -boftckk disisrnwii. - stud •"diameer, in. this valet doeis, key or othir mcheacal devses, dsMuld be se to prevm beain

Although it is comand shr di of the studs. mon to consider stud sh stren4th in determininug ulimate joint streagth, i Sctor should not be Pie upon for normal design torque asnalysus The jo*L oaiguulon dtortically neessary to mt the strength requirements having been detmined, it becomes imperative to asure cornpatible dfta desig of the machined surane for the ftenrs. Fractur resiuta•ce as well as fatigue condlatkion requise careful attemtion to semirgly les sigicant deA. Onerus- filet radii must be providMed in spotfaces. ountnborns. and keyways. All rath edge -arp curre mud be ciamf•ered, a"l b must be removed to minim= mes ctcnnratsoa. Smootb blending of inateusctions on critical machined surfaues is also w iesay to mimimize stress coacesaratio. Addit"ony the ow rbone or sotfae eim must be adequate to provide wre.b clearsa. for normal mauatmacace opavion. 4. Failure to covmmt dsllectim vd loading in coftgous structure. Many re d istancu s of gerbox mounting lug failure am attributiale to x.ternally induace So*& that wer Miored in the design anhiv. The cmst &wbox structura whi ", often ii a suTir than she a-raum summon to wic itandis mounted may provide a load path for bending fow torsin ractions p -sent in the airframe strsaure due to boding gar, roto thrust, or vibratory resposess. Examples are the aauchm•nt of a four lug amrmy earbox to a dec strocture that has vatscal beading odes, or the smilar mount* of sa iniuo0du r genbos on tail boom sructure that usr tail rotor dallos torsional ddeiodas with to obtain the safpith inputs. Beam it is doesk ty inhert with mountiag raedndancy, it is usually a m A *%r am mfa,!is sneak.e wte mom rr - r' rP -C---,-frme strucue locally or t' Wirotta aduistaol compliamnc at out or mor,. e•arbox attachment by oanof elaomeric members. Siemg up the awboa ose merely modmn the failure mode.

Oi

-uxs

4.U.13 Moalub ad

MageMum b fal inAt disfavor compared to aumaaim for cam and houuap becamus shorWting of srvtice life due to corrosion has hewn. a sowfia ant miteneam and spam replcmet aepes to the Army. Be•muse asaaeiim is a. tc with susect to all other mmals. failuet to a•ploy Oade design. prmws, and pmventive mainUance mmsume has cued an ovrunplhsis of the deicin of the mata. bare magsum acsc4 is afecmted ka by exposure to msrine axmaphe &hfu il umaprONLmd mild steel (Rtf. 101). However. corrmui of smeeiua allos can be

4

::•; -.

•

ii:•,-..

-r --- ''•.1•°

avoided succesefully only if the designer and fabricater follow the complete sequence of I. Design 2. Cleaning 3. Chromating or anodic film application 4. Surface saling or impregnation S. Painting 6. Asembly 7. Routine preventive maintenance. The nmot frequently occurring inadequacies in recent Army experience involve design and maintenance, Aluminum alloys should be used in areas of high susceptibility to corrosion. Ref. 46 reports the replacement rate for AZ91 magnesium main trmsma ion ca at UH-i ovrhaul as: Top case 16.0% Main cane 1.7% Support cas -2.3% Sumpcase .% Quills 1.0% In the cane of replacements of the top case, 1/3 were attributed to improper protection of bare surfaces during stipment after removal of the main rotor mat, and 2/3 to in-seice corrosion. The relative re#piafemfen rates smeat that the environment in nal the top cae operates spartzrularly conducive to corrosion. Aluminum alloys with high silicon conttet (6 to 12%) have been found to be supo to other alumiawn alloys and mngnesium alloys with respect to wear reastance. Properly dosigned spines of these materials will exhibit negliwear when operating with floating steel mating spliuwes. Magawsea and aluminum alloys commonly us-d in hldicopt bousin and ca- are listed in Tabe 4-6. ,.i..-re.ina ad .,tneai kat he.. i ,c.nat,..,t4 aeT-", can asure the higest allowable stregth .preaties. Tw tuseof MIL-A-21l10 control -flcatio rmaher than t of QQ A-I generally ill insur 25% hither adiowable fatigue stenth, alwlumahl• ciee 251)6 w tUow fatiguersT~h,patenthough the cogtmay be20 to 5M greater. it tin weight uavings may be n high as 40% if s". etrength delons tk design msl provided that mini. tri.ion ae nt imposed. The ,mum wall Mt..it Semide propartue of many aluminum foging alloys may be iinproved by cold working or mechanical artr m (Ruf. 102). Procurement and proces specifications for asngand are defined in MIL.C40l21.T caslins da~nd forgiu frg aSm dn in MI r eL- 1.iThe detail doesgn drawing t require t h f(lTowin Sprocesnmd a mi(NnD)u- ts Ic

)wbkb

-ibg

i. lmproqnsion. Tlermo• ttuig polyester mins \pg MIL-STD-276 wre omomended fo casting im-

prqgnation. Vacuum processing isesntial to remove gas bubbles from casting pores and to permit good resin permeation. Leak checks may be acrostatic or hydrostatic, although the former is preferred for sensitivity and cleanliness. 2. Radiographic inspection. Radiographic inspection is required in accordance with MIL-STD452. The detail drawing must call out x-ray views and should include a stress diagram to assist in determination of techniques and interpretation. to be employed. T(he x-ray technique should be able to resolve 2% of the thickness being examined. Film interpretation is based upon discontinuity gradations as defined in ASTM E-155. 3. Surface crack inspection. This must be , comaplished by fluorescent penetrant techniques as defined in MIL-l-6866 and MIL-i-25135. Inspection must be performed after forging, final heat treating or ajing. cold working, stress relieving, grinding, wdding, and maci~ning, but before polishing. tumbling, shot peening, plating. resin impregnation. or painting. 4. Hardness inspection. All castings and forging should he

chateSl for hardens by the standard SIn]

kg Brineil (or equivalent) method to ascertain that full heat treating and/or solution aging has been a%comnpled. This inspection must be performed prior to shot peening, plating, or painting. 4

1

Quls

External and internal quills frequently arc used to bouse a gear/bearing subassembly to facilitate moduhunmaintenance techniques and reduc the compkxiof the primary gearbox housing. Problems typ-cally encountered in helicopter applications are ociated with excessive wear or high tempeaature pcre of the housig bore that accepts the quill. The primary ctaie of wear is the ase of material combinatkms that permit differential thermal expansion with lomotsem becoming excesive at operating wtp ursCobnin*ha itmativte toupetures. Combinations that at elevated this problem while retaining mperatures allkeuv em of ambly at room temperatures. Hower, the tlmDml stress effects at -6erF also ma be conidmed (par. 4-2.3.1.1) to assur, that the material yield strKenth is not exceeded. Decause steel liners frequenly are used iP li•ht alhoy quills, the solution mom us be eatended to consider the efcts of four conceinric rit p with respect to their individual tolerans material Mrengts, fit-up, and thermal expensmo coefficients. Sealnt compnounds always hiould he is at extenal quifl-housing joints to prevent water entrapmen.

"4-67

¶

TANLE 4=4 rICOPT

ThAN&MISSON 11 CAS•E MATERIAlS AND APPLICA11ON DATA

MATERIAL DESIGNATION

PFOPERTIES.APPLICATION. AND RESTRICTIONS MOST FREQUENTLY USED AL CASTING. EXCELLENT CASTABILIlY. BESTCORROSION RESISTANCE, HIGHEST DUCTILITY. LOSES STRENGTH ABOVE 250" F. GOOD

A356 CAST

WEAR PROPERTIES, BEST CASTING FATIGUE STRENGIH

A357 CAST

SAME AS 356 BUT + 11% Stu, + 17% Sty

249 CAST

BE-T Stu AND Sty ABOVE 3E;3 0 F. POOR WEAR. Stu 20% ABOVE 357. CORROSIONAND FATIGUE PROPERTIES POORER THAN 357

z

SXA201.0

L

CAST

BEST Stu AND Sty BELOW 3500 F. CASTABILITY, WEAR,

(AMS 4229, KO-1)

AND FATIGUE ALL POORER THAN A357

2014 FORGED

MOST FREQUENTLY USED. FATIGUE1 Stu& St AND FORGEABILITY ALL GOOD.CORROSION GOOD, WEAR

_

1

GOOD BALLISTIC PROPERTIES. RELATIVE TO A 357 Stu UP 10%, SN DOWN 7%. FATIGUE STRENGTH LOWER, WEAR AND CORROSION SAME AS 249

224 CAST (AMS 4226) 5083 FORGED

_

_

POCR

4UJ3 FORGED

BEST WEAR PROPERTIES, FATIGUE LESS THAN 2014

AZ91 CAST

MOST FREQUENTLY USED. GOOD CASTABILI TY.

AZ92 CAST

Sty HIGHER THAN AZ91. OTHERWISE SIMILAR CASTABILITY. AVERAGE Stu AND Sty HIGHER THAN AZ92. EXCELLENT STRENGTH AT HIGH TEMPERATURES. RADIOGRAPHIC INSPECTION

-

LOSES STRENGTH ABOVE 250'6 F

-JEXCELLENT

ZE41A CAST _

_

_

_

QE22A CAST

DIFFICULT

BEST HIGH TEMPERATURE PROPERTIES RELATIVE TO AZ91. St, UP 9 %npStv UP 6i,. EXCELLENT CASTABILITY. RADIOGWAPHIC INSPECTION DIFFICULT

51t - ULTIMATE TENSILE STRESS, psi Sty - YiE.D TENSILE STRESS, psi 4-2.4 SMCIAL CONSNDUkATIONS

Akboqh many special displines affect dai e-

Ss

m ru

t aniemce indicates that two of

pFm-MgMmqporia

an vibration conrol aW dia-

nioka. 4-4.1

i

C

di"s amum -enkumul

dymmis awe treatW in par.

5-5. AMNCP 70&-31. while addia conskderations of oin Soveraiq. owlap. smd dampig are dotaile in per. 8-7. This panqraah addresses probIsmi awmedap with componnt rucmaat vibration,

particularly as it affects urav membefs. Remnc

Army

helicopter RFP specifacatiom have gated that ea remnant fresunis

"m be a minimum of 30%

away from the design contiuou operaing speed.

Wbihi this may be impossible to adcieve with sonte Wars whem all vibration modes arc onsiere, the intout may be satilaid for the polatially davgerous modes. As an aernatimv suffacient dampin may be

oumpka

t tod -eader -uch

waomat froquencies harm-

loee i.e., the vibratory stressia will be sof" blow the amdiraam limaiw for the sructure. As aisag deuia saeds for drives have rcsuid is pich bee vs4otie above 10.0W) fpm. many

*[

J•

AMCP 706-202

frequencies readily may be identified. In many infatigu, failures occurred that initially were attribinstances, designs will exhibit very low vibration due to However, loading. tooth to dynan.ic uted the favorable mass and stiffness configurations of the aestigation revealed that the fatigue nucleations flange, web, and hub, and therefore will not produce usually were loes ad in the bottoms of the tooth roots a detectable strain gage output. Considerable input or (n the insides of the back-up rim. The crack energy may be required for a realistic determination. propagation gemnally was radial rather than acrtss Although actual operation in the transmission is the the tooth ban, resulting in the loss of a large sgtinal arbiter the acoustical siren generally will promeat of the ger rather than a single tooth. Such duct sufficient input to get the job done. failures are typical of resonant conditions in which Audible detection ah3o is sufficiently precise, althe tooth meshing frequency or one of its harmonics though a microphone feedback coupled with orthocoincides with a particular natural vibration mode of gonal axis input from the exciter into at, oscilloscope the gear. is required to produce Lissajou patterns in order to often use Lightweight gear designs for helicopter distinguish between fundamental and overtone will exhibit various types of vibratory modes, such as responses. The observed standing waves are the with radial nodes or circular nodes, singly and in product of a forward and backward traveling (with combinations. Typical vibration modes for a thin respect to rotational velocity) wavmset. If the gear is Ref. in given are web spur gear with integral shaft rotating at a given speed w,two resonant frequencies involving that is ofconcern mode 103. Generally, the are obnerved for each static fundamental radial vibrawaveaxial into rim gear the put that nodes radial tion mode: form vibration. The lower orders (say, up to the fifth Forward wave natural frequency diametral mode) are more likely to involve higher 4 - f. + w/2. Hz (4-M) amplitudes and, hence, higher oscillatory bending stresses. However, relative resonant response amplig* .ard wave natural lrenuencv tudes for constant forcing input intensity at various diresonant Lrq!.%eicr vary enormously with fsei Sear blank co•r•furations. One gear blank m-ty s2. W (4-5) -respond most to a third diametral mode frequency where while another o the fifth. The flange, web, and hub design, all influence this relationship as well as the static resonant frequency, Hz number of radial nodes ratio of higher order resonant frequenciea to the funfotitional speed, Hz darnensal. The resonant frequencies ma best deA graphic presentation of the phenomenon is contermined by experimental bench test techniques using tained in Fig. 4-34. The fundamental radial node only the gear in question. Excitatki can be by mitihanical shaker, acoustical siren, or electrostatic resonant frequcies are drsjgnatkd on the ordinate by the number of their radial nodes. The abnignealy by an induction coil mounted very n;ear WM i's r.tinnal aerea anre-Inliv tfn nnr..al prra" the rim surface. Excellent visual de-trmination of tin speed a The inclined lie represents the gear mesonant responm may be accomplisbhe with a"ed con-orI_: if the gear web is ofasuitable teichniques pattern freqecy for a 41-tooth pinion melhi bdlanino"da otheinsitanc~ebsfmodal fPsraaaon.~tt fiuratio:. ,I' otw in , nodal and antinodaf driving at a normal speed of 20,000 rpm. Note that maw r-.y tic clearly dettckJ by manual probi g of the forcing f'untior represented by the pinion tooth int the forward traveling 3 node vithe rim a•A- %-vbsurfae with a lightly hand held soft brion nesr ground idl spee, the backward wagv S1fl ly,. AAible detmctl 3n is also sufficiently precise meh -

3% speed, the forwardw0 fromts I0 node at about i fkw mmsamt f•quemcy identification, although a nord wave at ner normal opeating speed, and to1 microphone fedbacks coupled with a radially op. The 3, andtOW at ovespera poed imint fromn the exciter into an oscilkoscopewsrd 1isnode wave 12 node vibrations a- all penmtially hazardous. pod, between pftetalu nandary to required Even if it wer de.or.rated that the cyclic stresses disuimaiah between fumdayvental and overtone wer below the material enduran.cr limit, obMasngne segnawios mqay beased.e reeponac jeetsaable acoustical enrgy radiativa would occur feedback phaug sial with equal -e. Whaen wo vibato amphiudes arm sf- a thee iatercb. teSig of the shape and mats of "ficiumt to prodm wsiificant stress levels, te most the ger is not very practicable in this example bocamu a 25% clhang in ear rim and web thickpractsi qumaiative etvaltutim can be made by ataass affus maxiamur chag of 3%in resonant ta, bamstrain gag to appropriate antinode regigns frqiency for a specific deior- (Rdf. 10). f tshe rio anid web. The hige• emeI lvl rsunance

-

N

-

SPIRAL BEVEL GEAR

o to

to

9

a0

11 a

aX

go

IN

11)

DRIVE SY$TLUS. 'U -L 1hU LOflMAT.OSftEoE___

Gear Teoth Mebing Sped

duction of sufficient damping to vudmn the vbrmimo sigificantly. The spiral damper ring or snap ring a shown in Fig. 4-35 has promc to be efeti1ve in many inisaes,a producing damsping ratios of 0.04 tog 10t With a resultat reductiont of Cyclic resonant arcs to 30-40% of the undauped mugnitods. A*16m.r

r

-A-

=

--

SW

I

1

-__IUI

dA-P.O

otte ty, an of damping (viscolstc ard torsional absorbers hat: baen evaluated with pautial mance in bulampte tflnitia s. Certain vnecdmiet damping tremcaat have boen shown to reduce some Var vibrefion mode amphwudss by 50% (Re. 105). -hruhadcapeesv

"AJ Dipm~estotchnique Many cockpit indicators warning lqlns. and gpips lit ths postal deflimic.o of daagaswi aids. Recet Army REP specifiations icmii indications for oil presein and Impsrathne low pesusaa warning, high Itw~mqaeri. warming. aid quantimty, and chip dea1e90m. Atddmitoa grouned iampulae wfchanquu rutial "hinc Rd mpending Oil-rakertypass warnin Rw oil beaings chas viswa! and am;l faduke deaL~a and lacti, and oilsm.. hag for spectrgraphic analysIts. However. the n~ thats by theisles haeve prowcn inndsuguaee to allow the wiae*rod OV saf and cast-daffve conditiontal

4-70

________

p~ 43

rypsal Sis oeinp. King A"Undimsim.

mnumaafah on existing Army helicopter drive subsyitwus Tou aitpaextent. the mussing ingredient ms consideration of these in the initial design and development test phase. Safe and cost-effective inapismantation of condational maintenance methods requires thorough definition mwand n ate ume of early failurt detection. digagnosisad pro~wmoa as defined in Ref. 106. Majo effort underway in Army-sponsoed programs sam aimed as reducang the time required for deei*om and as immproving sceduling. This philosophy rewts in sebmlsedaisg of unscedul~ed maintenance provided that eqAp~mnus-bae judgmients are availtli aommarig the secrity and mi of progasic of t4e deated fatilur.

-

-

--

~

)

J

Detection methods may be classirted as related to the internal oil system or external to thc oil system. and include: I . Oil system dependent: a. Spectrographic oil analysis b. Electronic or ekLetro-optical oil monitoring c. Oil filter differential pressure d. In-line electrical resistance filter grids e. Electric chip detectors C.Magnetic chip collector plugs. 2. Independent of oil system. a. Naise analysis b. Vibration moititoring c. Temperature mma.urement. SpectRogrphic oil anslysis has pro~en expensive and relatively unreliable because poor correlation exists between faiiure severity and detection. Little is vacually known concerning failure, therefore. diagnosis and prognosis ame impossible. Oil monitoring rouains hr, the cvailaatiora stage-. some techniques havig proven totally useless while others show sowe promise. Oil filter differential pressure can be correlated quite wen wid, iaiiure progression rat;ei of bAriqp when the rate of chanae of differential pressur is examined,. Usefulness of this system iscomprcmis~d. however, by the accumulation of normal wear, dirt, mad contamination parlicbs. Visual inspection of !he filter elemnwt debris by t. perienced tochniciantn can detet mine ahe generic failu": mode buta not the location. F~ilter grids arc best cmnpoyeJ cc a Wahion similar to use of oil monitodriat. The hasic diffacrice is ttal tnefriter grid paitjaz sw~Albr con..an~inant particles. hopgfully rstainiivi, Only the laiýgcr metal fudturc flakes. The of.s~ @1ri cioggaig is detected and may be coVCtvete Cteri(Mlly to Itrae-of-chiingC display. Chip detectionis have bftc used extensively in Army helirvpiors for two dftades. The debris parlidles ame atpturod a~agaetktly to bri~c electrical cnotmc points, which, in turn, ensrtize a caution fipt. The wsduincns of this metho is dependent upon the locaition of tkm dated". 'siA W ~ resulting crew action. failure indication often rest~ts from thr sdow accumulatioa of miormaa wear debris. and oftencn in awiteraw t oftn roosi te cew reult ina itisio, aortto 110vssttpaW the wCas. Use of chip colhctor phigsu eUhaiats~ an aIItl4A at eatr i t~ ie~~es ~iway lawi.nsmbmof plstwith greater deris starlac capacity and W. teual caution panel light aft used to aid in locaiMinof a fiailjr to a specific area or module. liton is aocoapqlished at daly oi Meiodi&c inte-

~A"C 705-202

vals and sufficient debris callcction isusually prc,.nt to permit an expericnced technician to identify the failure mode. Good sucmes has been claimed for such a system on commercial air carrier fleets, but uc~cesa in helicpter adaptation is comipletely dependent uponl wcll-designed installations that entrap and localize failure debris, convenient and accessible delector loca'Jons, and experienced technicisrns to correlate findings with other available diagnostics and to schedule maintenance accurately. Vibratioat monitoring and noisc analysi~differ only ini the sensing techniques. Acceleromnetu or other forms of contacting vibration sensor% measure vibration, while microphone sensors measure noise. Substantial research and developmnent effort.s are being mad.; on these systems. Ruiiimentary go-no-go Syak ins with a wamring threshold signal reqtiiring aigine shutdown have long been used in turbine engine installations. The real challeng lies in the electronic signal processing and its conversion into an identification 'and quantitative assessment of the failure mode. Signal analysis is being investigated by such techniques as auto-correlations, Fourier transkfutal Ma51711, jUWFA fiPFtiss

dmbleftj caiaaatL.~.

aross powcr spectral density. amplitude probability distribution. anJ real time dlassificatirm of wav&form by convolution of other techniques (Ref. 107). Tht operational succes of any such rystam is dependent upon its ability to isolate the faulty comprpncnt signature from the background created by otW~~ internal and external forced and resonant vibratians, siske bar~Ak and best frequencies, at an earlicr time tihan the chip detectof, and to retnain onlint to monitor the rate of progression of failure of fth particular component. Onily in this manntr can the remaining useul safe life 'DC pradmad. and maintenianc scheduld wisely. The ultimatae goal of these programs is automatic identificatiott and prognosis as well an detection. This goal requires coanpmehensive test data on all failure modes, both individually and in combination, along with impravad sensor reliability and an analysis system havisv;, minimum reliability an order of timag *d grnLter than that of the drive system being monitored. Tepraacwsrnctsvm, intAa Temtue prctorelativel se wa madsytwrlmpteatur ul toercrtiel integrated bulk messurmncit, seldom cam detioa other than viny advex" (virtualy emominico conditic") faikies. Capbility imnprovxnent roqapinz &multitude of individual conipowats *one=r caxn~iwsd with triw.ofdu nlssnawdfrbs ln ai6ai cukhavife umiays creted (ora selnevriwoni Rapid improvemients iu &

wtset"~ for bvlaaogw 4-71

I

4

0A

3:

At

dfiiv' subsylensntam NWAutkf larEZ4s UPon: 1. Improved kiixia dcsi,;n ra.uilure-forgivitkg d4sigjn lchriqbt" as in ptr. -1.21.2I. b. Integration of nonssay Gdicstk ais& mW gsospr provisions into the oa~nx!~v *tiaWr. 2. Redireca edevelopment ttstirw effort a. iog out of failuremi~ as~d MdWscImd in par.4-14.43. b. Compalat of pwogrssicon rate data and correatio

lvel.gearbox, wi

3. Implemen tasiono orfie ao~an

f

efectuive

un operating

of saI

a.Deemntio t0t

umn for

cvrow

Wa y

ar b. Dlriaion of. ipctosc failure modt4s from exciessivc continued operation.

43DIEvs, INTWRCONNECiT SYMFMS In a halicopte trnuasmncui and dc~lr utri

PRh,

~

kW i,

is erWnx1csa.'d

t% X~l

fth.

.&flaygtrI, CflCI'i

thfa wu.4&t ac- a ge.ýrb*Ný mad f10o¶ a itiria

to a

k

shafting must a-ecohmotlatt. These mounting systems may be giou;-W conveniently into three

0v

If ý-.

categories

Engine(s) mounted dirccly to main gearbox 2. Enin mone to irfa with tubulr strut&,, usin rod-nd bearings to relieve thermal and load deflsctious, wish the pzwrbox bolted directly to the airfrmie

Engine moumbx as in Item 2; gearbox mounted flexibly. The rotor pylon usually is integral with the that is. airframe mounted with elastomcric3 springs and hiSWe liks or struts. Syrteni I generally usa an internal splined quill be ~~sha;'L lubricated with leaf box oil. Thezspliuics mnný. hardened and ground, slightly cio~ned of straiZln: or indium hard and bobtstd, shaped, or roll-foamedt. yso Wa n rtn orenoyL tu* pouting and/oir sof, plating or coatings such as xiinickel, nylon, Teflon, or racolybdesun d"isf~c vhce Inva~a~sarcqpiýazsacuta for Systemn 2 can bc moc&

-".? k

W.psrdi

s3 m u~nd wu[!ciiAirgi- with rralaivcly$k arith citteve& Aaft1 Ul., retc.¶vrI tac

t~

rur'dMg*e

rtc"-n I 40r'r

ThOC 606010a *AraesI. W to pCci&ubCdyntamic axial defecion b vatiatiozas isctakic kugibh due to installauo-r. coj~~~.bearinsr, and benisl wukmt (hangusý 44,3CE~i~tA tL~J~a~Cmi&i;iwbeflL06)O tx4 x;tho&Si'T growth of the engine must tc &lcdticuor V-1mSs~uatW

Thia ncuvia4 darv w2ita tgks, Id 01prupsLa. t

¾

SjLa&Ii driveL uhaft rsqtirwaco M an 41 &.umin4 rv-*iwa~j kurn~od.by Sp.Aizdrveof m ,b.tparticular heicopter- dc4A.ý. iincu&ni thet &WA.t wzseý.' cuup.a. hagemiw vtgiu slce. w dftz,. and otAýt' VoJzlh)iiftio 6cweit #i. GO~I "P,

t¶haft dzsign. Car must be taken to stiructure provides support at wmtua te t Ota ew~~ii and gcarbx attach points adequate to ddfecion. It often is possible to comply ith ujhkzut-#ad cra'al loci I2IwtDZ Ittjuircnents.

pvamasily by the poi-va scqjamnsmee.L of Otnt,~i llo~ion to b: cactaxI vu ttt.Zity. mwvtIa~~imnbiPjA~y. %4 ik. 4 amuaAuoL "Iat r~&1% w.4da. ivw*fmnaeuta for relinbihiy. maiuwe discumrJ in par. 4t tnio.64Ni y~.m swuvability

)%AEVtM4Wr rctL'iVdy hilh dcfleciMý mjoA4nOrU# aitfrauW nw&I. hi. jgivcst loi tr4~-w;. v=OV sad Ag4udc&~tirww of I dtjý sa k~is nky arc org-

4Gc:%Sai

1.bM

Wdz t Lbagt all-iacuuimc ncWjuiraerw am, tswtrA pe641C"kb vahal.duab"'W iue

In

to

ftlrewijuurtmt

O U4NAIMC w ang

q,"dksc thi~t kuaksm aw-4dwcast in I~c pwar-

cow&aLUt4 byZ~il did: (Thotuam type), r"i-A 49g l mxicn tentoson ck dizra--st (ILandira t 4ixaq,menz Jbalaher tyipto,wcrtnwm4 tý.'kA gear ccupiw'L."ý The first 6.-&typ" SPw.a&Ny asf. preferredl tnx45iwc

t* &ru

Sysan YconVW~t

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d ni k~~ix*4. liaidt i~ aflU 4~fa.n~~aatat~tpaa~ysaa~n

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ottstk%1=44 dndbcntkWC in 014f

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Iubrication.

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The an junt of axial motion that must be acas functional failure of the shafting becoomes a catascommiodated usually will determine the type of tdrivc trophic Walure almost immediately, with acollison of shaft couplings used. Alt known drive shaft sysniý;s the intermeshing rotoi: blades. Achievement of the offer a resistance or damping force opposite to the dinecessary level of reliability requires detailed conrcetion of axial motion while triasmitting torque. sideration of operating xtresses and margins of safeThe maximum acceptable value for such forces usualty, critical speed margins, number and type of dyly is established by either the rotor vibratcion isonamic components (such as bearings hangers. lation system or the engine PTO desin specification dampers. couplings. and splines). rodundari'.y in limits. Damping force characteristics for various mountaiag and support itni~ctuMe and easw of incouplings and spline combinatio~ns are discussed in spection. Criticality of the interconnect system allows-f par. 4-3.2. 1. little latitude for rchal'ility trade-offs and2 comproOther input driveshaft design criteria are governed wise with weight, cost, and m~Aintainability goash. by maintainability, v~ulnerability. and reliability Optimization of design. then, must be in the direquirements and by additional engine pTO design rection of minimum number CL parts, low stress (high and specification limits, margin of safety), and Whb tolerance to ballistx Maintainability considerations require that the damage. Therefore the drive shaft tubas will be rc-laengine-so-gearbox shaft contain "quick-disconncct" tively large diameter, thin wall, and long (within a featur s. Ease of accessibility also is required to facilisafe buckling) length/diameter (LID) ratio and crititate drive shaft inspection and servicing, and enigiric cal speed limit. Intermediate bearing hangter design or gearbox repi xcement. Since these tasks must be must permit relubrication, with ready access to the performed at the direct support level. the absolute whok, hanger fot- visual inspection. The selection of minimum of special tools, fixtures, and skills should drive shaft tube material can noxuitate further conbe required. siderations of a"Ia motion due to differential expansion between the aifnframe (generally aluminum) Although vulnerability and reliability have been dtbmUsseJ nreviousiv.i is irritortant to co~nzio n thec 1104wtfvesbait (Step], aminu tivaniuamo comsequenee- of drive shaft failure. The large kinetic composite). Airframe deflections due to flipht energy of the input driv shaft categorizes it as a maneuvers; or load distribution also can contribute to potenitialfly hazardous o. lethal okiect should it the axial de~ections of the drive shaft. These deseparate at eithe.-or both oF the engines and gearbox Ahctins will necessitate couplings capable of abadapter. The ueof auitiflail devices, i.e., secondary sorbing the anticipated motion. if axial dameciotsi components -or structurt capable of capturing a failed ame small, then flexible disk couplings frequently are drive shaft. is highly desirable, the choice; for larger axial deflections, the geared In addition to limits on &xial force specification at coupling or ball-apline disk combinations are Watter the engine PTO pad. allowable moment and steady uWted. Under any specific set of requirements, the and osiLlatory radial Icads are usually slocrrned. primary 4esig empLasis must be reliability and Sinmc egine-to-geautox shaft rotation speeds are in ready-access fra sevice and inspection. Ow. rang 6W-20.GOO rpmr, compliance with the osciilawory lod limits generally require kincoatic and 4-W u u Satn dynaniic balancing of the individual elemets of the ~ satsse ~do igemi oo ThdrvsafsytmLadoasigeanrtr "trve shaft assmbly. When positioningt or locating helicopter to power the Ladl or antitorque rowo. "mai tolerano. between mating surfaces or elemcnts suc, couplings. adapters, and ahIts cannot be conbeft-een the main gearbox and the tWi rotor gparboa. This system smSt provide power to the taul rokir ay ecot.- ~ccsar to tro~d aequt~y.it masn y toelominatessarym th mi rotor in the event of loss of drive from cgitd trolle adeqaey balnceth mialy covleeo inuateexessve thme engine4s) In ormal operatio, the en~gine*s) vibration.driv thrnoug &fraewheeling dhs-* to tie main yewibox. During autorokatiom. when the fre-whselir4 "411 EUdim and Shellin unit is overrususing tW1 rotor power is extracted from An interconnec shaft syw .n for mul, Ole main (or the siam rGto adlorotAticnal, or kinetic flyweeL.iq.% liftin) rotot Wkdhpters transmits power between the artia. engine Vearbox (or the collector gearbox in multiTail rowo drive shafing win be subjacted to severe eq~sa eLi6copters) and the main rotor goarbox(es) trnasient loads and cyclic tersional ouciasioos as while also maintainin phase reiationshp between wagl as noual Moeady torqe inputs. Torque require rolors. The pitaary comuiduatioas for sisca an intermuass for moat flight conditions awe maderia is 000toc sha4 s"ste ane reliability amd survivability, nawne wit maxismum steody torqu requird during

4 I

-..

'A%

Ias

'

4-73

'r~

Sa - V.~

4

bover at high grow weight. The total power roquired to he ver ismain rotor power plus tail rotor power requkied to offset the main rotor torque, plus losses, r.A taW rotor also must counteract the main rotor cy,l4ic thrust vactor and an aerodynamic drag couple from the taillboomn. Conventic-sal rotor or propeller theory. including an effacieca factor applicable to A*. specific tail rotor can be used to, calculate the Mmady tai rotor torque. However, experience has shown that the transient torque requirements can be from 200-400% of the steady-state design torque. High levels of transient torque result from sideward flight in an advems quartering wind, from yaw acaelerations, and from unusual inik , conditions ivsulting from combinations of main rotor downwash, tail rotor blanking from aircraft structure, and adverse winds at hover or low flight speed. Transient torque inputs also cen be introduced to the tail rotor drive system by engine compressor stall, violent flight maneuvers rapid throttle movements (chops). or abrupt engine power loss. Under such conditions the abrupt relief of the windup of the tail r .4or drive-shaft combines with the flywheel inertia of the tail rotor and with secondarv effects of main ru~ui iwlin-at to cause aeveral cycles of extremely high amplitude torque oscillations in the drive shafting. Although occurring infrequently, this low cycle-high stress phenomenon can cause fatigue damage to thc tai rotor drive system unless the components of this system are dlesigned for torsional loads well in excess of the noirmal steady powter re~uircments. Transient design ctitcria for the tail rotor drive itipulated in MIL-T-595S and AMCP 706-203 are 300% of the power required to hover at design gross weight and denisity altitude or 150% of the maximum power rcq11irnd

in th~meet sver,. mzne-otuer .;.th;n

fi

infinite life critria. The throttle chop transient response is often the greatest oscillatory torque felt in the T.R. drive syrteni. The system can be modeled for the computer iiiring the cnginc-main rotor decay curves, the appropriate lumped mass and spring rate analogues, and the coupling discontinuities. with reasonable accurascy. A prcproduction flight strain survey will provide sufficient infornmation on the torsional charsctcris-ýics of the tail rotor drivesystcm to enable substantiation of the integrity or revelation of the unanticipated weak points. 4-3.1.4 Suherltlesil Shaffing Analytical methods for determining critical specds of a drive shaft are covered in Chapter 7, AMCP 706201. As defined there, the critical speed is that rotational speed at which the elastic forces arc overcome by the unbalanced centrifugal forces and the "bow" of the shaft increases divergently. Theoretically, the critical speed of aperfect shaft, i.e., a shaft that is perfectly balanced, homogeneous, atid equally displaced about the rotating axis, will occur as predicted by* analysis. The behavior of such a shaft is . dashied line in Fig 4-16. WN vihrstirnn shown occurs the rotational speed 17approaches the critical speed n,, where divergence occurs almost without warning. Practically any rc I shaft has some initial unbalance that provides a centrifugal driving force which increases with increasing rotational speed np. Suich a shaft exhibits vibration /rotation characteristics such as arc shown by the solid line in Fig. 436. While vibration levels at normal operating speeds

flight

envelope, whichever is higher. Such requirements are rather straightforward withFISCITC[F'5)

respect to fatigue design of the gear teeth and the can., tilever rotor shaft in that the need for infinite life criteris due to the high rate of cyclic accumulation (rotation speed) is evident. However, with respect to the remainder of the drive system, where start-stop{ cycles, throttle chops (T.R. inertia overruns), airburm engine restarts. and yaw control pedal excurvious accowad for the bulk of the high stress cycles,.____ a far lower frequenicy of accumiulation exists. In such instimice, past experience iv ith the fitting of theoretical spectrum analysis to subsequent flight strain survcy results is rather essential in efficient design work. When fatigue spectra are unknown, a ,

rewaable appiroach has been to design static yield

strengt lev&l to a minimnum of 3 times thc transient fatigue stress used for the gear teeth and rotor shaft 4-74

~

R*F [C 5 fVSHAI TyrCAL SHAF1

-I

II

,

~..~.I_____ tATSAF'PE

*,,,

Figue 4-36. Relative Shaft Speed vs Relatlhe Vibration AmplItude

V P" may be acceptable. the unbalanced forces can increase rapidly as i increases above 9, with the possibility of resultant damage. This situation can effectively reduce the critical spead margin to an unacceptably low level. The inference is that simply by balancing the diive shaft an acceptable critical speed margin easily can be realized. However, the ,ast of dynamically balancing the bhat and/or shaft assembly must be included in the trade-off, together with a careful assessment of the contributions of the end conditions and/or mounting compliance to the vibration/rotation characteristic, The majority of existing drive system applications use subcritical shafting, for which the lowest value of It,, > V,. Requirements for balancing can be met with ordinary balancing techniques and equipment; relatively short shafts minimize production and logistic problems; and ballistic tolerance design paramcters arc known. On the other hand the cost of a subcritical shaft installation with seeral separate spans

BEARING (1 SEAL)

may be higher than that of a comparable supercritical installation. The i:,anufacturing cost fot the short shafts may not be much different than the cost of a single long shaft, while the number, an-' hence cost, of machined parts probably will be higher for the subcritical installation. A single span of the subcritical system consists of a drive shaft tube with end fittings, drive adap'cr, hanger assembly with bearing. splined adapter, and coupling. A typical example is shown in Fig. 4-37. Design of the drive shaft requires a determination of the shaft cross section necessary to accept safely the steady and transient loads stipulated in tic pertinent design specifivation, and of a shaft length that will operate safely within the critical speed limitations. An efficient design generally consists of the least number of spans with acceptable critical speed margins and torsional buckling strength. Large diameter thin-walled tubes, generally of nonferrous metals; a greaw lubricated bearing sealed on one side

BOLTED JOINT ADAPTER

FLEXIBLE COUPLING

TUBULAR SHAFT

GES

HANGER SHAFT

NUT

ADAPTER 3llf

706-202

//

_aHANGER

ASSEMBLY

S......MOUNTING BOLTS TUBULAR SHAFT

\

TAIL BOOM STRUCTURE

Filme 4-37. Typical batrilgl Haiser AMedy -- Sedicfical Shaft Ase-mmy \

4-75

with a fitting for a periodic relubrication: and a flexi-

shafting will be partially or totally offset by the addi-

ble disk coupling arc typical of current design practice. The end fittings and their attachment represent a considerable portion of the cost of the manufacturing of such d'ii-a shafts. The fittings may be attached by adhesive bonding, riveting, boiling, clectron beam welding, or brazing. Tolerance of mating pairts must be closely maintained to ensure good parallelism of end fittings and low vibration characteristics. The type of ccuplings selected, their mass. Io"cation, and friction characteristics, influcnce critical whirling modes as well as torsional response modes ofthe shafting. A recent investigation of coupling nduccd whirl phenomena on turboshaft powered helicopters is given in Ref. 10S.

tion of a daniper or dampers. The elimination of hangers and the attendant maintenance requirements also may be offset by the addition of maintenance requirements for the dampers. Acceptable tolerance to ballistic strikes requirs hardware testing under simulated service conditions. Parameters for ballistic-tolerant designs for supercritical shafting have not been defined and dependence on individual tests is almost complete. A method for calculation of critical speeds and bending modes for high-speed shafting is well presented and explained in Ref. 110.

4-3.1.5 Supercritical Shafting Supercritical shafting usually opc. "-s at a mced between the first and second critical specu of rotation, although even high orders are possible. "The main rotor and tail rotor shafts, or masis, often pass t hkm

lo

h

•

fi e t

PrituCjti' .

snrp

h cfýn r

d

p r

opr h in n

operating speed. However, the critical speed is ,elatively low, the dwell tin-. is momentary, and aerodynamic damping forces are quite I.. ge. On the other hand. interc.onnect drive shafting and tail rotor drive shafting generally operate at relatively high speed with very little inherent damping. The advantafes of a super.;'itical shaft design aic the smaller number of detail parts and bearing hanger assemblies. The disadvantages are the need for dampers, which for reliability should be redundant, and the...... physical ~length of the"'• !hafts, im"-I:*-" ~ other .. . which ,,t• "•-may "- r"-" IRO"•*

VBl

IRtOKOl1t%,3.

/'%13V.

9

.Q1 201

6

13#0

%-41J11 10.%,

tars i'ay dermine shaft diameter and wall thickhe~s. Shaft sizes larger than thosc required by the power requirements m4N be necessary to maintain a LID rptio sufficient to avoid critical torsional buckling, or to counter a ;pecific ballistic threat. Once si.e has been determined, the design requirements for shafts in the supercritical speed range center primarily on damping und dynamic balancing.As shown in Re.f. 109, the ncessity for balancing to a very clo tolerance over the entire span er the supercritical shaft is paramount for successful operatio,,

I

4-3.2 COMPONENT DESIGN The basic drive shaft system components, coupa.gs, bearings, and shafts are discussed separately in the paragraphs that follow.

4-3.2.1 Couplings The primary purpose of the shalt coupling is to p r

v d.

..

4 'f

n

g ut|

- m

i s aah gnm

en t

an

di

x'; al

motion between various shafting elements and the engines and gearboxes. This relieves stresses in the shafting, bearing, gearbox, and engine components induced by bending moments and axial forces. The rcla!ive motions between these components may be due to airframe structural deflections. thermal expansion. or pylon excursions required by rotor vibration isolation schemes. There are six major types of couplings that have been used in helicoptcrs, and the selection of one among them for a given application depends p .. .greatly . ....upon the required dis.olrt ni paccimicrts &no thle ioaos that can be tirid

i

their supporting elements. The six are: I. Laminated flexible disk couplings (Thomas type). This type of coupling, shown in Fig. 4-37, is probably the simplest design for angulai misalignments I deg. It has been used oi the CH-47 synchronizing 5haft and on the OH-58 tail rotor drive system. Eachi driving sp;der may have two or three at-

It is incorrect to assume that a supercritical shaft-

taching points (four or six equally spaced holes in the oisk complement). rhe larger number is preferred from the v:ewpcints of vulnerability and survivability. The laminated disks are generally circular rings, although square and hexagonal shapes have

ing system will automatically weigh less than a subcritical system. Dircwtly comparable designs for a given helicopter application have to be made and the total instalted weights dctc'mi.ed accurately and compared. The weight saving apparently achieved by elinminating the hangers necessary for the subcritical

bcen ustd. One problem that has been encountered is disk fretting at the bolt attachment. This coupling features high torque capacity, light%eight, simplicity, and constant angular velocity. The torque capacity can be varied easily by the addition or deletion of laminates. However. increasing the

4-76

number of Isminatas reduces the angular misalign.

WELD JOINTS

ment capability or the particular dcaign. An ad-

ditional rclationship exicfls between the numbenr of &t-

tachmieat puiats and thc- torque capacity, and misalignment capability. A four-pax'nt attachnent (two-bolt shaft adapte;) provid.es the maximutm misalagameilt capability and Odso is the least expersive to mnufaclare. Thc flcxi~'le disk is capable of small axial drilicctions, and where predictted axiWa motions

LXEDIPRG ___

P-e low, this cotupling servu well. No lubrication is sliding involue spline to i-Zoomicouatr

V&rlM1jO1Fl; ;,____

initial shaft aw~mb!y length due 'o accumulation of

mrnif~triig~lcrancec. b-lowcicr, &.'x.n amial de-

-E'ALIGNING MONOBALL

i%4fion occurs tindcr operitting torluc, the dlip rt;-

sistanct of Itbe spline~ is so great thsit apprecia~ble axial .'orcc will be rploliceJ to the disk laminates. For splines of 0.is type the brealhaway slip force Fju rarely is less k;, -

j

.4 Q/D,..lb

(4-66)

-

Q ~-tcrque, lb-in. =pitch

diamttei, in.

In some special cases whc~re certain dry Nali lubri(.atiteris are applied to the splines (Ref. 93) break-

away forces of half this value masy be realized. 2.leibrdiap' ,ai (Scndix tyre) couplings. Thibcoulin (..e fg. -3~))~s boa ~e onthe OH-6 hel~copter. Thi, type couplir~g is gencrally capale f arnaimim ~aula mialinmci ~ ahout idey per diaphragm pair, burt very little, axial def~tao. ad~apai~iy thtefre ~a~st e ~axial a sl~1i~ Recrcultin i' i breaawa splipe i~ o it ih wiTC ser.CIes v snip uisbrekawy -agdmt stac pity ar ge'~ll force. F6, A~sj, the di~aphiag tc 5v~ snitv stoe faigue w~lueakt tocme oslltheytax ial eloftensol caai ar.d provide r.waly all of the axial deflection. ay e ncesaryto roUndkrm;;hconitin.%it or saybeneesarytype pro-i onuai coifaligin vi ae atideia imcnf- ige ex r centcbll(roftesimilar typ)baring at he n~cri1e~f~urcceterof hestak ~ tanscr %xialloids into the s-ipporting adapter and tofoc the ball spline to move. Although theoretically it i5 possible t&ýobtatin v~~'y low breakaway slip forces wiih the ball s!apractical considerations with, itspev-t to aainirr.ium length of the ball track groo',ea, vh'ntbevt of balls. and stAl p,!cvisions usually limit. 1hese rorce toa a r;aininium of .,

..

J's.- 0. 15 Q/D. lb

(4-67)

DRIVE FLANGE

Figure 4-3b. iiictl~ Dispbragai Coauplham All too often, the vibratory forces due to imbalance and r.ý;or vibrations lead to false brinelling of the ball track grooves, which results in turn in much higher forces w'Ah increasing service time. Althourph the flexible dianhraffm elemouts need no lubrication, the splines and the monoball occasionzily require lubrication 3 Axial~y loaded straight element flexi'hle coupling (Bossier coupling). This coupling (Fig. 4-39) requirts no lubr~cation and has the abiiity to accommoda,.e combirned axial motion, m-isalignment, and torque. A series stack of warped rectangular plates with a vertraliy located reczangular cutout leaving slender sides chatacteeiz*5 the coupling. Opposed corners of these plates are b~olted to adjacent elements and to end fittings or adapters. Design characteristics o'hscoup'ing arc defined in Ref. Ill1. Apiain odt hv enkgl xai m-.,nta; with some flight time accumulated on the H; -2 h~cpc (Ref. 112) and UH-I- helicopter. Thec atngular misalgnament cipability appears to be about 0.5 deg per ,-late elemen:. However, an increrse in the nu~nber of ýesments usW~ resttll in a teductinku in first 4-77

A

--

whirling aitic• ]s apad. A satisfactory lightweight design for an engiae-to-gsasbox shaft for mod•a•ac angle (of the order 2.5 des) 0.25 in. scillatory axial motion probably would be requird to operate in t super-critical rune if tte erine ohtput speed were above 600 rpm. 4. Elastomeric couplings. Cbouideeabl developme- t work culuinatiag with eaperimatl dight tasting on helicopuers ai s the YH-51 have been accomplished with this Vypc of coUplling (FiS. 4-40). FHowevr, all sucessful applications lhav had low angular misalignment and axA deflection requiremsnts. Efforts to develop higher capabilities (up to 2.5 dqg steady misalignment and 0.25 in. oacillatory axial displacmcmet) have met with failure. The low angle configurations LAw us•d simple rubýber ekmenut in shear or compreasiou. while for h*i _angles very thin, multiple layer, ruber-ametal-rubber combinations, such as are now co•t•ion in certain rotor system bearings Itait ben ustd. ibe principal development problens bzve boee customer fatilue d, -o r•ersed louding (alternating tension/oonpr.,Ao ) at high sagle/low torque conditcr,. The basic advantages of cbetomeric couplings amr his'" o.upiihria- (low shock an.d oie -.............

!ARD

IARIOeECA&ULk PLkTL•

Rfor

-

DRIvEsr'Fr

_-yokes. DRIVEsHrI hEM 'CU•uAREAusWed

COUPLwG ,ADAPTE

Figure 4-39. Bosser Coupling

7

.-u

i

no lubricaio required. no susacfuihuty to fretting corrosion, and polential savings of cost and mnintenance. lanei disadvatages am dereAioralion in an oiy environmint and aging and reduced ctritcl spaed dvi to high compliance. 5. Hooke's joint. The Hooke's or Cardan type of universa joint coupling (Fig. 4"4) is capable of mlativcly •Aigh angular mimaigi•ret, of the order of 30 deg at wadete spoes and I1 do a, h*ig speeds. Howeve, unlike all other couplings discussed in this paragraph. the output is not a cons.ant angular velocity, and significant bending moments arc induced in the attaching adapters and supporting structum.ý. Consequently, this type of coupling generally is cmployed as phase-matched pairs to cancel the cvsillatory angular velocity or singularly with systems that am very soft torsionally and hence can absorb the angular velocity oscillation. The H- 13 tail rotor drive system is an elxample of the latter type of application. These couplings have no axial motion capability and normally arc used in series with either a sliding (involute or square taoth) or a rec'-culating ball spinc. The input and output yokes of the coupling

at..h.. to th cros with cupped nidl-c bearings. When that co..,oncnts arc sized properly oscillation. angular and force givcl torque thel absivkaway sliding• of velocity tlhe adj-Afrnt spline Usually is wel within the axial load capacity of the coupling. Common failurr modze, re spalling of the "cup end needle bering. and fatigue fracture of the The needlt býarings require lubrication. ti. Gear Wtouplingl. Gtar couplings (Fig. 4-42) with higu~ly crowned enternal involute gear teeth :...ting with st'aight twoiaed intcrnal gca2 te,:? have been on helicoot, r.; far more extensive!v than all oiliAoter csmpling typts combined. 1 '.Xsc coo j~iugs arc capable of providirzi moder3tely high anguLr misaligninernt and axial motion during operation at high speed and tot'qur

Ž-tr

v,-ry low weight. Common

opfratin• • 'onditirci &:e i deg continuous and 6 deg ADAPTERRETAINER CUP

k

. ...B[A

SRI,4

N,,

"" I

A

SVAr T

ý,,Ai 1 R1 TAAINRCQu

SDrfIVE

Figure 4-40. listomwek Copling 4-78 I *

DIV

CROSS

ligle 4-41. $o.4e's J-•la (Umimal'il

.

-7

4

"ADAPTER hOUNTING BOLT CIRCLL •'•

INTERNAL STRAIGHT TOOTH COUPLING

BOOT SEAL

GREASE CROWNED COUPLING

Figure 4-42. Gear Couplng

transient 2nd 10.75 in. oscillator; axial motion at frequencies in the range of 10 Hz. Unlike straight o.

spliies o~r recirculating ball splines. the breakaway blidinst

force F,. cait be very low; actwtlly F, reduces

-..- _-

~

as the mi'aa~ignment angle increases. A comparison of th;, force ~o, a typical gear couplin., n.shown in Fig._____I_ ccn ~act becween loaded teeth is at r.*high sliding velocity r~. t; the angular miselignments. Conse, quenly, a stmerimposed axial motion is resisted only by tht rela, w'vely low dynamic friction coefficient rathaer than d static value. G.,-r coupling operating limits are thermal rate than 6i1.3up,- w~hich is the limiting contideration for the ri .- ,oupling typej previously discussed. Specially developMr gresse lubricants have provided the best load carrying (least friction) capability for gear2 couplings. However, the operating environment (high

etrifulal field, high mechanical stroking fire-

qucncy, *ý--d tlevatiA tempcrature) combine to make the -erst ma~jorit) of grrase unsuitable for this app'iitMiss.

-

___

_

n -. a iuiy M I SAIL SPL!III BALL SPLINE A-. *-%-fCftULMEDN

I

___

W -1

-j

-

IR_____

____

IAIUNTAGL e

Figre4.3 Bma'kapay SliWG Foresei NalgmC foe '/arlkm Spiae Dekvkn 4-79

A.side from the need for periodic relubricagion (60D-hr intervals ate common) the greatest difficulty with gear couplings is providing adequate sealing for the grease. Guillotine slider seals and elastomeric boots are most often found in high angle applications while modified lip-type shafts seals can be used for low angle (.

operating speed range of the propelle,. The operating range in the typical diagram of Fig. 5-38 may be seen

to be free of critical speeds up to 8P. The dynamic characteristics of the propeller system

6PI

7,P/

DE

140

defining its critical speeds for the various aerodynamic excitation orders i.e., the rotational speeds at

speed diagram st -h ds that of Fig. 5-38. A'%propellei critical speeds, there may be high dynamic magnification of the aerodynamic loads. Therefore, the rropeller system should be designed so that the lower,

1 1 8P

I I•C~rAION ExciTATI0N ORDEq----, ORDER-v' 8

preceding paragraph is determined by the structural

0 Ftgure S-3.

1000 500 PROPELLER SPEED, rpm

1500

Propeller Crical Spea]Diagram

AMý

1

3@

tho operatin speed raWg with at lemast 10% m rgln. leamt, the values of crtia speed charje with blade anaje due to tht blade twist and centrifuagal effects, blade ange should be conaidered In the evaltuation of reaptionleua mode critical speed relative to the operating range. Also. the effective tetnition stiffness differs for the tiree kinds of propeller modes snhown ;a Fig. 5-39 because of structural coupliPS within the hub ThIis effect must be Included In dynamic and response calculations for the blade. Iii general. the retcntion stik"Nz. h~ lowest for the me actionless modes and h4iest for the symmetrical modes. Propeller aerodynamic excitaiions with.a frtquefl cy order of one greater or one less than integer multiplas of the number of propeller blades combine at the propeller hub to produce bacl~ward oT forward whirl modes of the propeller, respectively. Because Of this whirling action and the rotation of the propeller, these aerodynamic excitations appear on the gearbox-aircradft system as iotating shear and imomfenlt loads at frequencies corresponiding to mutltiplesof the number of blixies. For example, in r' threeblinded propele, ecttooa rqece fZ n ?a w~ elt by laic ovgearo as a or-wnnra.&cu J This interaction of the propeller dynamic systain with the aircraft sy stem must be taken into account in calculating the propeller blade whirl mode critical ONLY OUT-Of-PLAN4E MOTIONS SHOWN IN THE LOWEST EILADE

speeds. This con be done by a complete coupled Amalysis of a rcta&lng fleible propeller atta"he to a stationary aircraft dynamic system. It also cm~ be calculated by first determining the variation with froquency of aircraft system whirl Lapedanoe, e~g., angular and radial deflection oi the propelier shaft for unit shear and moment whbi loeds, and then Including the aircraft impedance in the propeller critical speed analysis. Aerodynamic excitations at frequenck. that are multiples of the number of bladeas excite the piropelIcr In a symmetrical mode, producing vibratory fore and-aft and torque kowd at the uame freluency on the gSa box-icrf sytm Jutaswth thewhr modes, dynamic characieristics Or the aircraft and transmission systam muast be included when smymetrical mode propeller critical sPeeds arn computed. Again, this can be done with a coupled analysis. or by the impedance technique discussed previously. Because the torsional impedance of a transmission system usually is low, symmictrical blade modes that are primarily inplane (putting vibratory torques an tie shaft) will have conside-rably higher critical speeds than would be calculated for a fixed hub. of theJ,cmitriftinl sUMfbaaun cffcct iL. twist of the blade, propeller critical speeds will vary with blade sanle, This effect must be considered in placing the critical spewds pioperly. in general, it is customary to place the lower order wvhirl and symmetrical ctitical spends at least 5%out of tbc nomiral operating range. Less margi4 is needed for thmes critical speeds than for the reactionluss modes because of the much greater damping supplied by structural interaction with the aircraft systemi.

'r- T zOnca SE~NDING MODE.

~

the dynamic characteristics of the propeller

system hove been determined. the maganitudap of Owe response to the various propeller aerodynamic excican be determined. For modes that arc being I'\tations WHIRL excited well below their critical speeds, a real-variable resonse analysis may be used (ANC-9). This always is possible for IlP-acrodynawmic excitation and / sometimes for 2P. System response of the higher order aerodynamic excitations may be determined by an energy method, SYMMETRI CAL using the calculated normal modes of the propelkr and assuming the structural and aerodynamic damping from experience. Another method is to use response analysis, such as given in Refs. 50 and 51, with complex variablew so as to include structural andi L __-Daerodynamic damping. Thec foamer (energy) method uses the normal moJes and natural frequencies ob. REACTIOMLESS from dynamic- analysis of the propeller systan, I'tained and letermines 6ce rsponse of the blade to n particular aerodynamic excitation order by equating the ft"roplle Vatworne Miode 5) Figue -.

-

-

--

-

000011

AMCP 706.202 energy dissipated through damping with the energy introduced by the excitation. Experience shows that the effective overall damping, aerodynamic plus structures, varies with the type of vibration mode, being about 0.02 to 0.04 of critical for reactionless modes, and about 0.04 to 0.06 for whirl and symmetrical modes. From the response of the blades to the various aerodynamic excitations, one can determine the blade stresses, retentioq and s'iaft loads, and, finally, the ',oads applied to the gearbox and aircraft. Certain excitation orders put vibratooy torque, but not vibratory bending moment, on the gearbox; others do the opposite. Also, 2P-excitation on a four-way (fourblade) propeller puts no load at all on the gearbox, as this is a reaction!ess mode, i.e., all the loads are reacted within the hub.

is added vectorially at right angles to the normally considered pitch inflow. Hence, the total 1P-inflow angle is affected. Likewise, the dynamic pressure is changed by the cross-flow component, but for the same gust velocity the wing lift is affected less by a lateral gust than by a longitudinal or vertical gust. Vertical gusts have direct effects upon the pitch component of the inflow angle, the dynamic pressure, and the wing lift. Each of these factors influences the flow field and, consequently, the excitations and loads. As in :he case of a lateral gust, the verticai component is added vectorially to the forward airspeed. This changes the magnitude and direction of the velocity inflow. The current method for determining propeller vibratory loads during gusts uses a quasi-steady-state analysis to evaluate the flow field and aerodynamic

A propeller must have the structural capacity to withstand the combined loading from its response to all of the aerodynamic excitation orders superimposed.

excitations. Although the propeller speed and blade angle may change, depending upon the rise time of the gust, it is expedient and conservative to assume a step change in the inflow to the propeller. In other

S4.2.3 Gaw an Mamyer Gusth and m.ineuvers can heve significant cffects upon propeller vibratory loads. The more obvious

words, the propeller is assumed to be placed suddenly in a different aerodynamic environment without any

change in blade angle or propeller rotational speed, __ .L . ,,- ,.. .,. . p. . .. ,.,i,,.d in the manner discussed in the preceding paragraphs.

effects are caused by changes in the aerodynamic flow

Propeller vibratory loads incurred during maneu.

fields and the consequent excitations to whic'ý the blades are subjected. Secondary effects are the result of gyroscopic motion and inertia forces. Because the

vers are determined by using essentially the same procedures as for gusts, with the exception that the maneures snayr genral are limited toat ose in-

basic frequency of the vibratory loads is the propel-

volving vertical load factors.

is parallel to the flight path of the aircraft, and, there-

maneuvers in~volvintg vcrtical load factors are calcu-

fore. subjects the propeller and airframe to a change

lated using the procedures given in the preceding paragraphs and considering that the effective gross

Aircraft design speciications (MIL-A-8860 series) fer rotational speed, many stress cycles can be acdo not include the time duration of each maneuver cumulated on the propeller during a gust or maneuver. This is in contrast to nonrotating airframe cornnor a breakdown of the maneuvers as functions of ponents, which are subjected to only one major load airspeed. The maneuver spectrum (see par, 4-11I, cycl during a gust Or maneuver. AMCP 706,201) must be available to the propeller Sus' ca co .,e t i -isav .. ........... ,-- n lg-n .... ....... lateral v. Th.....-........... designer so that he can a-.,--- "tat n tcutuc ,,,c u, directions: longitudinal, lateral, and vertical. The the propeller compone,,Lb ,iis be satisfactory. longitudinal, tore-and-aft, or component essentially In the design analysis, blade vibratory loads for in dynamic pressure. The steady torque and thrust on *the blades change with a suddenness that depends upon the rise time of the gust. These changes in load can be relatively high for the large propellers used in

weight of the vehicle is its actual gross weight multiplied by the vertical load factor. Maneuvers influence blade vibratory loads not

V/STOL aircraft because the blades of these propellers are operated at relatively low angles of attack. The change in dynamic pressure also has a direct effect on the IP excitation factor, and the I P-stresses are affected accordingly. A longitudinal gust changes the lift on the aircraft, thus imparting vertical accelerations &ndchanging the wing circulation, which, in

only by changing the aciodvnamic flow fields but also by the resulting effects cfgyroscopic motion and inertia forces In a pullout or pushover maneuver, the angular velocity of precession fl, is equal to

turn, has an effect on the flow field.

where

(n,

l)g V

-

rad/sec

The lateral component of a gust can be treated as a

nz

- load factor, dimensionless

change in yaw inflow to the propeller. The yaw inflow

V

- flight speed, fps

5-62

(S

'

AMCP 705-202 The resulting blade loads can be calculated using a procedure like that used for calculating the response due to I P-aerodynamic excitation. For this analysir, the load is a function of the mass distribution of the blade and is applied perpendicular to the plane of the propeller (out-of-plane). Like IP aerodynamic excitation, gyroscopic motion induces a IP-moment on the propeller shaft, which, for blades having three or more blades, exerts a steady bending moment M on the aircraft as expressed in M - lp., ,fA-lb

(5-16)

where

4,

propeller mass, mass moment of inertia, slug-ft1 W propeller speed, rad/sec Inertia loads result from the vertical load factor applied to the propeller. As in the gyroscopic analysis, the load is a function of the mass distribution of the blad,, but in this case it is applied inplane. The I P shear force F on the shaft is simply -

F = nzWp, Ib

(5-17)

where Wp - weight of the propeller biades and hub, lb There may be other special occasions where loads due to maneuvers should be considered. For instance, a tail propeller of a helicopter may be subjected to large precession rates in yaw while hovering.

J \

54.2.4 Stall Flutter Propeller blades must be designed not only to handIe the applied aerodynamic excitation loads and to have the appropriate dynamic characteristics, as discussed in the pieceding paragraph, but they also must be designed to be free of flutter. Classical bendingtorsion flutter is not of concern because of the large separation between the fundamental bending and torsional frequencies of propeller blades (Ref. 52). However, stall flutter is a major concern because of its potentially destructive torsional vibration. There are two apparent causes of high torsionalblade vibration: aerodynamic hysteresis and Karman vortices (Ref. 53). Aerodynamic hysteresis can cause divergent, self-excited torsional vibration and is, therefore, true flutter. The Karman vortex excitation, however, is not true flutter but a forced excitation. It nevertheles is similar to hysteresis stall flutter and can caue large amplitudes of structural response and possible failure. Torsional dynamic divergence due to stall flutter "Occurs because of the phase lag in the aerodynamic

circulation variation with airfoil torsional motion. Beca.,ac the vortex formation must travel to infinity before full circulation develops, the airfoil angular motion tends to lead the aerodynamic change in moment about the elastic axis. When the airfoil motion and phase lag combine appropriately, aerodynamic energy is fed into the structural system qnd sclfexcited divergent torsional blade oscillation occurs at the fundamental torsional frequency of the blade. Although methods have been developed for analytically predicting stall flutter (Ref. 54), expcriencc shows that a general understanding of stall flutter and empirical relationships usually is sufficient to evaluate whether a given blade design will be subject to this phenomenon. Tests and analyses have shown that stall flutter is dependent primarily upon three factors: the reduced frequency, the blade angle. and the airfoil Mach number (Ref. 55). The effects of Mach number can be combined with the reduced frequency to give the stall flutter parameter SFP. b SFP , ,b, SF, d'less (5-18) aM'w/1 - W where , naturai torsionai frequency,. ad/sec M - local Mach number, dimensionless b, - blade semichord, ft a = speed of sound, fps When full-scale and model blade stall flutter test results for many propellers under static conditions are combined in a plot of SFP versus blade angle, points indicating the onset of flutter form a gener.l trend, as shown in Fig. 5-40. 1 he envelope of these flutter points may be used as a design basis. Although Ref. 55 shows that a blade whose SFP is greater than 1.0 will riot flutter regardless of blade angle or power. Fig. 5-40 shows that, for !ow blade angles, a blade may have an SFP of less than I .0 without being susceptible to stall flutter. Because the design line in Fig. 5-40 is drawn without regard to such secondary effects as camber, thickn'ess, planform, sweep, and center of twist, it is, in general, conservative; i.e., although the SFP of a blade lies under the curve, the blade will not necessarily flutter. In general, increasing the camber and thickness and shifting the center of twist forwurd will increase the blade angle at which flutter occurs. The effects of planform and sweep are more difficult to assess, because the stability of the blade involvts the integrated effects over the entire blade. Thus. although some blade sections are stalled, the blade itself will be stable unlss the integrated energy fed into the blade is greater than the structural damping present. 5-63

I

. ,s. , J

4P., AMCP 706-202 1.1

-

I

j_

S ,0.9

CM, increasing the aitfoil camber is one way of increasing the forward thrust or power at which a blade

-

-

-

-

will be subject to stall flutter. However, in all cases, if the $FP is greater than 1.0, the blade will not be susceptible to true stall flutter regardless of blade angle or loading. The other possible cause of high torsional bltde

0

!

.

5&4.4.1 Helew Dw Steel suitable for one-piece hollow blades, or for the $par of Spar-h blades, arc low-alloy Aws eqwvalmnt to AISI 4350, vacuum mcltied, in both the 364o Rc hsrdnm and the 40-44 k hardness ranges. Thee steals must be protected from corrosive environments. The leading edge or the entire airfoil may be protected with eosion-resiatant coatings or olaiwn with lea-durable paint coatinP on the internal surfa•c. Impact dmage is a serious probibem for a one-piece hollow steel blade. Wall .thcknfa that may be adequate for carrying stiuctural loads may be thin enough to dent locally. The mrent reduction for local impact daner is the combined effect of h.e gouge stress concentration the local plastically deformed material, and the stress-raising action of the dent. Frequent impaction and local removal of gougf and tei plastically deforme surrounding material can be und to proct against the effaft of this type of damap. Anotlnmethod is to protect the

squarencas, axial positioning of the Wade, eccentriblade with a hard, damnrreistant plate Such as city and squareness of the hub retention on tse pronickel or chrome. Howo:r. the stsenth-reducin8 effC.e_ of the hap] ,ostine n the ,eA stru-tu....re tA peller shaft, and spinner mounting runout and tilt. Thc balarnc r,%.Aim-era of the hub and spinn be considered in the initial design. usually are met by removing material or addin,,,,\.JC balance weights. After the parts of a rrosraler are b•lanced RELATIVE "separatcly, the assblWd propeller is bahvi a WEIGHT statically in either the hou•iz.td or vertical position to 0.0003 in. time the proJle w bj- &"at &i balance weights to the hub. T"f-bladt should Ih at. .

SOLID ALUMINUM

fi,,Az balance to obutir, cruise flight angle darkia thisflight. the smoothest operetiao in

If additional balatwing is rquimd. it may be performed on the air'crraft dynarni.'ly by nameS systematic trial w¢,;tt methb or speial instrj-

'Er-

mcaatin,&L&-hat bepulse syichronizer unbialanceindicating (PSUI) unit or a vibration analyer. Because of differca in bUde augular position loading, and poi4ly in nacelk. system respo, m in ftlgt it may be mwcr.ary to supplement groud •.ymidamic bakadnc.1d 'S of ;k= prp• with infjlht dy-

b.la. nc

SE

SP

75%

STEEL SPAR AND FIBERGLAS SHELL'

.

minted substantial weight rvductius, a unmmaaiM in Fig. 3-47. Advaned *apoitie blad con•,udrticn

-- 74

H

•.••-•-n..

The continuing d&tvapat of new a -riala and coLstruction techiWqa, far propelle'i Wdes has perhas the potential for cvm mome isproMvin t. "Thepat qgrapks that fow deal with variou kirn. of Made maltiWs. Adtiomul information i,' containd in ANC.99.

AN

-d

ZZ-

T ITANIUM SPAR AND F IB.RGL&S SHELL F.w

$.47

.

assalde w-4 Wc*ittae~sa

-dM

-

A

S

Iean is %144in satVtCtoryV thniques for fabricating individual cowmponena from the taps. Brazing and diffusion bonding &,c two niroccase with gret potential that peoiaw the ctipbilaty of being dewcopc and monomies.1 manul'c-, In~to highl rqprouci turing -ows." Dramatic increases in stwoLth-toweight and sliffocso-vawtibt ratios over sither wat;4 or titanium arm possibie with the new advancedco pouites. When loa5Tht is wwaetially unidirectiontal with Amreinforcing fibers and the matrix material is o hiighly stressed, an epoxy matrix appears preferable because of its lighter %acigt.However, a nasal matrix offers hgher s®&h and stiffness, where noaded. MAlo. the allowable sts&eIL1 and duign roduki of .naterial with an epoxy mutrix noraniay miut W,ZdjuWte downward to ull,3w for moisture abaorpt;io ir. service and for time griadWa modulus dsware Wada continuous Cyclic: stresing.

When spa-gwe1 vnixastii is iiJ"d. with the spa mtn14r Vnd a sturrundas the majvw Ing shell of a 441cict rnatzrsiJ& tltt £2& is pvotacted from erosion and &iiapkcýdaub t',;b-rt heOlt Intermeal surfaces of the "s rm'u be pztscccss kbwever. (Mass cloth ot i n-splhwtv~d pkiw4ir rmatmrs5 can be used for the ac odynznk~ sge3l awlt bc:Wv to the spar with adhesiives. The iaatw and wxcrAial propcities of the shiell can be talored by selected orientatinu of Whe lay-up. Hollow titanium load-carrying spari are being developed, using SAI-4V alloy. Because of the excalhant corrosion resistance of ahaniw alloys in gSc*ral, no corrosion prot*Aiion woaquiitd. AM airfoil euvelops adhuuively boao4.) to the spar jrovidics both erosion and impact dsmaje pwte~tn. Thorophilugs that can dqrzds the lore no pfwLwr.W fa4ut uvragdh ottitaniwn mand be considered in this typt. of conatrwlion.

4

S4&45,C Fbs Maiculi vhMatuitt Riid wuztaha fowu materials with a daxuity Ln the rnaia~citicr (ild n mmf~ta bennodwiady;,,M of S-12 lbý/R cvrmnitly tit used as tk heanI'A wawd ,~43shav # W ct c k-uv nx kVCw b= an caidyd forv twiil Wý .uin. et& Aiwvb bilktA. Thv. rw"3.j n&a*1I1%4

54&441 Cg"a~o Won OUi

4

of the pout -jihrj:&v*L kjyx. Cz."A nadic LaE. ssd nSn kvnudiC.Twhaior~ f4x lay-tpdu~e4 in ttwW aloha, mi;"Az, wfNn.-n Ww thermal coaductivity ALAtii ~~an wtic to ad rercasns ~ ~~vtA~algUratios with adequate WsA4aw-v A~1.~ut~e rrr~ ~~oc~ai~ e4M'iev adlumiou, and satisactors-tc'sabebya poviiC sj~alskcnh ~rcnasha= Of Particular impcwianca in bta4 ~ ~ RtWtSizThShy. ww ~ ~ ~ o toC -w IVAq, ek-, "i eUSaxt&4 1S7 LU* winsale is Elip. ability of she. foams to beratim icatica fabeý"i ,..um of tM bac eke in unite o OMulet caxitw 01r DeeM -9 c*Atnt fitct S~q imU epasty MIPW.6r S4A4A Stvanlta, AdVtjh." and bwoa i4 cM. Jl6WVbitru 6UI) wUwrfl &;,&kki L~ tli . aa aOUs ~~~ ~~devloamwnu; wis .. Tegnatx w44 Fibeurn,, Vttlp ,,

-4

.

.

.

..

Av

.sasafils flat uhwo -~

Thc4ý b 4 % &q&&& t am"s t pWalak e 2AW ~Thuy c~o," bi- ma4uwsao

pksa of fibao"a ccaia~5( trocihl@itu . as ikuq*~ taLitusjs for me&AQ au.Aw TWc~ kv& &at *.auJ advvs P

itet6,

~N44WA "gVen

Vim Ia aftMO I wfl if tug SWIVAM?.. ftmiJQ& C%~rqý lwUia6Jtd(M4w 1ite4.. Mai'A @i~rA

@~ i"a o1wcecii-tw

t4OL~iaAahi hItfi.044 nd IL;tt

th~ba'tsA~y~wLIuOw

htloud&.a

2%~~rwv

WAOWSM4VL

gYUtJMe

1%

rLdi n4 &&IOz Ows be;iQ Wart,

jaiML.JMO am vw-wziz W"t i nile1* etba w tacVOf kiv - *-if Inc adt"Uvtn boqz4 p~~.vtt~ i~ ~~nhuc .i jAs ini im jý#itd edt sr proo is tkb km ca* ghd da4 cv ta4Jk isA t~cso www4imiuknm.a is Atkitnd, V#4a biJra k CtOVya4* CýAht~fr tMtEILA.Lk "144. FAflCA*E 1.1114-1 % tiC-* Mfl2.%&Otvi &QOPa %, Q>'Ar 4g~o btA~chzwp scon ts t zw-ýU aofpw &QM 4b~.a rajej lwt C4~ k~rw gbfUtntt C!. Mj1Q

&%b4pb S43

6us4atl~ k~k aid a%~o uwttaaw m&.iýs bwaAWc ia~~t~~tat1~U

k:.

*,

Iin-place

fsiatlh,- 4lou akJ

'I

f

bk~~~~adtn wiib mcýts k;.ss&L Akbkc

--

N'

k

C4

amp"s of piopelsr design, determination of blade fatigue strength must be basd upon both compr

E MEAN J'

fatigue tests of full-scal blades and vibra-

*hensive

,

-- T-rr_ -n--n1 WPLCIIIENS

12LT

tory blade stresses measured on the aircraft under conditions rcpresentative of service operations. The number of vibratory stress cycles accumulazed In the service life of a propeller is so great that

vib~ratory stiesmes below the endurance limits are

A

neoesary for most operating conditions. For instanc assuming a I P vibratory stm. and a rotational speed of 1S0 rpm, 9 x 10, cycles are accumulated in lOD hr of service. Some moimentary or intermittent operating conditions can occur in which stres amplitudes exceed the endurance limit. Each cycle of such stress uses up som" of the fatigue life of the propeller, and it is necessary to establish oons•rvatively that tht accumulaton of those cycles can be tolerated. Curreat practice is to apply Miner's Rule for cumulative fati•gc damage. Vibratory stress limits must be derived primarily from controlle laboratory tests for fullcwale propelie blades, supplemumted by sopcimen teats. Ground and flight measurement of propeller vibiatory sresseis during aircraft operafion is described in Chapter 8, AMCP 706-203. The instrumentation required, many of the co. ide•ations involved in planning and executing a vibratory stress survvy, apd the interpretation of results are included. Certain of

_ /

RILL SCALE

C...

"--

TO ... C,

-,

CYCLES to CACK

SCALE

DSEt-eCTON DLOG

SCALe

Fipre -

ladgue Strenh Diffemece Dutwee SpWms wad Ful-cale Teasm (mpreunted by the coefficient of dispersion; i.e., the ratio of the standard dcviation in fatigue strength to the mea'n fatigue strength, at a particular number of cycles) is gencrally in the ran'. of 5-10% for specimens, but may be as high as 15-20% for full-4cale component.

Trh6-

a.-

-t .l

.-

I..

..

nn..,.d

A-nA

upon p.evious tests on similar components, a.tiipated service loadings and environments, and related sevice experienwe. The various regions of the blade - tip, mid-blade. airfoil transtion, shank and retention -t midbt be aonider d as to thsir dhanks and

thes subjects are summarized in the paragraphs that

fabrication details, their steady and vibratory

follow, with emphasis on the interpretation of results. An example of the application of Miner's Rule is in-

loadings, and their emvironmental exposIres. The fqufreafnb for fatigue testing of propel blades arC similar to those for rotor blade dicusied in par. 5-4 and in Chapta 7, AMCP 706-203.

eluded by refernce. 01 @ (dw

E. dlmce U

Stmmm

T"U

54&i

Compiderable information about the fatigue u.rwuth of prooller blad materal an be obaiued from specimen tsw. A dtiscusion of types of testing, numb•r of samples, an statistical in•tnpratio of fatige tat dat can be fond in Rd. 57. StL&t2

hilse

Tels

p-m totig can womplamt, but never rqlw buing , of fu-sclet producti composesta. b of the S i fsni-" tioe dileamows and dieoa ia rss sate, fI-..ak comp-Lie.ts tend to have boh loer mesa 1[t'. snh•os sad a ItI

rtAe

scstter diam coetvvtioal laboratory sped-

mwn. A typical ilkmust

m oft

dWfwr m Is sbown in Fq 5-4, wim the mes ftigu etrenth of flwls coaponms at logWr umbom of cycle is about cm-half tha of the spamima. The data scatr

3-76

7-

-

r

----

T,--D

"M -I

H

no.

pedim" Ta 5-8.2.1 Airraft Teats The instrumentation requir4 for masuring blade vibratory streass in flight is described in Chaptr g, AMCP 7M6-203. In gcneral, survey tests are programmed to e-

compass all sigaisicat re"vice oprauing coitions, with adequate allowance for variability and, whm

possible, for rute cbha•e andlgrowth. The resut of a propeller vibratory strem survey customarily ar summarized in plots of vibratory streis agnin the most peminent variable - suc as propele qse, insped povwe or time. The curv usually are

selected to show the highe

stra

sampl of a gtr summary

-k

in th tp, the

mid-blade and th shank rWoa of the bla&. An

is shwn in Fig. 5-

49. A full sciof sudscurves will form the bei fatigu Iklieanaym.

for the

(

Of '

,

lS.C

at certain critical speeds in. r.,dom 8pound windlanvironment, or some instancs of stall flutter. In such

..

-

" \cases,

-

•newvers, "duranco

-o

I"shown

_._

it may be neonehary to avoid specific operating Sreions In order to assure structural Interity, For vonditions, such as takeff and inflight mawhere vibratory stres levels exceed the enlimit, the cumulative fatigue damae must be to result in an acceptable fatigue life for the in Apropeller. Fatigue life determination is discued Chapter 4, AMCP 706-201.

•5-9

ANTITORQUE ROTORS . Gin

S49.1 GENERAL The tail rotor of a single-rotor helicopter is designed to provide thrust for counteracting main

rotor torque at all flight conditions and to provide QM 4.•

Zo

.directional •that,

,1

So0 10

D

M INtNCAT-t

2W

3W C0

variable thrust for control in both the torque and antitorque directions. Yaw control is effected by variation and modulation of the collective pitch setting of the tail rotor. The collective pitch is controlled by control pedals and normally is adjusted so at the deign point hover condition, the pedals arc in a neutral position. For single-rotor helicopters with the advancing bladc on the right, (the conventioniui

AikýPUD V,

IUFULUi, Urhom'

s

.g:

~

u aee Mmlans To detrmine strucural integrity, the measured starus in various regions of the blade - after due Consideration of beckgroUnd data and allowance for

increases tail rotor (positive to the right) thrust, producing a left yaw. Likewise, the right pedal will producc a right yaw, with the tail rotor loing to lower values of thrust ai.d finally into "reverse thrust". The tail rotor design goal is to produce, with minimum power and weight, the thrust necesary to meet the control and antitorqu¢ requirements. Tail rotor requirements must be met without the occurrmce of any undesirable vibration, whirl, or shake

chavae

characterisics.

FW* 4.

-

TY~sa

Sin s

c r

must be compared with appropriate

material strength information. The strength data are -ued from full-scmae and ojieciamsemins. in choosin appropriate struegth data, the kvd of mean Wrw from davp computations is muflfieutly Accuin. For olpeating conditions consideed to be ementially commuon the srns leves mum be below the endurance limit. Normal cruise flight must be tr#Ated as a cotinuous opma•in condition, and, in ose inK-BaMWl atio rmal diab mid dsme dold be

wr~ed simark. A lhauther omelderaa

The tail rotor is designed for the most severe ambi&it conditions iic.udifl :l-sm-icop.c.,-"itia,-, tilude and tamperatumr, and the critical altitude for the engine. The maximum thrust that the tail rotor must provk• without blade stall is that required to counteract main rotor torque., while also providing th specified positive yaw acelation and overcoming tail rotor yroecopic preesmsion eects, in the maximum 4inc.fled crowind. Prmvmo mum be indudo to

counteract distuebmacso so&ase gusts. Coesidraties also m tbe givm to tbe tmlosm due to inatre-o useam om- mai6080M W u wth th 'Aertical ue Ba, main rowor, ad othr pMru Nuuue i is dmemi howhehr myo"Wsam i Kmm of tAhe ilnpar. wbm thdW ~ sunoma of dos pespulr *0K be Eapede bas shw~s tha if th, tai rotor th~a Poorly rMpg"ae ALa. WoMdiMe Whaere thel VWkseQUINUMMes At the W"ila los-upen oMiiM s aftm eablishing that

-h

toy stv -- tqgh ikbe low m owud - nmg become t lo vu eipesble vaiat s in cnrewuenm. Smi moifions mftk In~smoperedo.

(e4.,bowpisMs fyaw •, 100 aiWewisda minlma dim*, ealtitude) Ma ", d wrpirwe" fair foiwad ligh

easly will be ustheLed kk

vc.ss, ths -- 77

tail rotor thrust cap~b;' st;11 be an 'yzed under forward flight conditions, including critical 550 HP,,, -Rb (519 maneuvers. Without such verification. i! may be R),. *[lb (-) necessary to restrict the helicopter operational cnvo. lope or to redesign the tail rotor when deficiencies are discovered during subsequent Rlight test. + l':4' Prior to final definition of the maximum required TI, lb (5-20) tail rotor thrust, consideration should be given to in. lb (-) creased thrust requirements resulting from the inwhere creased crngine power available with engine growth. -masn moment of inertia of helicopter in Refs. 58 through 61 give additional data on tail sdug Spcf-yaw, Military no 8 ft, rotor design. There currently are SpcfX - distance ',rom centerline of main rotor to cations applicable .specificallly to the design of anticenterline of tail rotor, ft torque rotors. -'vaw acceleration, rad/sec' Q.,-, main rotor torque, ft-lb 5-9,2 TYPCIAL ANTITOEQUE ROTORS R-main rotor radius, It Tail rotors in current use employ from two to six (9R) ,- main rot or tip speed, fps * blades. However, there is no basis for limiting the number of blades. The blades arc retained iai a hub i t, ainrotor to.utrequread to providpe reqfire allow collective pitch change rangigis from positive to main rotoretorqon, an oprvdbrqie ngtvanlsTw-lddtail rotors L-aihy re Teti oo hut l The ailroto thustcalculeted in Eq. 5-20 is the net thrust. taking into account all interference losse bladeshv niiulhigs nete thre. r mre due to tho presence of vertical tail fins, flow field, and type, 53 commonly is uWe'to coto higes.apii dynamic effects. The rotor witi be designed to meet flappingth ontro thern magniude. n tal rotrs eurmn on nE.52 ttecii preconc required for banding moment reduction also tetrs can comenste e usdt fr te aeodyamicand cal boveiing temrtrature and altitude or th. critica centrifusal twisting moenste thates theroyai bandet engine. altitude. Normal forward Rlighrt condi arc centifual wisingmomnt hatdries he lad to of secondary importance since the vertical ions fin low collective pitch angles. The blades may be usually is designed to unload the tail rotor at tail cruite. retained in the hub by bearngjs, tension-torsion High-speed autorotations and rolling pullouts are straps, or clastomeric bearings. tociia xetos Thefr lads tilued otor ar muh stffe in In addition to producing the thrus: required. the all modes than those used for main rotors. For inti oo hl edsgndfres fcnrl onrl so aeo rotor shouldbe designed stance, in the torsionall mode, the combination of maniloroot.Th

"

a,,,

This. tierandr thantrhestiffness generalyis frvetors ih tabne

th hgerlaine masadhg (Ref.

59).mal

5-9.3 VAIL ROTOR DESIGN REQUIREMENTS Thetai roor hdlproucethe thrust noccsitry for helicopter yaw cnradthe attru our-

mnsof the main rotor. This thrust must be pro-

that a linear control will he obtained in all flight conditions iner thr tail rotor does not openatein the structural integrity, tail rotor design sAll conside scoasi pirobt~1s. A

onvitonmeet found with tail rotor hdlicoptcrs e.g., in theprsene f atal fninthe wakeoftemi rotor ats ideward velocities, near the ground, and whim the heiotrhsapositive or negative yaw Band pon he pnutmcmdond~os between the tail rotor and the rest of the helicopter (Fig. 5-50), the tail rowo tkrust T,, required is found frm. Eqs. 3-19 an-0wbX and R, umaured infeet&e5.Gaaur

a

K

5-9.4 INSTALLATION CONSIDERATIONS Durinj design of tho tail rotor. consideration shall be given to the actual installation on the vehicle in determbining the taquired thrust. The rotot may be designed to operate with the tail fin downstream (tractor configurcion), or with the fin upstream (pusher configuration). The rotational axis of the tail rotor also may be canted with repect to the fin to obtain a life component from the thrust vector. The locatiou of the rotor will have been selected to provide adequate clearance from the ground and other parts of the helicopter, and with provisions for the safety of ground personnel (see Chapter 13. AMCP 706-201).

"ALI4. Tractor C..iguratim In the tratctor configuration. the fin produces a blocksgjs that causes a thrust loss. Tests (Ref. 58) indicae. that the net thrust available to satisfy helicopter requirements may be estimated from T.

/I_07S

t

lb

7S A~

05-21)

/

T

-thrust

a.--

for control and antitorque, lb

_dk ~ce

by the

by a lag of the tip path plane with rhspect to the control Axis, which produces an equivalent to cyclic feathering. As a result, one side of the disk is loaded more highly than the other; if blade stall is encountered, the additionai precessonal momuent must be produced by the unstaled side. This effectively reduces the thrust capability of the tail rotor. The rotor blade :nust be sized to operatel at lift cotfli. cients below the value for sttall throughout the operating range. The increased loading caused by rotor precession must be provided for in this sizing. The effects of operation in a side wind alsc must be considered on the basis of a uniform variation of thr~ust with pedal position. When rotor-induced velocity approaches sideward velocity, the rotor will en. counter the vortex ring state. This characteristic, shown on Figl. 5-52 (fromn Ref. 58), giives undesirable flying qualities and generally is avoidtd by pilots. It is prcferable that the induced velocity (disk loading) be sufficiently high that the vortex ring state is not approached until sideward v~1ocity exceeds 35 kt. 5-9.4.4 Direction of'Rotatift When the helicopter is in rearward flight near the ground, the characteristics of a tail rotor installation Ufhtc10

total disk area, dimensionless

"5-.4.2 Pashe CoaliguratiO In the pusher installation, the production of thrust creates negative pressures on the fin and tail boom on the side adjacent to the rotor. The integral of these piressures over the affected are produces a for-ce that must be subtracted from the rotor thrust. This force can run as high as 20% of the tail rotor thrust, but can be redu~cedby increasing the ax-adistance between the rotor and the fin. At a distance corresponding to

VV M. W

tam

ýý

a

~Wa a

n.

aaa * .6WIS

WIi'hcihc

'wp" moving

-e

.

25. S/A

0.264 -201 .0 Q

4..VU

-15

used with caution.

UH1CC; S/A =0. 143 1o0~ ZUH-1C U7..

UH-1D

w

AH 1G G-

X_

effects. Therefore, the pusher configuration should be

o6"

produ~c

AHI1G

offect of distance for both tractor and pusher configurations 6a shown in Fig. 5-51, taken from Ref. 58. It is possble to design the pusher configuration with lower lo-sss than occur in the tractor installation. However, because of Dlow blockage, rotor performance is influenced to a greater degree by wind

S-.. peslia

forward

undesirable flying qualities. The ai rotor can encounter a large ground vortex produced by the main rotor, which causes nearly a 20% deciease in tail rotor

.

S/A =0. 136 ___b

Wheki the tail rotor is oparating at a yaw rat.., aI/ moment is required to precesa. the gyroscopic forme. This momeset is a functin of yaw ratc *, tail rotor_ angular velocity g,,andpolar moment of inertiaI1., and must bea produced by acrod 'namic forces applied 90 dog dhed of the direction ol precession ins the case of rotors with Rlapping blades. This is accomplished

__ 0.8

_ _

0.4 TRACTOR

Fl~um, 5-51.

%%

IUHI

0

0.4 PUSHER

0.8

Fin Seperatle. Distaaee/Rotor Radius 5.79

ramr

Ahm

an ad~m"'%W al

-mm o

3.amr~,i

t~-

u

nbotei

iG

f mb of__ -bw

AQ

-wea

4tdPwr

rwa vahsiwg tIN rolvdl trota. o f k trrn0 ise (Rd pow, #4 to the 3 an intnedyi cm imno himn rwkkr prseto ZXIWof the&;ltdM roqsirdc Hroa CJ~rtAiI IE "t3Un mch rolor rduction. v

ca~ mddhisd ta her4oyate, sti

ane I C E0d

i akeiulwsh coatovr1 ail roto

r

to~yat

dvs!;xxt1 :o vowst wich zbe bott4xx bkgk sovig

o

N

forwnd.¶ I.,~~

~

---

j14--

VNORVORTEX

LEFT

'64

AIGrI

2D 30 DIS& LOAD1NG T/A. psf

5-.SEml~E~tEq10 Possible diastie rdwsuiortk

tail rotor thratt 4im

40

SumM to tlcingiR3sS oa byt'ac tai r-jnor of h~ot exknatgae

(Jhith rwvihxat r~t~d air deoaity) should 1w atos~da .

nd by subvr'lqnot

tesing.

IMm"

I VG PAR t duo i u a alrtrwl pbykO:1O5 5-3TAIL Thecfa pt4~rhssi1 oto "nl 6140 ~ the. losc~nl, tip wgjmtL blade twist, sAnd airfoi vxstion. Etmaaa; tbkese csailncs inf)ucaa the iIaJwb wevight, 4i is =owary tW dtnnino "c~ inatoritatotpin ct&r to op~iiaiL the dcsign.

r

.*AS

j

5-a. Tal ar* Pnfnmmme,

Few Sied.

siderasrd (Egi? to the left. Tail rotor wsiciht sod boo'm size inamaae to disk loads decrease. Thus tjhj final selcton of tail rotor ViLk loadng wilt depend upon ai OvCJHrade-omd ff of required jpamr vemus tai rotor dis'uaer.

S-9.5. Tog RwnWT1# ed T1s~ rotos Aft d4M~imd 1t0 operat at tip speeds Of 54-tA1.2 TogCRw U"M L0144g 6W14= fpc. For a g.ven thrus requircmwt, taW ini t?4 5111* rtlir, the powerC required to rotors opv-ratiog &llowtip speeds will nad highe psmatt thrust depmads upon the dUk Ioadi4-, &urd t.Iiditti to obtain the requiroJ operath~g C,. The ibis on the 4iwtsr naoktld. As noe~d in Fig 5-53, leoý tip wo'j akw) hwuc~scs the torque, of dhe drive systra. i1,agc factor botl, increms the overe~l Wiijbhý the pCM,, F@CWEG0 £WCR*C with UJiCSmad &A Io*iur i.e., thrust bOG~iUJg lbjkg, *erw4acts- To of den antitvrqre utyatm. Highmer ti p speeds can result minirAt the powu; rcovirod. kv tdu kmdJOS am in blade itnakymic comnpmssblity losse wrth cot sn...we 1'.*Iwrn, as rmosd iip.-.3itPrwoovdiug ptwr loam, *bh oonsol forovis blad 4,36!mw 0hai ,, disk hwxL-jg be nsffackflily b11% to kIW .4. ercon prokams sand hiowe noine

I

A 88

YOLMt&WISJ Sins

than 35 kt dUiaiin

'

$-

Boris (Y-4. 5-54).

>

4

air

1W

doe

Ya in raw 4Eq.

90

I

Y*ASIR)

/SKma/n

'.'' sowdoad aped wd/an divId bar E". S&, aNudes 5422 ski hM ib blab Me. Pql Lw tiiota pom..msmsd to a qpdfedb muagouuimd is yaw. -"a

-rvd

labvsoerd to osbouud is 3btwht dwaureigs mud t. Wbepawfrwipee.lod diuiska (sad, 11stm paph PS sdumy) of fthtai9 rawn a d. Sals or kow-epeud Aqt sec Tb. twist neqire dffwiecy bseome pumu to nxmama PCysW. t du kmSWea HoWsova, wish WmM An .psk~ai rotor Wfd''Smai blade twist

70_

-u

-

ot. bu m -W -O hWOu7Wei sofwe a VAIL ROTOR TIP WPE.fps Vlgu 54C T1.S VftbS 10 ooRaw Nam

wow

_F

S4j. Nad. Husk er sof ITMe romw eoiiKy & eam Nc q3W* WeP C aMY WAY Nude chord &Wd had. numbber. Usaeuiy, s is boat t0 kip OhN vale Of asdvadurJ blade 00lidf ob-44l betweso 0M and 0.06 fo vrnon of Srui.a. lure said aselaioma. Tlw blad number b en he sany

I.between

)

,(AMC.

The in noldity rmqed, dopands qa tip spaS, disk loading. maxiuma thiust reqund atd the airfid ±toadun The chai. of fte aidaod affect th a-mmoperating CL. this directy intaad* the solidty. The product (ecb) of the reqwnC blade chord c4 Dm5tunes WOlad numt &i can be found froma Eq. 5-22. knowing the maximum opertig tip speed buiopter yaw rple, ran blockWdap ff, and thiugns:quin.ý

ncut;

6K c~ C.(B)'1wihonly CAC~p-

UA ui~+ ($22 38A ~ tvhcc 1 3BR~ b - number of tail rotor blades B - blade tip los fator, dimncnsionlcs c-e.*T&;'ive '!d ord, ft oissrtA pe1 tWaic, AUS-fl' - pola! rn~;of K m rutio ofn!ty*]teAdwqgthrug to " taii Lrito thrust. 6imensonlc A - tail totor radius, ft To -'tail rotor thrust to oompeoSawc fog main tokw torqv-., Thfor X - disteoce frewct ouerlina of msain rotor to geatedinc of Itol rotor It

/ ,"4 xkTQ "Are +

-+

4

4

4.

"ta

MieAh

whusu-

u80111

imlwsoveral Wction airfoi The coice of theW skeet a dotaMauium -m mW wfostaaacs nakdo hl the .awainvo apacWE lift .OdThimt Of 16 6"W

Ar 66 "AM&'

.Ivk

a

Wa~g p~cluig mow_-- charactensicL Many WJa rates we ACA 01012 and 0013 airoi atin whech h~av a pkchimg nmend of numtatialiy zero cv wr the usualhoratigw. naaeU tp =av and wgjht .f tail rotor are diecjy tpadmi~ upon The mauimum operatiftS lift Coadfkjd=; she nuv or cambere airfoil is bein cm.66=t The amo important chuaractaic of two camubued section cocukidee xuitable f£g tasil rotors ame skein MTable 5-S along with the ChbiAbCtUU&iIWi ofw gr veUC NACA cci:_ for cmsran The ?4ACA 23o1 airfoil is typkiJc! a clam of aim.toil sections wLeW camber is puninai

over !hM for-

war ation fthe airoil. Compared vith the syanmetricol akirfriil U'Ie aanUSa in an improved CmaWx9 a.~'s nrWin pitcbing mornnt coefficient Cm. JCs ACA 6M,412 is typice! of the taminnam-flow urlt(MACA 63.6", 65, sairi 66 strie) ralvsc to jJ -tWvi low values of mrinimum section dra coefficient C,,, by inuintt-mnoe of laminar flow over mutW. J thair MI:f8O;'; aiocs with pldgb-upcd lift and drvag charsrTeisticu flUe to th", fact that the cr-ýcrm it dis'Mbuts6 over thiw cv&ri chord, the pitchino&marinAv czmfficicnit isqcfesr, n-shown In Tabk 5-5. T.; vigrtw at C4,, am shown for cowmarigvs Ona10i ly. VWOW 'sa. . 4~aMig must include Wsrei opri';jng vsiu of RtiC Lid Reyno!ds n;A.bf-Aamrilua leadig 04t~ m~iougt . Fa7itu-x Iq qtry nom-oeq otrtcctiots tn sldail actioo 4 wihý .tul iu deficiet IU rotor capahhtý

r

;4W

-t;

a-21

2

by suitable deap mWd balacin of she roto "-.7 STRLJCTU1JAL CONSIDERATIONS Mlaul. aqatave (acsedowa) pitchin moumat can b. 54.7. S.Swul Druaiss Niluad. For campos by usin a avseisly stiff blade Withpt'00. nd y baancng te ~The plawnm:, of the ustual frqumncims .1 the taZ: with pndymmme sems (Frg 5-5.ing mAy blsewemle aft~ofrtri eattmne&J INICOCIS in insuring5 the MtiUCutdwIsaw srodsaw~ (ig-5-35 itmay emategrty of Ohe syifm. The dials of main ro,ýar 1 so th Ca.mposm of slwingS forclasa;s excitation must be considered. For ia. swodynamuic tbtaqpve arodmimi Onbedi dmAspu toontiue,two, %u'. a sixz rota ortu mayla she tail pitching momts. Howeve. when She CG asmve standit %e winsixg aio, corrpmW~- s ato ineto th aft Ike centrifugal ente~ring moment (tennus raicket 'Z4k At ra. Ae e.Atm-W mfeunyi efa) maybe inceasd and the 4mAuited nie. in sydlit also my be found. Asiaryofti smaui, moiest may 1W occu. Aother method OF sion sources is ashown in Table 5-6 (from Rde. 59). emo isby ug bom.In d pichm ham~g Guideline for ;Aacemnent. oft51wtal rotam natssal boo. haadl that pisthinbostmays by ingaor firequencies have be=n denelcped (Ref. 59) and are.reio- tookaio bWO my hOn ~ti SicopemI SooAtai datrin respnse vihaio modesn 15 Hz~. Sold becnidrd S-*A TAIL ROTOR PERFORMANCE 2. Thc na.&nal frmwicnce. of the tail rotor htotdG0 wku an bw hem rotor Once she sIlIW not be coincident with, nor Sn eo.w ptouiiity ?to, anyShn performusce can be eanimated for the zer velodexciting farce firriqtwiecie for steady-state optrating ty coadition by usin ate sme promdciuus as frw a with the excjta.. coifoa inc hovering main rowo (tee par. 3-2, AMCP 706-201). -

tion sources shows iri TaW.s 3-4-should be swa:idaS for at kni th form two modes.

S

* TABU:L-1 Additional daisof the placement &tthcs.' natural AERODYNAMIC CHA.RACWIRISTIS OF froquiencics m-ýiy be found in Rdf 59. SEVERAL AlIRFOIL S*.CTIONS, SUITABLE FkORruin ldn TA~kROTfl ADI2IYh inc"Auizon of' the firsa four lowest frequency modaý (au1-of-p!kanc and inplanc) is sufficient to etatady-stata rotor behavior. The aerody-

AA111Oil.

CI'4**reprtwait 1 58 0 1 78 -0015 1 67 -0071

_____

NACA 0012 NACA 2J)12 NAL(A 64 412

natnic blade loAds may be calculated by classical techniques,

I

MuCOMPE NOSE

with the local blade segment aerody-

narnic coefficient defined as a fraction of Mach numbe and angle of attack fer the c;Iuserating condition (see par. 3.2, AMCP 706-201). The efkcL. of induced velocity, Prra1ri rotat, velocity, fin intfr-

--

and elastic flct-i~sck velocity. as appropriae. must be included to obtain tist proper overall Ic.rence,

s.ATING U *AWENT

loads. TABLE 5-6 SUMMARY OF TAIL ROTOR _XIT

T~~~_

COMPONENJT

a

OF CF

[IO

____C

F8EOUENCIES

TSOURCE

NEAIEAIVSOTWK#Y.

IPITCHING

AEFEJDYNAMIC

/O&TSTEADY

UNBALANCE.

flb¶ilr

OUT-O)F-TRACK STATE 5XEDSYL-TEM

fi"

INPLANE lt

.4;EXCITATIN S-3. Com rtlemi for Negfuth Pucibchn mtammal With Cadvj Aq-e Aft ilal (C

IFIANSIENT FIXED SYSTEMINI

INPLANE

ibiW,

EX C;TATION

'Xefrt

BLADE MODE

I~

.,

OUI-OF-PLAN'E IP ~ N

OUT-OF-PLANE MIi H10 AE)ECTION ________

INPtAN* 1'x-o.PLN

4k

S41.73 Emb Sfiuua, Assi.el When blade sections. stiffnes, and man dlistributiavg have been selaqed and the externally distri-suted Iceds calculated, the neat task is to establish the

integrity of the blade. The internal strain

qmacy or tbs Loom. in torsi. was omiplid to the ch-ag in rotw thrum due to lhteral vulociy (6flet) was eleoauuaamd. The so that negative dqmpi c'hi'p ta tail rowo thrust due to the amotio is pro"60441ona tohashfila1ppiNg b &andlths pMb.*-fiw COUPhmq tan '1- Item WOfaed is this ma tat" aMW Live k was the best way to damp the sysuma. This type of problem can be avoided only through c"~efWl canskeato ofhe dhelicopser. cmiao Ucm

There ar. tkvo gencasiateigories of design loadin wonlitions c -c;sidercd in bin-kdesid-p. ultimate comditioms and !fligus coditions.'Ibe blade mus have an ultimate sirenth "4% greater tin the highest pea load santicpaWe durnng the lifetime -3( he sysiem. The bklr~e also rnuat have fatigue satength snjfflcient tc pre~vent a failurt 1= La alacmnatinj loads, I~~~Uatpence has shown that fatigue 'salnly is the more uaiticul design condition. Inma analysis a spectrum or atigue loads and tie:, expecte6nmrc focurn.*i the lifemtim oftkLet~c~' sdvlpd h sdrvifo !.tr;ir l nnucl tFL bulair'cý, climbs, T.%*s,;,' lostr. gadl ground-airgrotril c,,cla (sz.ý Chapter7 4. A MCP 706-20' and Chapi;' Ahife ;)-W203). The fvý, , stxcpnlh of the cuy.;paw,nýt ucencraily it

54-7.75 Fh~ase d DhwssMEu Tail rotors arm no scaled-down main rmuors Thur.. fore, fl'atter and divergence problems gasally an not as severe as for the main woanes. Usualy. the tai rotor Wlades arm much stiffe then those us main rotors due to fth opeasting eaviroomeat. Also, the relative inertia. as exprissed by Lack number, is of the order of 2.5 times that of' temiror.Fnly, aspect ratios of tai rotor WilJes are azuch lower. Although these fmactrs reduce the tendency toward flutter addivergence, problems with tail rotors have occurred, so the proper combination of 3 mad pitch linkASfiffnee Watut be iV,,' 11W InMeaddeam for the main rotor (par. 3-4) and the data in Ref. 59 mnay be used to determine the design details necessary to eliminate this problem

structural

(streWl distribution among the elements of the blade is wemputed: the resultant stres Levtls are compared with allowable iewels th~at test anid/or experience have shown will preclude failure during the life of the sys-

tr9A

cr

j

I ~~,

4,

.-

given at, i_1 '>tt 4 a S-NI ý-rivn dcneuive from test date. The rngnitudt ai~d shape of the t~rsrve art a function of material, stirs v ilichtaui'sW1, en Amonmental conditions, and magnitudt of conuurrent srejadl inrcss. The influence of steady stress on ftiti4,,;m strengt m.a y tý& considerrblc and m'jst be coansiderd in fatigiuw analysis. In tail rotor bK~ts, the eleady &za generally is equal to the altcrrt.narj sMtrcc, a t. reduces the allowable alternating strews by aprroxiniately 20%, depending on tix. mat~ia. As with main rotors, tail -otor blader dt'ign i! ""iterative process of developing a blade with adcqvcer aerodynamic, physical, and dynamic chaxactuýs!ics, and fatigue integrity. If, at any stape in the dra3ign process, an inadequacy is de'aced. the design process begins once again uintil all required chatacterustics are achieved. 5-9.7A4 Auroeastfizlty Problems due to acroelasticity have been en. couaterecl in the installation of tail rotors. Thesn problems have -esulted in undesirable flying qualities at bipb speeds. An example of susch a problem, termed "tail wagging", was encounterted on an expenimental helicopter. The rotor was mounted on a fin that projected atove the tail boom. A natural fre-

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"531

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flawy #S4f~fxclWd MaOil"uA AW Hdkispur Des W&6 kikgod Mos NACA I 31014, Wawbdiagii DC. Felbrsuy 1%7. 23. WI~laa K. Ocmewar.r ft~ysm&Iaup 4r ildiwpzar QeedW". Vbnaes. Prosl a pVO of th 13th A4n94al National MWAinG of the Aseeka HficepuW Sociaty, Wuhlwgotr, VC. May 3157. 24. Onepa W. Drkam, The Medeuiks IrasaWty ad EasY Aspesa %f AO&O" cw Nhladgd

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A

It~1.

t5.

36.

0t.

lB.

19.

20.

21.

J.

MI. LOOtOU~,

.4

K.

W.

IRVIVC),

USC9

0ý

Om.

L.

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MX'ma

%J.M%.um nnutaa

aaq

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K

35. K. 0. tAct and F. B. Gufafa.. (bAnf4 LaNot".o( LawjfAdwd Slattay De~jv iai4a ltiakeper RA~ in Fumed #~I&. NACA TH 2309, Wasahiagou DC, bLeah 195l. 36. A. Omaaow, Eymuams and Precalns fer

50. J. C. Houboft and G. Owoo.\ tW~man Eqssfl,awe fmuc. Chw~i:sma of Medsiaf 4 %wd Aw&V. a4d Train f Twined Naulfn basar Dliet. NACA TN 1346, Waebiqom. DC, Februar 195-F. Nannaiy CalwingftO Awre~pod (lar. S1. J. C. Ho.'h* Caspled 8adtq ad Twatae acunitaedcs #1 ft~r Rena. NACA TN 3747. Vtfewmuns Vj flwaV Rimahq DSwks w~for Waebhqiou DC, Otober 1956. Arbflwy Lna IAS Preprins Nc. 539, S7. Runul f.. luiquis, hddctzps Gun ktap w Janufty 1333. lWAdsdl Umsawdy Aen*%ade SUN EOfecu. 52. J. It. De~c and M. J. Turner, ropewPe YR 72-M, US Amy Air Mobility Retch anid lutuC, iou-san or the Actomaaical Scimi Developmnent Laboraloim. Eupia Dsrntuw;A. Ilk No. 6 (June 1949). Forn Eutsi, VA, May 1973. 53. Ii, A 1heertimd 4ppvacr to a Seed, cf .¶aE 38. John C. Kelngn and David L. Kidd, A Study 'if Flwer h.ob4ma Cotmell Ordajut Sboaoo of Tmvbdt-Poemvd aHmpelpr £4 flve Svaim lmnAAcronwtis: Eqinueriqg. AFO6it5 599 Juuu hdlly, Poomefixw of the I.t16N Anne &. Nat6ona 1961. hissing of the America sl~icopvcc So(ciety, 54. F.a. Cauta nd C. Niebach. Pat WA 4aMeIA, Washington, DC. April 1953. Imitabidoy at lila Fensuwr £puedr. Vdh. Ifl 39. SAE ARP-704, Heicopter Esgle/Reiwr Sysemx Smil Rlaaer. TR tý4-IC. US Anujt Aviation Ceeupuadibty. Society Of Automiotive Eaesrs, Ha4auiiiI Laoao~ Fort Evlik, VA, Iý'ruauy NY. Jane 191%2. 1969 40. Josph L. Peakowtki, Aaaaowagic Contirolof Frre is5. J. E. IM~er, The Sifta..' of Venom Poranwees. Turbine Engines, ASMdE Winter Annual Mccetin. lnduatq A~ich Nueesder,. en Puope &vDod 4 1. J. Shapilro, Pdflncles 9f Heilcopser, McGraw-Hill Book Company. luc., NY. I1W* 42. AFDMA Simaidards, Soclion No. 11. Aetehogs of

'IC#~IFN 3357, Waahi..gon, DC, 1955. 3* S. G. B~, 'Popui-er Ftialncing Probk~hi&s, SAE Jokircal, 53, No. I I (November E945). Ewiluawdq Load Rwiotjs futw RAtes Jkmrý 57. STPAý i A, A Gaakk for Fcuigue Testing aadi Ac AniFitinM igMaaofaaatrcrs AssociaS~wisekg! Analyxis of Fat/je Dta1c. Second Si tion In., Y.lion, American Scicty for Testing Matersul, 431. EetginteMVi M~ammaa SF-9O. Rolki Bering Corn1963. pony of Ameiica, West Trenton, NJ. 58. .t. Rt. Lyiin, ct a;.. "4Tail Rotor !)esi1 0, part I44. AFUMA Standards. Section No. '4, Maethods of Aroyaic" unlofhemrcnHlEwe/eec:kmiz" Rthic Bal 14una eans.An Loa Ameica Friction Loadrwig Maafacturc ~ A~inss cAnt:copter Society, 13, No. 4 (October 1971),. FricionBeaiuSMarufaturat sso~atoo, 59. P. W. R&.ke, et ag', 'Tail Rotor Designi, Part H2 NY. Dnmc,'ournal ofthe 45. P. Rt. Payne, Helicopnkr Dynamics an eiyAmerifcan ~~~- Struaursid Helicopter Society, IS, No. 4 (Octobecr nartsacs, Sir Isaac itMan and Sons, Ltd., LonwDnmco don, 1959, pp. 361-36-7. 46. SAE ARP-926, Aerospce Recommended Prac. 60. F. Robinson, "Cmrront rThnaS in Tail Rotowr lice De-sign Analysis PrceaAerrfo& adw Mode, Design', Journal of the American Helicopter Effects and Criticality Analysts jFMECA ), SocieS~~,3.N.4(ctb~ 7) ty of Automotive Enguneers NY. 6i. R. J. Houston andi C. Ei. K.- Mowr~s, "A Note on a Phenomenon Afl-cting Helicopter DirnctionaI 47. R. E. Pctersen, Ssresm Ccncernradion lesigr Fac. Control irt Rcirward Flin&t", Jo'anal of Amertors, John Wiley andi Sons, NY, ubao Hclkitipter Society 15, No. 4 fC-ýc~to-a 197e') 48. CAM 6, Rotorcraft A irwoothinecrs Nornoi Cote v2. W'. Weisusr and G. Kohler, ToM ioior Desige gory. Appendix A. Main Rotor Life I~ktermiwnnGuidc, Tk 13-99. US Army Air Mobility lion, Frderal Aviation Agency,41ashington. DC, kcscarch and Decvoprnicie Laborztorica, Revised January 13. 1%62. Januasy 1974. 4.J. P Den Ilartog, Mcs*ewcaf Vibratbm. Tirzd 63. N. N. B:fr 4 , "Rtzults of a Ta" Rotor fJ.0rcctionEditioi, Mrflraw-1i! lBoc-k Co., MV4, pp. 331. of4-RCUto,'tts:t". Journal of the American Hcli' owtc, Ite', K ... 2 (April192

5485

AAs" cAarr~a 6'-

FLIGHT CONTROL SUBSYSTEMt

7

64UTOFSVM3OLS U -. roIZ tsaapi

duritaivo of rolling momeo wade rqs to rai d"a ft4b/rad-c 4 dihedra ustability; duivatve OF realiai meWenft wiih rweplI to yaw angle. ft4lb/ia 1, - roll anwue saditivip, dmntiiutxw of moant MWa Vsqua to Conro .q~, ft-lb/ia.

aa aI li n'-'

mI"dt d

vh

w

erlniir sim,%"hstrm-n

t liw aimpty theiis S ,guz, sheikihas bawia stttuibbly , M W21 intomarunass cofaw Mated doptu SMCPt do -Ill into S k ceqon awulIo that mussthumam folylow quift ht[lo ie b lafluamo Of hudicogita Mobiity reqsdrwenss Visas

a

ggwt uAI h razwrs0 Sn

VA

~

mit.1 h

r.Ik/u

et.

No-yawast&Kbty; dsvaevccfyaswimgwnonmn N4 with rejajsct to yaw an&P, ft-b/rad N& -N YAW cn1trot asaaaasvitY; dMexnve Of Yawtug ntortueatwith arepec to control tDf4 ft~b/n a -number of pitch lPus. dc~ncnionless suj:nrnber of nonrmdundanz c'oapcocnta i 3 having the. failulio rate: F, ', a numrber af r5uvidsnt womponcmis having th.ý fuikerv rate P, a~ -number of norzrtundant cost piyAudnts

4-Ij. Poleof Dhpsbre Ty; kelly, the prelimnmary 4meign msafts insa drbanilion of the flight controls and a first e.W"Lrs of stabditlt augmcutauo:-i subsystem dwaaralristc teat arm bdcived to be sufficient to permit cortspbaacc of the licopter with the stebhisiy and earnzOl .ipcciFcto.Thcptjainminr dcakg dali. indu&4-con. trc' kinenaitics as firnacs on rotor blade or aaadynamic control sbarfiicc trawl, gciecrai anraregivnt of controlas, and mtdcbaaicsl &a-tx~s Tkc PAV' limia try dcsgnKvL% as a base Point fromn which

.a.atrc iolaumse jrat 2 aLnuiarc. at redundant comnponents

dcýAtvn akealswul arc proposcd for thc purjposc of amtpyovi'g ayste-in capsbilizy, relkitbilizy. m~aintainability, arnd cost. Thwu design alternatives then arm sub-

P W-q

o &

A -CR of oftaa No mcp itch dmpng duiuir ofsi rag-eft 6-1.1 DOEON MEtHO ae qse~d stabliy' duivativa of pisthieS hwcTn= aidrtrn .amrfacsur - &afte a IM ayes Hant with Muuze to crwanl vdocKaty; intqlrstoa of arirfmaze watrols, and uambiity amgft-lb/?Ps matatio ssbisyama - should ooudaac iterative ae jWiAt saabdity; doaiVa~im of paitkz% nka' and/or competitive trade-off studies. Thaci sadfis AV~ ampa viaý tpers 0o pitci ansit, $A-b/md will evaluate the purfoimace. cost. safey. roiabwilaty, - ladA coaziol auiatv-Ay: dcnvative of and maintrawnc charecteriatics of owe or sevarul pss"u* W~tib il L nIVec t onto contw @systems as they relae to the mission maquime- amagft-b/in. MM Mi sthtPRhabilty qmWatioa. refute i, -yam 'Sampiag dcrivativic of yawring H50to within this chapter. S

K

6-1 GENERAL

*

Z,

havirig

the failure raw~P2 fcilure rawe of uajlrqgatc of c~omponents, hr'nea~t. resrp vertica dtamping; derivative of verlica, force with rerpet to vertical velocity,

Z& vcst~ical control sensitivity; dcrivativc of %-erti"Ia forcte wih rcspect to contra;4 input, 0 - xjiafizWrwaton aout ý.yaxnýd

a -kocd, rotatoracl radise

totetd to A Uad esAiifgff trade-off Mtud) Cos maniaTypical considrratioais to' 1w. rivvi.wtd arc: 1.ywii The level of helicopter stability rcquirc-d. wvith or without augmcntation 2. The paramet'.r& that should be controlled, and 3. The automate4 ta~.s or tutopilot, (plot relief) functions that should !beprovided init

swell as toexrnldsubcs

S. Th(.use

of

snl,,da-

rmlihn.

- qZMP Imtrwfr40ac.redisMaccd systems; augmenetation avý'nuavr locaeo~i; in)l00 (and-I wetheraumu' w Corfroo ~mlcd/amc Wtorparalle to the pilot's6 iit puts

V

4eu

AMN 'LCsMW qau AgiaC WkWUe 90 eMWUWga frelsL A oflason cm a oaid a Vapi wtoo rom is p1- issAd MLOC-SI.

aquas10 6m83 SW

.-

rMW t oth LA1M

smaN. AND

a -

lU

-e Mk- de-w%

n

SS

04bi dmmigMa, hqrne"

A.1

MiL-HAI i imasdims Isw wbiA do sablty o do lbs huluer a" be wa mesd. lbe fcwqans addkitlud passders to be Ac~e hor em* awsud lms. qaphbmiis to.waasud MiSim t. Co baits mn lbs wi4l. keefL Mmd InWO ONOaAieS dka5 o aShir .MOWgh musefif in eMpurlee with aG&e pusMa. (CO linis. we.a (1110COi..OalPs week htr mosmmtsied bel-.

-%

2.Allowable rotor WeS variationa

3. Fmatral load wdlgpusio. " ")&Ys ias~ se dependw Won the- fa~ovs; aim, &he heasieshsomd Is ovahiasac throaghous

"aP MnadON of wdis do eaht y hbdnpew tees qma mn ph-m MqI aissali MW a wa.n~ be-.les mise Onimmbetm am pnaga fts Wediq. Uses MW P e dugrma twa f lki no"id ai -m bs qadmu as vont 14 e n.'I COrW"bee, ioke IkAa sAu8=4"" p'aw - evalma& of smu/mA*M.e inosfi prsar and uelabef of lbs Ayine qmAm pa@ fttw S~eOf 111s pcssyM. Kt 1118 daMS. 0@u6soMS ie tima for NpsWaaim of "i"Mus so amy gmenus problem liadicusd by simletias. A msioita. cupa4i of evaiseiia many avsa of lha cors ai etmisd i to te 1.

W-. Compliacs wit sabiky qecifmcstons A*oud be evvualneas and all the hates weighed a to

6-2

stability ckdarcteruans of the baskc, ~aaupmcnacd airframe should be essablishd. The helicopter sysvem i's liky to underg swany thanges in its anionic controls but only limted airframe d~.p duna *blf cce

OWatzd UdJin"e thý Pio'a A&iY to cO"N" dbe vaceadhimpta ofis hursc repnses to external daujten Of myal.Aw~ u rapidiy with thes amt the magninadec or tdo s95 which a steady attitude or trim is ujaiawm, ald t1K exis16nc Of any cecil~lationh or lightly damped to. Engineer work to quanstify the rtirg of quialities and dynamnic stability so that, they er t~e t ro vidruitur.: vehicA wpih afrnsmk stability and to makc predictions or comparions for existing helicoptesb. The latter includez prediction with sufficient accuracy to Supplement fliht ttsting activities and to expand the enajncee's uiosstanding of safety-$-flgbt topics.. , 6-2.1 CRITERIA AND METHO OF; ANALV'ha.S The following list presents the isicre signiricWii, topics relating to helicopter handling qualities: I. Contnal poweti, sensitivity, and intcruis couplin; 2. Inherent or augmenitod static stability and damping 3. Characteristic- ro,,ts 4. Type of automalat control sysW~a and vsntablas '*;oatol~d 5. Fatsz feel 6. htgnitude of respose 7. lWAswv of cmoaalc uampoaaat uou stAbilty.

TOLSactions. 4-12 AALnCA

6-LIANAYFICL TOLShandling

There ame two basic forms of matChmatcal reprcscts~on ssesin or veicl trm ad sabiitr bWialt1-pcfiurbatiofl equations and total-formcqut-' tions. Tie typical skazii-petnusbation equations noted in par. 6-2. A MOCP 'I4i-20 1, expedite the atecalinent of stability at one flight couditkiv, and can incorporate nonlinear control loops -tcadiay. These cauations also are adeptable to paremnctric studies wsing analog or digits! computerm. Total-force equations con.plctcly de~scribe tkt absolutt forces acing uponi the bdlicuptc. Howeve, tisty require & rather l&sIc codapiefielit or either anaog Or idita computCr 'equipm;Au for thei; aolu"sIo.i7s typte of solvtiof. is nwussary wheft invesigaiq larige variations in fli~gh. conditions ýic., speed, attitwfr ard lazgu "h sagA. Windj tutnne tesint, shou4 L..- conoufta. it] crier to reine the mathematal woe.ý. Insatnatioccoo_ comnin aodYasMic cisaurwactrm o, Se faqW

STADIJTY SFI1F ,

di

ATUON

ve F

MIL-4lM rfllw d ois kabs gadidiasý 6Kr w aabhahaqol buakk qlamaeliusal Cnyasaip. wis dos spirmmita, bow~r. dos W Imm~ -Sauil)P pee ~doci dom dymamc Nabiliy beaoh .m~u le in ebut -mh wW46 Am -s -tam na ,-

root becomes swaM! atas

R. dlsubssocaccx as ain ?qxtrio-i reapvnsc. with tilt bot co;nsia-i1 set primarily by tlic amount of roill m; As The hi,;hez the damping, tab showie~r the

dpaeun.i:~freqefi).

Mixt cons .ar. Tile reqJartsncnt for mininiman datmpn4 is sct by MIL-ti-50l, e.g for a typical, light ($Yl)Vehicle unicta 1FR ~oiiditions, the require-ica.ii~a C 3-se. tume conutawat. For cargo hell-

T~s.ac R-Aqaed Damnping Ma2t.2 The mA5pagb ji ltufan i for citner VFR or ILk

ir. Lý xh of I sec. Test results Iron. a nap-of-the-earth (11,41 t),.cai Of A wt-por. CArtconnaisseraice heli-

phatioiiaime.imataw=. Thi adnipngqrakV

Wee a"GWL ;A oliQVUA;

0ptz:r. the IUS 1peuficatiog allows time constants

-t

auu I

0 2a.

bts rollcniEnLfrmC3t

nd uppllletal ataortro- 4*stro rexmware

1. maial3,wel-do~wý eouar les ikey o *tk inRes. ad 3 Te vluý o rlldamping

F4

fl,

12

0.8

STABILITY BOUNDARY VFR-MIL-H-5501 SATISFACTORY THIS SIDE

t

4: >-0.

0.7

1.4

0.6

-1.6

0.5

2.0

~j-0.6

3I

RADIUS EQUALS UNDAMPED NATURAL FREQUENCY

z 0.0.4 -2.5

0.3 *

STBE Cos-'4 art

O

tA

HOVER

0.2

4.0 ,5.0

0.1

__10

pi06lkt

N 12D kt

-"3

-2

= v"co~n

".

+1•

REAL PART OF ROOT

I

STABLE STIM

d20

TOHAL AMLITDEsec TIMETO ALFAMPLTUD, Figur 6-2.

sc

10-.0 10.0 c

UN UN~ALE A

DOUBLE AMPLITUDE

Ota~racedestic Root Plot

can be asese more effectively by the roll subsididnce thmc constant. For missions requiring agility, the timid constant should be minimized, ccr,sistent with other factors such as pilot accelaertion environ-

loops needed in order to reduce this time constant. Because of"the obvious requirement to minimize time delays, any lags in roll control azctuation also should he reduced as much as is praticablec.

ment and lateral pilot induced oscillation (PIO). In the twt rfercnct-,ý the low roll time constant ws obtained with a ri$jd-rotor vehicle. However, the same response =n be obtained with a conventional, articulated rotor with stability augmentation. Root and trimient anialymt can indicate the augmenuttion

At speeds above 40 kt, other lateral modes of the helicopter are comparable to those of fixed-wing aircraft. Unfortunately, MIL-H-85O1 presents no VFR stability criteria. Poor, and even unstable, Dutch roll characteristics can occur with MIL-H-850 com-n pliance. P•or Dutch roll stability under conditions

m

-a

MM re r" Wsim N WANdi Oag polwuts fidr v hu-m .ean the pilot. The mpoiemessW

sid 1obe M":.aiaews M- Mam. (WIntaq the aft pyle.) i the principal aces of in-

and DS* rol ased in pur. 3.7.4A mad 3.?7.$ of Rdl. 4 shoeS be ueed inite. Thu an i6iwir io

pom

Saued-sing upedflcsiona, sod am. be ma with rodirsnairy nbilft arefsfao. Amy coWqn if Ow MWW gn44 sassmad .•"awednds (rdeoammwo•y a "ateral p1qcd'° rImNN to - a "qWiral-W mode) •omd as be i " by the pdlo.

for sms-rar voces.

6-2.1.3. Veim of husman Tb. ra of added Abit g eqm ussia uo Un ni 's moft Smd be eiewe thoougallyo m eroot pots.q"Mw jy w* sWiad to down.ism aaynuica pis, soin eoatas air-. so othe flbs conditis Ths iiwms~ata "1i1i3 hh0~e Abba, Snhhy• will help ave tmin aid coo by hadieia rieas of Th mvAagied (imbmra) ailky of th bhecreed suibly an roam wheo uablit too moocolpe aO be oaioodwWfly •o •e wih ziect wb SeM t"W 'Mn S Pwu&Stw WAS U Px plo formats imuoinas to dee amt awonlgendw saad fRgh uewvlape boits, evn shoog I. MSine hebuopsar a airuqass. ateritical groa is planned Thiis a smwey bWarv aaqecusio.m aitusion cam mae w which &aqansic as e wighlts and as bash etr•sm of thOCG roaoa failuen requjir ta

-"

to fly only with she minest

-wd

2. Pause hdoeul'er WsuM aispoeked. as i

lhaw 1.

vehicle abiliy. A oo-Ocliwaem and pwrfcsM ip trade-off exists banwe the a -ef e

but at the upper bom deisted for tLh sinna 3. Helicopte with stability aupagaioe for the

de•s of flighMt eo

conditions -

sys&

ophuii

on ad M fae

Itsms I &d 2.

dre of ishemt nauility provided ky tshe lsic hblcopes Oefignion. In efCting sad a tads-off

-•systemu

6-•2

Typ of CadW

.atdy. the lad of stability required under failure coo-

The types of nubiity sugumetation applicable to a

ditiomw ! be defiwd siance it may be ubjec to variaion with he flht control symem concepcua. For c ,, ie_ oa eL.a•a-

given design vary i the parameters they control and, Iurdoe in how the aid the pilot. Co o.... .

ships between augrmtation sophistiaion and inhen• stability: I. A reliable, multiredundant augmentation ca tolerate a significant degree of inheen~t in-

1. 2. 3. 4.

Rate Attitude and trim Altitude and/or airspeed hold Heading

stability. A vehicle so equipped, however, m,'y not be capable of flight with all augmentLion switched off. 2. A sigle-channel augm•,nit.on system will require inherent vehicle stah.'" ty suitable for cmwpliance with the VFR requi'mements of MIL-H-&SOl under augmentation system failure conditions. 3. A dual-channel system will require an inherent stability level somewhere between the first two cam. A hardover failure of one channel (while operating in the normal dual mode) can produce a vehicle response that causes the remaining operating channel to experience a saturation in opposition to the failure. Such a failure, with inherent airframe instability, results in a rapid divergent vehic!e response, whereas the vehicle response with increased positive inherent stability becomes slower and more easily controllable by the pilot, A method for improving the inherent longitudinai stability of single- or tilt-rotor vehicles is the addition of a horizontal tail; "differential delta three" (rotor blade flap-pitch coupling on the forward rotor)

5. Hover position 6. Special augmentation. Rate controls provide improved damping of all augmented dqrees of freedom (potentially six, including three linear and three angular). This type of augmentation aids the pilot in coping with the shortperiod resnonses. but does not prevent Ionhlterm roll drift and possibly a sluggish or unstable speed hold, even with a stable stick gradient. The addition of attitude loops and trim functions (such as lateral accelerometers and speed-hold loops) aids the pilot by providing long-term trim-speed hold and strong sideslip roll attitude and pitch control. The pilot then must provide only the power setting, altitude, and heading control. Finally, the use of altitude-hold and heading-hold removes the need for pilot input to the controls. Hover-position hold loops, and controls designed to impart stability to cxternal sling loads in forward flight, are examples of special augmentation controls.

can te used with tandem-rotor vehicles. Lateral-

6-2.1.4 Trashet ResponMe

directional stability improvement is afforded by the tail rotor and a vertical tail surface on single- or tilt-

The characteristic root analysis only partially describtc the handling qualities of a helicopter. The mag-

646

mMW

usldlimmbemcs I also mot be desierld.

For- p

casur

impsabs. cn

kdL Tale 6

seaskivity

m uirmess age Sim by MN5L-14501 for &Nbutthe

"eic aaL

Sineaomphmiua"

ýwt

w09

deani of stabil"t aqimsmmaai oao- which may ad to in abs eosaiit. This e(ba m bs overcame by the incprpasion of coNtr ol t hmd-Isad awamnsIL TlS aP uS ol 0 a-s um of ho-pi. mssaic fim Smn

-

mama so steral ditanmce, but also provide dsh de d astivisy to t in addition to apfltmmi. for oliwtudiaal aew r vol sensitvity. MEL-H-US I0pents a evamisso for manever sabiity. It stats that, following a qd sep inpu, the accusatwio and p-b rate a Awk be cwa downward and canvsrgmto no usm than 2ec alter the mtart oflte input. A bort time is deirad for attaining a abaEnd valut of nmeleratbiwospoci lly whse inanevering capabiity is estial to the muia . Anallses can indate methods of maximizingthis per-

-

aye1ser

ft.a

naý"ae

feed-forward and loop closu. '

%Nmwanminnc nfl

Such analysis should

include any significant structural dynamics. For xmmpl elasticity of the -oo mounting cause a major increase in the time required for the response to become concave downward (Ref. 7). "Italso is deirabit to analyze the tranient response to pulse imputs. This response will indicate the helicopter behavior in turbulent air. An asseasmet of the attitude time histories. snd the time interval required to rea•mire the initial trim flight condition with alternate stability augmentation schemes, is ing some form of pilot control reed-forward, the pulse should be inserted downstream cf the augmentation system so as to simulate a turbulence encounter ,c.nce pilot pulse imput into such a system is not equivalent to an atmospheric pulse. In addition to the transient response analyses noted, transients associated with the following maneuvers should be assessed with respect to SAS authority, control margins, and vehicle capability: 1. Jump takeoff 2. Rapid acceleration from hover to maximum level speed 3. Quick stop 4. Autorotation entry and recovery -5. Hovering turns 6. Pedal-fixed turn entries and recoveries ,7. Fixec ollective, constant spend turn entry and recovery, The characteristics of residual, limit-cycle as-

pres-s. a ast of SWcwae limis for

- by a dim

w.

Fig. 6-3 is a gmphical pImssanom of allowa

-*

iatfr naa 7%u WMuriMM 1 ninsit S*moish the bigh-lrequacwy houto ansuola ras sntablibs 26• ifrIsy rane and andisplaces sKablIb the likOW-frequec mi

-

611.9 O Fman e stability and contral requirements qcified our ivod for thUe hliaopa eqamenIt the design ob-

jecaives for the flighot arol ay~stis Therefore, each

potaet flight control systm dmgn muss be inveiated to iumm that ma doe not violau• tho operational rsqummirmsm. ODA&. a tability a& metMAtio

syste

will exhibit sadsactory per-

tonmnae for mmall-perturbaaion maneuvers or disturtanes, but will impai vskid. stability severely duein lag disturbances and atmosphei turbtmlem because of its rate-limiting or maturation aka...

k .. hnm e tt tabish the a operational limits of the flmit control

system through analysis and exp4riment. If the system limits vehicle stability, it must be reconfigured appropriately in order to assure; spedfrcation complianc. Flight conditions for which such studios must be made include high-rate vehick motions, regions where stability may be £cntstive or highly nonlinear, and turbulent air. The influence of flight control system sensor outputs upon vehick stability also must be considrrmd end monitored in order to assure proper design and a pitch angle sensor in a bi-4y axis coordinate system will in',oduce significant inter-axis coupling at large bank angles, because the gyro operates in an earthaxis coordinate system. The effect of atmospheric or self-induced turbulence upon helicopter stability and control requires significant attention. A vehicle that has satisfactory handling qualities in calm air may exhibit large attitude and rate response., including poor speed hold, in gusty air due to poor modal damping or to control loop nonlinearities. Response to atmospheric turbulence shall be evaluated over the entire flight onvelope. MIL-F-8785 provides critejia suitab', for gust analyses in cruising flight. Attention also shall be given to special operating requirements, such us an external cargo hookup, where gusts or wind shifts often increase the demand upon the position stability augmentation system. Many nonhslicopter rotorralft are subject to ata6-7

ANCP 7W~202

bility degradation from self-induced disturbances closn to the ground during hover or low-speed operations. This phenomenon, often termed "skittishncss". is believed to be caused by the rotor wake

failure but the vehicle characteristics must allow safe entry into this condition. During entry, the characteristics of the vehicle shall provide a reasonable pilot reaction time from the point of power failure to initial

reflecting from the ground and rcimpinging upon the fuselage or wings, or being rcingestod as rotor inflow. With tertain combinations of vehicle con-

corcective action (I soc minimum, 2 sec desired), and should permit a pilot of average ability to maintain control with adequate margin. Once stabilized in

figuration and flight operating variables, this re-

autorotation, the vehicle should be capable of mild

coto ytmis fietosprssktiheste desinermus hatsatifacorystablit insre

su-

mentation adcontrol margin levels arc produced. t TABL MAINIJM 6-. APLIUDESOF LIMIT-CYCLE OSCILLATIONS ________

with

________ ________vchiý:les

OSCILLATIONi 414k

NITTC -IF-8786

L,.IIEAR ACCF'.ERAT!ONS

ThIý,J"1

I

85 1,.1-949

M114-876

MIT T-9

AEA

00 VERTCAL-n

proverrents.

5 and 6 describe the difficulties of obtaining a

/

SAA2

/

~

auturotation.

Mit 4 -C7R

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

.-

Interaction of stability augmentation system Magnitude of control trimi change resulting coilective pitch rcoucioao riC4Cb~d~y to C1isiw

ting the average pilot's ability to efltct the autorotation maneuver, and of defining potential im-

'DiSIGAD HLIý_OF IR F%'I"ENIRefs. FO PUE

I

Bufieting due to wing wake

.00

12 9

OSIAIN

II-I

3. Flapping and blade clearance with reapcect to the

4. 5. 6. irom

USALI-54

M!L

fixecd wings

2. Restriction due to blade stress limits

AT[ýNS ANGui AR fAT

-OSCIL[

AXS

falblwaaevlu.including that needed to maintan hdralicand electrical power, during entry into autorotation. Other factors to be reviewed in connection with thiis maneuver are: 1. Margin of control power available to overcome disturbances, especially near zero load factor and for

~\ ~

Figure 6-3. Allowable Pitch Control System Residual Oscillations 6-2.2 AIJTOROTATION ENTRY Most low-disk-loading helicopters have the capa. bility for stabilized autorotation in the event of power

satisfactory time delay in the event of total power failure at an airspeed near 200 kt. With dual engine installations - where the probability of sudden, simultaneous (less than 2 scc) failures is extremely remote - it may be reasonable to consider only singlec-ngirie failures.

6-2.3 SYSTEM FAILUJRES When a stabilitv augmentation system (SAS) is used - whiether it is electronic, fluidic, or mechanical the potential hazard of a hard-over component ~failure exists. MIL-H--8501 requires that the pilot be able to delay a corrective control input for 3 sec without the response exceeding an ar~gtiar rate of ~0 deg/sec or a ±0.5 g change in normal acceleration. The influence of flight conditions and CG position upon the severity of the response to an SAS failure should bc reviewed. Frequently, an aft CG position :oupled with flight operation near the blade or rotor limits is the most critical situation. Stress levels upon recovery ftrom a failure may be an additional factor in the ability to satisfy the failure requirements. Responses to SAS failures can be reduced in magnitude by the following meahods.-

AMCP 706-202 I. Reduction in augmentation system authority 2. Increase in inherent airframe stability 3. Multichannel redundant systems. Selection from among these methods during the design process is done after due consideration of the other flight control system requirements such as per formance (especially the authority needed to meet gust and maneuvering requirements), reliability, maintainability, and cost. Failure-effect studies, which note the consequence of each component failure, should be conducted in an organized manner. These must identify: I. Any failure that cannot be tolerated, such as an oscillation due to loss of feedback 2. Any compromise in control margin 3. Failure causing multiaxis response too difficult for the pilot to control 4. Ability of the pilot to switch out failures 5. Consequence of subsequent failures. As an example of Item 5, after the first failure in a dual system, the remaining system must niect the failure criteria or the flight envelope must be restricted so as to meet the failure requirements. Pilot-in-the-loop simulation is a valuable tool for

pushrods, to both the swashplate and the blade pitch arms. Cyclic pitch input to the blades is the sum of pilot control input and stabilizer bar teetering motion. Viscous dampers, connected from the stabilizer bar to the rotor shaft, control the rate at which the plane of rotation of the bar and rotor follows or lags the tilt of the roter shaft. This lag in tilting stabilizes or damps the helicopter pitch and roll motion. Additional data may be obtained from Ref. 7, and pars. 6-2.4.3.2 and 6-4.2.1 of AMCP 706-201. 6-3.1.2 Hiller Servo Rotor Another early mechanical SAS is the Hiller servo rotor. The two-bladed, universally mounted, underslung rotor has a gyro bar fastened to the hub at right angles to the blades. On each end of the gyro bar is a short paddle blade with airfoil cross section, whose pitch is controlled cyclically by the swashplate. Cyclic pitch imparted to the servo rotor tilts its plane of rotation, resulting in a cyclic pitch input to the main rotor blades. Stabilization results from the lag in the response of' the servo rotor to tilting of the rotor shaft, and the consequent pitch and roll damping due to the lagged response of the main rotor. Additional ur u

ubtainud in zci. a and pars. 6-2.4.3.

.......

U4*

impact upon detail system design.

and 6-4.2.1 of AMCP 706-201.

6-3

STABILITY AUGMENTATION SYSTEMS GESAS's.

Par. 6-2 contains numerous iz-fercnccs to stability augmentation systems (SAS). Owing to the inherently poor stability of a helicopter rotor, satisfactory flying qualities have been achieved in many cases by altering the inherent characteristics artificially Such techniques are celled mechanical stability augmentation. The pilot workload associated with early helicopters was very heavy. The handling qualities requitiments of MIL-I1-8501 have been developed not only to reduce this workload, but also to increase the mission capability of the helicopter. The result, however, is that it is virtually impossible to satisfy thrse requirements without modifying the in. herent characteristics of the helicopter with a rather sophisticated SAS. 6-3.1.1 Bell Stablllzr Bar Perhaps the earliest mechanical SAS is the Bell stabilizer bar. A bar with weights on the ends is mounted pivotally upon the rotor shaft at right angles to the two-bladed, teetering rotor (Fig. 5-7). Mixing levers are connected to the bar, and through

111n1y

6-3.1.3 Mechanical Gyro Refs. 9, 10, and II discuss two applications of intermediate-size (10-15 Ib) gyros in mechianic&l The first, produced by Cessna, is a rete gyro that is connected mechanically in series with the input from the pilot control stick to the control boost actuator. This system acts primarily to damp roll motions. The second, the "Dynagyro" by Dynasciences, is a two-axis, hydraulically driven unit with rotating damping arranged sn as to align the gyro wheel slowly with its mounting reference in the fuselage. Outputs of the gyro, i.e., its pitch and roll displacements relative to its mounting, are fed into hydraulic boost actuators that are connected in series with the pilot's cyclic pitch boost actuators in the respective directions. This design is similar in principle to the Bell stabilizing bar, except that blade pitching moments are prevented from feeding back into the gyro. The Dynasciences SAS also hicludes a hydraulically driven rate gyro mechanically coupild into the hydraulic boost actuator that controls tail rotor collective pitch. This provides yaw damping. No electrical power is required in either system. Ref. 12 describes an all-mtchanical yaw rate gyro for single-rotor hel.copters. The gyro, located at and driven by the tail rotor, tilts about a longitudinal axis in response to a yawing rate of the helicopter, and 6-9

AMCP 706-202 phase to reduce the d,:flection (Ref. 17). This concept has not yet been developed fully.

mechanically changes tail rotor collective. Rudder pedal displacement moves the reference poiat of the gyro centering spring, biasing the system for turns.

6-3.2 CRITERIA FOR SELECTION *S-3.2.1 Augmentation Requirements it is virtually impossible for a helicopter to comply with the handling quality requirements of MIL-H8501 without some type of SAS. Selection of the type of system to be installed requires evaluation of the deficiencies of the unaugmented, or inhereny, characteristics. The evaluation criteria include both the specification requirements and the requirements imposed by the missions assigned to the helicopter. Refs. 18 and 19 discuss the tailoring of helicopter handling qualities to mission requirements exceeding those set forth ip MIL-H-8501. Par. 6-3.1, AMCP 706-201, presents recommendations for contrcl power and damping. High-performance attack and troop support helicopters require high control power in order to achieve the necessary maneuverability. Good damp. ing in rol!, pitch, and yzw also is required in order to prevent the helicopter from being oversensitive and diicult to hoid in a given attitude. Furthermore, helicopters become more divergent at very high speeds, with the result that speed compensation of the stabilization system may be required. By means of simulation studies with alternate helicupter/SAS combinations, it is possible to determine a range of gains for the SAS that will cover the extrcmes of opetational requirements. During flight test of SAS prototypes, adjustment capability can be provided by means of calibrated potentiometers or resistors (decade boxes). Final values for system gains should be baqed upon adjustments made under actual flight conditions duplicating those of the required mission. The test program also will establish whether or not the gains can be constant, or if they must vary with flight speed, gross weight, or any other parameter. For a smail obsevation helicopter, the requirements of MIL-H-8501 generally are adequate, and the simplest mechanical SAS may be sufficient to meet them.

6-3.1.4 Lockheed Coutrol Gyro In later versions of the Lockheed control gyro, a gyro bar, consisting of as many arms as the rotor has blades, is mounted universally upon the rotor shaft above the rotor hub. Pitch links connect each arm to a pitch anm cn the following blade. Push rods also connect each arm to a point directly below on the swashplate. Sp,-ing capsules in the linkage between the swashplate and the control stick enable the pilot to exert a moment upon the swashplate. This moment is proportional to stick displacement, and is transferred to the control gyro, which precesses in the appropriate direction 90 deg of rotor rotation later. The tilt of the gyro results in an input of cyclic pitch to the main rotor blades. Rotor tilt and fuselage tilt follow because of the relatively high flapping natural frequency of the hingeless blades. Additional data about this system may be obtained in Ref. 13 and in per. 62.4.3.2, AMCP 706-201. 6-3.1.5

Electrohydraulic SAS

In order to achieve acceptable handling qualities, many helicopters use electrically driven and sensed rate gyros to measure rates of pitch, roll, and yaw (Ref. 14). These rate signals are amplified, shaped, cross-coupled where appropriate, and fed into lcctrohydraulic servo actuators in series with the conventional control boost actuators. 6-3.1.6 Fluidic and Hydrofluldic SAS The fluidic SAS, which is operated by air or liquid, is analogous to tke electrohydraulic SAS and may be substituted for it (Rcfs. 15 and 16). The fluidic SAS, with specially developed angular rate sensors having no moving parts and with integrated circuits having n3 external plumbing, offers advances in reliability and significant savings in cost and weight. However, it represents an advanced state-of-the-art, and it still may suffer from problems such a3 leakage, temperature sensitivity, and nuii shift of the sensors.

6-3.2.2 Helicopter Size The gcneral category of SAS, mechanical or power-assisted, to be used is dctermined by helicopter size, Only the smallest helicopters can use allmechanical systems, because the rotor feedback forces that the SAS must overcome are corrcspondingly small, In helicopters with power-operated controls (par. 6-4), the SAS need not operate directly upon the rotor but can operate at a much lower force levcl in the control system below the power actuatorr

6-3.1.7 Fsp'lup g Moment Feedback Rigid-rotor helicopters exhibit strong noseup pitching moments with an increase in speed or in upward gust encounters. One method of counteracting this tendency is to sense the pylon bending moment rind to apply cyclic pitch in such a direction as to reduce the moment. If the pylon is flexible, its deflection due to rotor moment can be connected mech-mically into the cyclic pitch loop at the proper 6-10 I

".L•

- W '- ,

,•,-

-,,.

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.

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. .

.:

..• . . ._

: ,•

:.

. .

"

• ••

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:=

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. :

AMCP 705-202 (between the pilot's stick and the actuaton). In practice, if both hydraulic .ad electrical power are available, the SAS gytos arm made as small as possible and their output signals are amplified (electrically and/or hydraulically) to the power level required to provide inputs to the control actuators. Dual or triple electrcal SAS'i can be provided below the final rotor control actuator with less weight than a single mechanical system.

)

-

6-3.2.3 Type of Rtor System It is possibk to use the Bell stabilL-er bar or the Hiller servo rotor with rotor systems having more than two blades (Ref. 20). However, some of the obxcure refinements or kinematic relationships necessary for the success of the system may be overlooked. For example, the orientation of the gimbal pivots on the Bell rotor is critical in order to prevent driving torque from acting about the feathering axis. In the Hiller system, the amplitude and phase of the feedback of blade flapping into the cyclic pitch control of the servo rotor paddles are very important to the effectiveness of the system, The Lockheed control gyro, which is processed by forces applied thiough springa, is applicable only to hingelems rotors or those with an equally high flapjoing natural frequency. In order for a flapping or teetering rotor to exert a momeot upon the fuselage, it would have tc, tilt relative to the rIotr shaft. To sustain this tilt, the control gyro would havc to be tilted by an equal or greater amount, depending upon the b linkage ratio. This control gyro tilt, 90 deg out of phase with the swashplate tilt, would alter the phase of the maximum spring-applied force upon the control gyro, causing it to nutate toward the saahplate tilt, and eventually to line up parallel to the swashplate. In the case of hingeless rotors, with their high control power, the required amplitude and/or duration of control gyro tilt are too small to permit any noticenable gyro prew ione Inigeneral, whenever a rotor-mounted SAS is modified from its origin form, an extensive pro.... grim of developmental and qualification testing is necessary. The internally mounted electronic, hydraulic, or fluidic SAS's, which are more flexible and less dependent upon rotor dynamics, are more adaptable to any rotor system. The electronic SAS gains are adjustable individually in pitch, ro!l, and yaw directions; can W. made variable with airspeed; and can be cro.•-coupled if eesired to compensate for adverse airframe crou-coupling, H emode. 6-3.2.4 Helicopter Coafgraion -, ingle'rotor helicopters can be equipped with any

type of SAS that is compatible with helicopter size ani the type of rotor used. This is because pitch and roll attitudes both arc controlled by cyclic pitch inputs to the main rotor. The yaw SAS, if used, operates by controlling the collective pitch of the tail rotor. Thus, each SAS input to the helicopter control system i3 independent. On the other hand, tandem-rotor helicopters obtrin longitudinal control by use of differential collective pitch of the two rotors. Obviously, any longitudinal SAS will be required to change the thrust of one or both rotors. Rotor-located, mechanical SAS's - such as those that are used by Bell, Hiller, and Lockheed and that affect only cyclic pitch - are not adaptable readily to tandem-rotor helicopters. The necessity for mixinL all controls from the cockpit of a tandem helicopter before they are impressed upon the rotor makes it more straight-forward to introduce SAS control inputs in series with the cockpit controls before mixing. However, in large helicopters that contain many linkagas in the control system, even normal amounts of play in these linkages may detract from SAS performance. Therefore, the SAS output signals for the respective axes should be mixed eicatricaiiy in the same mannvi and ippici tiOrt a" arc the mechanical controls. Then the SAS control inputs may be introdu~xd at the input to the upper rotor control actuators. 6-3,2.b Suppression of Structurl and Rotor Mode Responses, Vibrations, or Gusts Helicopters whose blades have an inplane natural frequency below the rotor speed consequently have high response to horizontal pylon forms at frequenhies of rotor speed plus lag frequency ( wt). - If que.c.t aorotor speedpus agvifrateon either flexible or rigid body, has a natural frequency that is near the aforemnntioned sum or difference, there will be a tendency for annoying, large, transient reGponses to gusts or sudden lateral control motions. In the case of resonan=- at the difference frequency (01 = Wr), selfcited destructive oscillations can occur in the Wir or on the round. In some cases, the SAS roll axis ha& coupled with the r-sonance and aggravated it. Thus, steps rhall be taken either to eliminate SAS res[ onse to the mode or to make use of the SA3 in st ipressing it Unless a epecial design effort is mad,, the total lag of SA, sensors, signal shaping, and actuators at the high frequencies of the transient oscillations is liable to shift phase respons, into a region that caurts divergence rather than attenuation of the It may be necessary to install separate sensors, filtered to respond only to the pertinent fhequuncy and then phase-adjusted so that the final SAS output 6-11

-ý

A'

7ý,'

_____-I....

AMCP 706-202 is at the proper phase. Both the SAS and the entire control system must retpond to this frequency. Ref. 21 discusses the theory that n-per-rev vibrations may be reduced considerably by suitably phased control inputs of the same frequency. This type of vibration suppression requires large amounts of power, and shortens the life of the control system considerably. If such suppression is to be used, the control system and SAS frequency responses must be approximately 15-20 Hz. When dual SAS actuators are inserted in vertical linkage, the mass of the actuators may induce small control motions in response to vertical accelerations. In a specific case involving the collutive pitch lever with friction lock disengaged, the weight of the pilot's arm coupled with the vertical motion of the hellcopter produced a sustained oscillation. Mass balancing of the control linkage and/or the use cf viscous dampers are methods of curing these oscillations. Gust alleviation by means of control inputs responsive to gust-sensing instruments still is undeveloped. Closely allied to gust alleviation is airframe load limitation by control velocity restriction. However, the cos•trol requirements of the two tend to conrlict iocau-. gust aaieviution requirv-x rapid conStrol rcspons.ý. Becaue both programs have as their objective the reduction of airframe loads, the gust alleviation system should perfonm the duties of both. Ultimateiy, the SAS will include the fun~ctions of gust alleviation and load limiting in addition to flyingquality improvement. 33SSRELIABILITY The expr-ssion for the fa.lure rate P, of the aggregate of components in a system, as shown in Ref. 22, is PA = ni PI + / -'\l 1+ n2P2 +(!L2

!"2, hr-' (6-1)

where nr I

rtLumber of nomcdundant ct ;nponents having the failure rate P, -number redundant components having the failureof rate P,

- number of nonredundant components having the failure rate P2 n'2 - number of redundant components having the failure rate P2 A mechanical SAS having fewer than a dozen parts, all of which have a very low failure rate, is u!trUreliable compared with an electro-hydraulic SAS with hundreds of parts. Oa the other hanm, the weight penalty of providing redundancy in critical parts of the electrohydraulic system ii not great. As seen in 6-12 n2

r-~-

Eq. 6-1, the sggregatc failure rate is highly dopendent largely upon the at-iount of redundancy, given that the design insures that failure of one medundant component does not affeW the operation of the other. 6-3.3.1 Safety The designer must be cognizant of the influences of the inherent stability level of the helicoptcr and its SAS performance upon flight safety. Added stability margins can improve safety during night flying, or during limited-visibility situations caused by the presence of dust or snow ,louds. During such operation-;, the provision of improved stability levels allows the oilot to c¢ncentrate lesb upon flying the helicopter and more upon other pilot duties. The flignt control sys'em should be designed to allow the pilot to detect or diagnose a failure, disann the failed system, and effect corrective action. This requirement may involve soine for- i of onlinc statusmonitoriag for the variouL control system elements. Design compliance with the current Military Specifications does not preclude th,- possibility of inadvertent flight operationt with one channel of a dualchannel augmý Izstion sy:ni inopcra.v¢. In this type of failurt, the differenct, in flying qialities is small enough to be undetectable by the pilot. Thus, the pilot may enter a flight condition in which fail ,re of the remain;ng SAS channe! cannot be corr cLed within a Zasonable reaction time. For certait critical situations, a need exiets for automatic control activation. For vxainple, electromtchaniLal SAS links should revert automatically to a mechanical lock if hydraulic pressurr is los. .This eliminates the possibility that a sloppy extensible link will create control difficulties while the pilot is attempting to cut off the failed system. Another exampie involves external cargo-handling or -lowing operations, where it may be necesry for the load to release automatically if the applied moments exceed safe levels of controllability. The rerul s of the failure effect analysis (see par. 6-2.3 of this volume and Chapter 3. AMCP 706-203), including any supportintz piloted simulations, should be reviewed and verified 5y flight test. These re.ults then should be incorporated into flight handbooks in the form of wuning notes or flight restrictions for various failure conditions. 6-3.3.2 SAS Faihnur A discussion of SAS failure madu, limitation of authority, and time delay criteria may be found in par. 6-4.4, AMCP 706 201, and in par. 6-2.3 of thib volume. Failurts of rotor-mounted. gyroscopic SAS's

"

j)~ are not discussed. These systems sAll be designed so as to be at leat as reliable as are the rotorcraft primary flight controls.

6.33.3 F'ah-sife Prlmdples Fail-safe desian, redundancy, and self monitoring priniples also are discussed in par. 6-4.3. AMCP 706-201.

L

6-3.3.4 Battle Damage, Vulner.41llty Steps shal be taken to reduce SAS vulnerability in cases where loss of all stability augmentation would abort a mission Duplication or triplicintion of actuators and hydraulic systems is a valid approach. SAS actuators should be designed so that it is possible to lock tthem in a centered position in the event of loss of hydraulic pressure. Duplication or triplication of hydraulic lines does not reduce vulnere bility unless provision is made for automaticall) cutting off the oil supply to sevcred lines. Levers, bell cranks, and pushrods can be made large in size and of lig'st-gage, low-stressed material in order to ir"-uce vuIneainhlitv tn--Imall arms fire. Critical components not readily duplicated should be grouped and protected with armor (see par. 14-3).

j)

6.3.4 COST' 6-3.4.1 Developrmat Coam The cost of developing a new SAS generally is in proportion to the ad vance in the state-oif-the-art repmoented by the dev~lopment program. A conventonal SAS for a conventional airframe can be obtained from off-the-she'f components, whereas a new concept for a rotor-located SAS, or for sensors based upor new teclinoiogy, may require a large expenditure in order to bring it to production status. The new cont&pt must promise a sufficý-nt increas: in cost-effectiverness in future production to rcompensate for the U.gh cost of development. 6-3.4.2 Puodsacton Cost SAS production cost can be reduced by adhering to thc following: 1. Simplicity c-f design 2. Use of integrated and printed circ.uits 3. CommonalitY' ot circuit mod-Ales 4. Extensive use of value engineering principles, Production cost increases may be expected with anl increase in: 1. Number of system components 2. Quality or precision of components 3. Number of nonitandard parts \ 4. Number of parts that can be assembled in\corrft-tly

WrC 706-202 S. Number and interdependency of adjustments to be made in final assembly 6. Number of parts that can be damaged easily in assembly 7. Dogrce of cicanliness required during assembly 8. Unrealistic requirements, or ovcrtzmphasis on snua icpiesc n a. Wcght ri~duction b. Compactness c. Functional complexity d. Reliability e. Maintainability f, Structural integrity. 6-3.4.3 Msllaterancc Cvt A simple, mechanaical SAS composed of infinitelife parts (as in the Bell s".bilizer bar) requires maintenance only in the form of regular inspection and lubrication. In the event of battle damage or other failurc, repairs can be performed by a qualified mechanic. An electrohydraulic SAS, on the other hand, may require the sei vices of an instrument snecialist- an electronic technician, and a qualified helico.pter mochanic. Thus, the ucif-temt cireutit should be devised so as !a indicate exactly which section is defective. Removal and replacement of plug-in modules represent field mainteniance at lowest cost. Added to this cost, however, is the cost of maintaillilg adequate spares. 6-3.5 TECHNICAL DEVELOPMENT PLAN For the dcvcelopmeni of P conventional (e~lcctrol'yd..aulic) SAS, the plan outlinrd in MIL-C-1g244 should be followed. In addition, the airframe and rotor dynamic and aerodynamic propcriics cvcnwluly shoull he inclucied in the initial system analysis (MlL-C-48244) in order to show thec possible existenc of a~rame cross-coupling and the need forr anticross-coupling in the SAS, as well as to show the overall behavici of the SAS/airframc coml~ina' ion. Six degrees of freedom of the airframe, and quasincrnial modes of the roior (inplane as well as flapping motion o,' the bltdes), should be used. The resulting equations art used later in the simulation studies required by MIL-C-18244. The simulaticn not only will 41low the pilot to evalute the system, but also will permit demonstration of the scverul types of failure of the SAS, and will indicate time delays permiscible ;)eibre starting corrective action. In the development of unconventional SAS's, eapecisily those involving modified rotor dynamics, s,-vwal changes from the procedure in MIL-C-18244 ar mommended Unconventional systems require more inisial system synthesis, or concept selection, 6-13

A M M7-2U2 than do conventional systems; and model studies should be undertaken as an aid. The models can range in complexity from simple mock-ups of gyro and linkage arrangements, through dynamically scaled wind tunnel models, to remote controlled flying models. Tne paragraph of MIL-C-18244 dealing with model studies notes that experimental models may take the form of full-scale, engine-driven rotor and SAS assemblies, suitably mounted upon a truck bed for measurement and observation of dynamic behavior under forward-flight conditions. The maximlium possible experience with and knowledge of the system should be gained before the start of testing of a man-carrying flight article. Full-scale wind tunnel tsts, although expensive, .-an be used to test the flight article progressively to conditions beyond the extremes of the projected flight envelope. Further substantiation of the airworthiness of an unconventional SAS and rotor system can be obtained by operating an identical system on a tieiown teat, where a given nurmber of hours is requirc-i for each hour of aciuai flight icsinig. The documentation and data rcqulrcd to establish the satisfactory fulfillment of the technical developmant plan are described in MIL-C-18244, substituting SAS for automatic flight control system (AFCS).

6-4

PILOT EFFORT

The helicopter designer must consider pilot effort, or control system loads, from two points or vie,.. The first concern is the significanice of cntrol feel with redtr by the normally fly nd qualities. phyicald t fiying to ssocfin gard atios. Pilot. ofappliednorcea physical association of applied force and the mapeuve'ing response of the aircraft. iherefore. the cont.ol feel in maneuvers plays an important role in the assessment of handling qualities. Stick positioning also is a fundamental characteristic, be cause it h3lds the helicopter in the selected trim attitude when thu controls arc released. MIL-H-8501 provides for stick position trim and hold by specifying breakout forces and force gradients. The other design consideration is rclaied to the structural integrity of the components. The components shall achieve specified factors of safety when subjected to loads due to pilot and copilot effort, artificial feel devices, power actuators, etc. MIL-S-8699 wefort. covers this aspect of pilot 64.1 CRITERIA FOR POWER CONTROLS Whenever the magnitude and line-.rity of control 6-14

:-•

loads permit, direct mechanical control saill be uwed unless there is a valid requirement for power controls. Direct mechanical control is the simplest and most foolpioof control system. Howcvcr, power-operated systems may be required wbhn the control system environment contains high control forces, feedback of vibratory forces, or mixing of control forces. The control system designer must verify a need for poweroperated systems before adding their cost, weight, and complexity to ,he helicopter design. 6-4.1.1 Control Forces Not all helicopters require power actuatots. For cxarrple, on small, single-lifting-rotor vehicles, a system of weights may be installed in the antitorquerotor controls. Centrifugal force acting upon the weights balances the pitch link loads, and the system is adjusted on the ground to compensate for the control forces in cruise. The pilot t•annot trim the system in flight, and accepts the unbalanced forces in the pedals in hover and flight modac other than cruise. However, this is a small disadvantage in comparison with the aimplicity of mechanical design. Also, it may be feasibie to design a bungc spiring th.. wdil counteract the steady download in collective pitch ' Medium- and heavy-lift helicopters generally require power actuators due to the magnitude of their pitch link loads. Pitch link loads are sensitive to rotor blade design parameters, both aerodynamic and inertial. 6&41.2 Vibration Feedback The control moment of a lifting rotor blade is a steady pitching moment with various alternating harmonic components superimposed. in the nonopoet system,eermoc.Intenn rotating control these components appear as n-per-rcv forces due to the n number of pitch litiks passing over the attochment point where the nonrotating controls support the swashplate. The preoence of these vibrations in sh cyclic p stick generally is intolerable to the pilot. Vibration absorbers can be used to reduce the amplitude of the vibration transmitted by means of a simple mechanical system. 6-4.1.3

Kinematic Effects

The generation of control forces and moments along and about ¶he various axe of the helicopter is accomplished by combinctions of collective and cyclic pitch on the rotor(s), as discdssed in par. 33.3.1.3, AMCP 706-201. The motions of the cyclic stick and thrust lev,;r (and, on sonic helicopters, the motion of the pedals) are transmitted through the swashplate to the rotor(s). In some installations, the

. .

AMCP 706r202 control input

are transmitted to control-mixing

assemblies, vhcrc they arc combined botforc reaching th: swashplate. The degrees of rotor blade angle trol travel in the cockpit, are the dominant consialcration in establishing the mechanical ratios in the

INUUT B

UTPU1 A-6.)

_1 _.

INPUT A L _ 0,1

o0) OUT'PUT (A+B)

(A)INTRIM POSITION

mixing Even if control forces are low and the vibratory components insignificant. there is a €,oss.

talk of forces from one control ax-s to another because of the mixing. It is unlikely that the mixing assemblies, whic" contain components sized for stroke or travel relationships, will produce satisfactory force relationships. MIL-H-8501 sets limits upon control force cross-talk. Fig. 6-4 is a schematic diagram that illustrates the mechanical mixing of mixing coitrol assembly.signals. Fig. 6-5 shows a mechanical

I j.--s OUTPUT

INPUT B I-

o

INPUT A

f

(A-)

4•

OUTPUT (A+B) (B) POSITIVE VA._IE OF INPUT B FROM ýR.I

t

4

Figure 6-5. Mechanical Mixing Assembly

"%

INPUT B

OUTPUT OUT(A40 ) -

INPUT A

6-4.1.4 Control Stiffness At high airspeeds and disk loadin3s, the onset of rotor stall flutter can limit the flight envelope. One of the many parameters to bc considered is the compliance (stiffness) of the contrc! ,y,'-. particularly of the swashplate and. its support. Hence. another justification for power actuators is based upon rotor performance. Fig. 6-6 illustrates the installation of power actuators for tandem helicopters. 6-4.2

OUTPUT (A+B) .0 (C) POSITIVE VALUE OF INPUT ArROM TRIM Figrc 6-4. Control Mixing Scherntik

HANDLING QUALITY SPECIFICATiON

The handling quality requircinants of MIL-H-8501 sshal. be specified in the detail specification if the rotorcraft under design is a pure helicopter. However, if the rotorcraft is a high. performanc viehicle with fixed wings and alternate means of producing 6-15

horizontal thrust, the detail specificat-ons may specify requiremesits from 5oth MIL-H-'8501 and MI4L-F-785. The requirements for contrul feed forces in normal helicopter operations are found in M11-41-8501. The maximum and minimum breakouts and force gradients are defined, alor~g with th-, limit formes. No gradient is specified in thrust, because a collective sick holding system - e.g., adjustable friction om a bpc-

gcnerally is provided. There is no require-

ment for any gradient cxmcpt that it be linear from. trim to limit force. MIL-H..SS0l identifies the maximum control feel forces that arc allowable after a failure in the power boost or powur-opcrated system. The limit force in the failed mode is larger than, but of the same order of magnitude as, the limit load in the normal operating mode. Consequently, if hydraulic boost is requ;.ed fc. normal operation, dual boost probably will be required for the failure mode.

STATIONARY SYSTEM

12ROTATING

SYSTEM

-

~

ROTOR BLADE

I0

POWERED

ACTUATOR

(

STATIONARY SWASHPLATE AUTOMATIC (CYCLIC) TRIM ACTUATOR Figure 6-6. Powered Actuators (Tandem Helicopter) 6-16

POWERED ACTUATOR

4.4.5 u-. 3

IJMAN. Apcr.quii~ofora~ ectve ystm cntrl i a A per~qu~it Wan ff~vcsytemcotro i dcugn dvfiiti~on ofcoiritrcil au~mentation nee~ded as function of tota pilot workload. A force feel system may require no Pilot Control in order to maintain a trimmed flight condition. -

Ceutmi Force Co" "Thcontrol force sytem should provide. 1. Trim position identification tha: will enable the pilot to feel an out-of-trim condition and to fecl and identify trim when returning 2. Hold control in trim when the pilot is flying hands-of 3. An increased force cue to indicate increasing sevcrity of mantuvciring wherever it occurs. An increase in gradient with inciea'ing airspeed is recoinmended. Care should be taken tu avoid force cues introduced to the longitudinal control due to collective inputs. The optimal system would provide a constant relationship between longitudinal stick forces and resuiting aircraft load factor during maneuvers. The control force fc.zl system provides an im6-43.1

mediate and :lignificant cue to the pilot, indicating the

*

woptcr ricpoajonh to control conainand in any flight ct.. Jition. This tightens the loop of pilot control and vehicle response, and enables the pilot to realize optimum control. A lesser performance leaids to use of the feel system only as a trim hold device, and the pilot may prefer to turn it off tinder demanding control situations.

*

6-4.3.2 Dev'dopme,.tal T"~ Moving-bamc flight simulationa can be useful in developing the optimum control feel to suit the heclic pter mission. In the movingt-base simulator. pilots can draw upon past expcrience to identifydesired force ;eel characteristics. Stick force pro-

*

~portional to rates of control displacement, helicopterK angiular rates, and to normal accelerations should be investigated so as to insure the design of an optimum system. As the functions of the artificial feel system are ipcreased, the complexity of the feel unit also incicases. A design requirement for a specified linear gradient in the region of trim and a different linear geadient at greater excursions can result in a feel system with more than one spripg. Furthermore, ifsa rtquirement exibts for' nonlineapr forct versus doflection characteristics, cams or linkages can be employed. Fig. 6-7 is a schematic diagram of an ar'jficial feel system. Flight safety at high speeds can be increased by reducing thc occurrence of high rotor loads associ-

ted with excessive control displacement. Adynamic pressure-scnsitive (q-scnsitivc) control force fedl system produces minimum forces in hover sand maximum force gradients inh.gh-specd flight, wherc the sensitivity is ~rrcatept. This concept is an alternative to Use Of Acontrol ratio changer in the pnmary control !ankagc. The q-feel system car~ be modhanical (with q-bollows), electrical, or tlectroicat -el spoie u hydrauslic, In fixed-wing icat -el spoie u slight'y different purpose. Thc pilot flies the aitplane: by sensing, among other cues, normal acceleration and control itick forces. Response of an aiipJAne IS suhta4h cag nnrm!aclrtinjrui of elevator deflection increase with q. If the artificial stick force per unit of elevator deflection also is made to increas with q.then the relationship of stick force to normal acceleration can be made to approxima.te a constant value of stick force per g, regardtess of flight speed.N Military Specifications useful in the detail design of the artificial feel system include MIL-H-85fl1, MILS-8698. NIIL-F-8785, MIL-F-9490, anid MIL-F18372.

IC

pa.

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PWJACATF

--

tip

ITRWM coNT"V CBE

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(A)TRIM CONDITION CNRLCNEE 7

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(B)TRIM CONDITION CONTROL DISPLACED Ftgure 6-7.

CONTROL DISPLACED RMTI

CONTROL DISPLACED FROM TRIM

ArdtWIca FWe sad Trim Schematic

64.A AUTOMATIC CONTROL INTERFACES Inner loop stabilization signals are summed with the pilot's commands through electro/hydromechanic&l actuators in series wi*.h the pilot's controls. It is important that the high-ftequency, small-amplitude stabilization signals do not reach the cyclic stick in the form of formes or deflections. Thus, there is a 6-17

~ -

AMCP 70620 need for a "no-back" (a device to prevent the foed-

data were taken and the environment in which the

back of forces) located upstream of the SAS series actuator. A stick boost also will perform this function. In addison, if the helicopter is to be equipped with an autopilot that introduces signals through actuators that move the cockpit controls in parallel with the pilot, there is a requirement for compatibility amr.ong

new system will perform must be assured, or appropriate adjustment of the projected rates must be made. Another rationale for duplication is based upon failure considerations. A power actuator may provide the required reliability; but if a failure of the ac-

tVe inertia, compliance, and damping of the primary

tuator is catastrophic, a redundant actuator is re-

m.chanical controls and of the parallel actuator.

quircd. Further discussion of this subject is con-

tained in pars. 6-5.2 and 9-2. 6-4.5 VULNERABILInY The close support of grounid operations exposes the U S Army's observation, cargo, utility, and armed helicopters to small arms and automatic weapons fire, The unprotected, singlc-channel flight control system is vulnerable over its entire length. There are a number of weys to reduce this vulnerability. One naethod is to make the components so rugged that they can sustain a hit without losing their struc-

6-5

MECHANISMS

The rotating controls in the main rotor sytem nor-

mally include the rotating swashplate, the pitch links, and the drive scissors. These components are shown in a typical arrangement in Fig. 6-8. Functionally, the rotating swashplate translates along the rotor shaft

and tilts in any plane as dictated by control inputs.

tural integrity. However, this is seldom feasible, especially when space and weight must be controlled rigidly. Cert.,m areas, such as the cockpit, will be proItectec with armor piate in order to safeguard the

The swajhplatc translation and tilt arc transferred to the blade pitch horn through the pitch links and, thereby, control the main ro.or thrust vector. The of the rotating con(a)ihc fix ruoor tMe positionp scissors iu drive 'haft and rotor blades. and cruis rchaivc

crew. The same armor can te used to shield the mechanical controls. However, it may not be feasible to run armor plate all the way to the swashplatc. A redundant cintrol system not only helps Wo solve the vulverability problem but also improves flight safety reliability. To be effective, redundant channels must be separated physically. Consideration must be given to single-channel jams and disconnects, to adequacy of control if the remaining chann-.l goes to half gait, and to the question of whether the configuration should be active-active or active-standby.

(b) provide the load path for the conversion of drive shaft torque into the tangential force required to induce rotational motion in the rotating controls.

64.6 RELIABILITY

the rotor blades complete one revolution, the pitch

6-5.1.1 Design Factors Structurally, the rotating system shall be designed to withstand the alternating (fatigue) flight loads introduccd by rotor blade torsional moments and the maximum loads introduced by severe flight mameuvcra or during ground operations. The fatigue loads are periodic, and alternate primarily on the basis of once-pcr-rotor-tevolution. In other words, each time

The overall reliability of a flight control system depends upon the reliability of the individual components and upon their arrangement, which may be either in series or parallel. If the helicopter system specification prcscribes a minimum acceptable value for flight safety reliability, this value may be so high as to require dual mechanical controls. The detail

link load completes one stress cycle. Therefore, a high-cycle fatigue evaluation is rcquired. The primary loads are discussed in pars. 4-9 and 4-10, AMCP 706-201. and the fatigue evaluation is discussed in par. 4-11, AMCP 706-201. In addition to the primary Pisht loads, special consideration shall be given to socondary loads. Failure

designer first must establish the single success path;

to evaluate secondary loads properly may lead to set-

then, if system reliability is inadequate (a value less than required by the helicopter system specification), he must add redundancy, beginning with the least

vice problems. Among the secondary loads that shall be considered are frictional moments in rods and bearings, and bending moments created by centri-

reliable componn'nts,

The reliability of a component is a function of the

fotgal force. A typical pitch link rod end, with a selfaligning bearing, is shown in Fig. 6-9.

historical mean time between iailures (MTBF) of that

Bearing motions of -6 deg are not uncommon

component. When historical failuic rates arc used, similarity between the tav,ronment under which the

during each rotor revolution. The normal force (pitch link aad) times the coefficicnt of friction produces a

6-18

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M~CP 706-202ý

ROTOR BLADE-,

LISTATIONARY

SYSTEN.7',

ROTATING SYSTEM

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2 TOIALS 3 TOTALS

Figure 7-4. Example AC La"s Analysis Format

--

.i-.

The exciter functions so as to supply DC to the min&! blocks and/or connector restrictions, main field winding. T-he poles are on the Ftator and temperaturealtitude environment, availability, and on a disturbed widing in dlots on the r, or, Both characteristics of air or liquid cooling members are laminated, as in the. main _ncrator. S. Adequate systemo growth - designer mrust con1 n C'.9iatuf A3 4"A tvk'irUGA~tU&, UAnd Provides LO% tv ocurcc. a power aetecting ad-r future growth when main field through a bridge of rotating rectifiers. the ALERNTOS) 7-22 C GNEATRS 7-2, AC ENERTORS(ALTRNATRS)In order to achiove a self-sufficient generator ovie not depndent upon any external power source An AC power source capable of insuring a power for excitation power - a permanent-magnet pilot exby specified that than, quality equal to, or better MIL-STD-704 can be achieved best by utilizing the citer completes the lc~tsical portion of the AC g-Derator. This generates either polyphase or singleattributes of the conventional salient-pole. synchronAC power, which is rectified either within the our aterntorphase The heheopter ACý generator is comprised of the generator (stationary rectifiers mounted within the niain generator, an exciter, and, in the mejority of generator case) cr in the voltag#- regulator (Jpar.7-4) to provide DC excitation power to the exciter. Decases, a permanient-magnect pilot exciter sharing a pending upon system requirements, the magnetic common housing and shaft. Modern-day alternators pilot exciter may provide control power, protective are brushloss; 6:.I,, no brushes, slip rings, or commucircuitry, or operational power necessary for proper tatots are employed, functioning of the distribution system. 7-2.2.1 Electrhca Desigc 7.2.2. Mecbmalcal Desig The main generator consistt of a stator and a rotor. AC generators are. housed in either aluminum ur uniare The stator is built of steel laminations that housings. The selaction of housing magnecsiunt the co'itaan and periphery inner the orn formily slotted output w~ndings. Thesw windin. are connected in a material is dictated by weight and/or vibration requirements. Lamination eteel for the magnetic cirnormal three-phase., (our-wire mnrnnrcr, and are discuit iseither a silicon or co~balt alloy. The latter conplacrd so as to minimize distortion of the output struction results in a relatively expenisive generato., lamiiof oonsists field voltage wavefrmn. The iotor or but provides a weight advar~tage o' almost M0over a the and "pol,ýs". form to as wo lunched nations system using silicon st*.tl punching.. nuniber ef poles and the rpm fix the output froPrac~ical generator speeds range from 6000 to 12,quency. The main field winding is wound on the rotor rpm for 400-Hz output. Variblefrequency 000 the Je prcivi poles @ad is excited with DC. This in ratings to 120 IrVA art paT~ical to machines sufliciunt provide to necessary mafutiomotive force spews of 20,000 rpm. Applications of 6W0 to 12,000 lUMe of force (flux) In the magnet circuit of the main rp'q require grauz~-lubricated ball bearings with gaweator for adequate, all-load-condition, voltag bearing lives of MW0 hr in an aveane helicopter en2W.1iuron, 7-6

-

Vil

Vl

__

_

_

_

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7-2.23

DESIGN DATA EQUIPMEN CHARACTERI_,_os TIS-

CARD LOAD ., ,1DSORT

IPT TAPE

were not considered. In the case of blast-coo!cd generators, inlet airflow and pressure-altitude characteristics of the separate forced-air supply must be defined adequately in order tis.al'ovi proper use of a gcnerator cooled in this manner. Cooling-air temperature versus altitude and ambient tempcrature data are vital aspects of an adequate cooling specif.cation as discussed in MIL-G-6099. Contaminant

EDIT LISTING

--.

ERRORS DETECTED

Cooling

Cooling, a primary requirement in an AC generator specification, may be accomplished by air or by liquid. Both air-coaled and oil-cooled generators are for helicopter applications. For air-coolcd generators, either self-cooling or blast cooling may be employed. In the case of selfcooling, an integral fan is located on the rotor of the machine, and diffusers or baffling are employed to direct the air over the hot internai surfaces. If the in. stallation is such that ducting is provided to the fan inlet, the pressurz-flow characteristics of this ducting must be considered so that all through-generator airflow requirements are met. A self-cooled generator that exhibits good performance in the laboratory may burn up on the airframe because duct restrictions

EUIO 1 0ORMATIONemploked W

EDIT INPUT

706-202

_olingAMCP

_

lprotectiorn of cooling air is required.

Oil-cooled generators fall into two categories, conduction-cooled and spray-coaled. Oil-cooled with inlet oil tempe,-atures generators are practical from -65' to 330 0 F.

YES CANCEL RUN

"Inthe conduction-cooled generator, oil is circulated through closed passages in the housing and rotor shaft. Cooling is obtained by conduction of heat to the oil from the hot windings. The bearings use the oil for lubrication as well as for cooling, and rotating seals are required. The weight of the

Y

ic n- .-. . .

h- - ,-

that of the air-cooled generator.

TAPE

LOAD ANALSSI REPOR

Figure 7-5. Typical Autoration Flow Charl vironm-nt. Oil lubrication of bearings generally is necessary tor generator speeds in excess of 15,000 rpm. Bearing life in excess of 10,000 hr has br-n

~) achieved in helicopter generators using oil I

i-

cation. A typical oil-lubricated bearing AC gene Ator is shown in Fig. 7-6.

In the spray-cooled generator, the cooling oil, in effect, is sprayed directly on the windings. This rest-its in an improvement in heat transfer, along with a weight reduction of approximately 15% compared to the air-cooled or cgnduction-oil-cooled generator. To date, all spray-oil-cooled generators have been applied to 400-Hz sy:tems, and operate at 12,000 rpm. For comparison assume a 90-kVA rating; a modern air-cooled generator using magnesium housing and cobalt alloys weighs approximately 90 lb. A spray-cooled generatnr with the same rating weighs 55 lb. A generator weight of approximately 0.5 lb per kVA is aciicvable with apray cooling. If

spray cooling is used, it is necessary to zcavenge the generator cav:ay, i.e.. to remove excess oil fesulting from spraying of the windings. 7-7

TTRýT

AM-'P

Figure 7-6. Typical AC Generator with Oil-lubricated Berrings 7-2.2A4

I7.

App~lcation Checklist

The criteria that follow should be considcred in the dtaian of all AC generators. This is a minimum list of characteristics that must be defined and adapted to a given application: I. Rating: a. kVA at required power factors b. Voltage at terminal& c. Phases d. Frequency 2. Speed, rpm 3. Maximum weight 4. Envelope, diamiemF, and length S. Mounting detals 6. Cooling requirements

\2.

Applicable Military Spocificatkicas regarding

I' . Waveform 12. Performance requirements under unbalanced

load conditions. 7-2.2.5 Variable-frequency AC Generators Variable-frequency generators are practical in ratings to 120 kVA at speeds to 20,000 rpm. The previous discussions relaive to mechaniacl design, cooling, and application checklist generally are applicable also to the variable- frequency g--ncrator. There are twe significant performancne characteristics peculiar to the variablc-frcquency generator that should be considered prior to its application: 1. Voltage transient performance at high speed Voltage regulation problems over a wide speed

range.

generator and/or system performance, eleciromagneti; interfcience, vibration, etc. 8, Minimum africiency 9.Overloads and time at each overload

Voltage transient performanice at high speed for a wide-speed-range generator (e.g., 1.S: 1)can result in severe system problems. This is because the maximeum voltage attainable from the gcnerator £t

circuit

tainable at the low speed. Upon application of load,

10. Short-circ, tit cturrent capacity

7-8

and time at short-

the high speed isthe speed range times the voltage at-

AMCP 706-202 severe d~ps in system voltage could be experienced. Wihregard tovoltage regulation, whnaspeed range approaches or exceeds approximately 2.5:1, high-speed instability c~n result. Because the regulator is called upon to adjust from an overload at low sped t noloadoWading power factor loads at high orde

of 5:1.portnce.Theweight

7-43 SrARmF/GENERATORS, DC CKNERATORS AND STARTERS Stawe-of-thc art DC systanms for helicopter applications are designed for operation at a nominal 28 V. with power quality defined in accordance with the requirsmnuas of MIL-STD-704. For the majority of applications, advantage is taken of the volumetric efficiency and lightweight properties of the DC starter/generator. Nevertheless, there exiit many applications that, for various reasons, eniploy both, a DCgenerator and DC. starter: The construction of the DC startcr/generator i, '

~comprised of a rotating artnature and stationary

field.d The armature is constructed of Aistack of steel a"laminations uniformly slotted on the outer periphery; the power windingV art connected to the coinmutator and are placed into the slots. The stationary field consists of laminated main poles, interpoles, and a solid steel field ring to which the poles arc attached on tlhe Imner peziphery,

LOn

the main poles are wound the main field win-

dinga, connected either in parallel (shunt) or in series with the armature winding. The interpole coils are conn, -led in series with the armature. A fourth winding - distributed in slots, placed in the pole faces, and connected in series to tht armature - serves to support the main field and to overcomec the do. magnetizing effects of an armature reaction to the magnitic field set up by load currents flowing in the *armature winding. This winding is termed the coinpensatieg winding, and generally is employed with gencertor ratings of 200 A or more. The starter/generator nornuilly require a fifth winding consisting of a single turn in stries with the armature and wound on the main, poles. This winding, during starter operation, aids in increaing the torque output per limpere of input current. Some manufacturers leave thii, winding connected during geerator operation an a differential compound windigthat aids in the tagulation over fte load and snpeedi range. The discussion of the mechianical construction of

the AC generator is applicable generally to the DC starter/genierator as.Ioretochvethe lightest possible weight without sacrificing mechanical integrity, both aluminum t.nd magnesium are used for housing materials, the choice being related directly to th:ehria nvironinefltrequirements. advantage of a DC starter/ gnrtremp3oying this high-permeability material, cmaetoaunit employing silicon stiedl punchings, isof the order of 20*. The starter/generator normally operates from stand-still to speeds of 6000 rpm in irelation to starter mode operating speeds. For genefator operation, it ordinarily '-overs a speed range of approximately 2:1. with 3000 ryi. i the usual minimum speed. and seldom is applied whta- maximum speeds exceed 12,000 rpm. State-of-the-art DC istarter/generators generally employ greasL-lubricated ball bearings. For hclicopter usage, thz bearing life generally falls between 1000 and 3000 hr. The brushes that ride on the commutator and cn duct the current from the pwe surc - in the case of the starter - and to the load - in the case of the generator - are made normally of carbon and copper. Because of altitude requiriements, the brushes are treated with u compound (such as molybdenum disulfide) in order to provide the necessry filming characteristics under c jnditions of low oxygen and moisture. For starter/genecrators, brush life is limited to 5M00D10 hr, depending upon the severity of thec

start. For those applications requiring Senerrtor

operation only, brush lives of up to 2000 hr arc po. sihe.. Air cooling of the DC starter/generator and generator is standard prazf~ice. Contan~i~nn" pro. tection of this cooling air ic required. This cooling may be accomplished by integral fan (self-cooling), by blatt cooling, or by a combination of the two. Pre.cautions are necessary in order to define the cooling conditions adequately. Becmuse of the problem of providing adequate heat transfer from the-brushes and commutator to the oil, oil cooling seldom is em. Pioyod. Following is a checklist of minimum in'.ormatic noo.esaary in order to dcfine ad1equately &stafter/ generator for a given application: 1. Engine type and manufactukcr 2. Intended installation 3. Envelope requirements (diameter and length) 4. Maximum allowable weight 5. Maximum alIawabke overhand moment 6. Engine mounting details

-

7-9

AMCP 706-202 23. Percentage of maximum generator output used

7. Applicrble specifications, if any

at engine cruising speed

8, Type of cooling - blest, self, or other a. If blast cooling, prc.surc available b. Temperature of air 9. Ambient temperature range 10. Altitude requirementsi II. Direction of rotation facing engine pad 12. Engine to starter/gener.dtor pad gear ratio 13. Power sapply for starting: a. Battery, type and voltage b. Giound power unit, type and .altage 14. Engine torque versus speed curves for standard conditions and -65 0 F (or the lowest applicable temperature), plus a notation of whether or not thtsc curves include accessories and gearing IS. Engine light-off speed if not shown on curves 16. Starter cutoff speed if not shown on curves 17. Maximum allowable time to light-off speed

24. Voltage regulator type and applicable specification. For engine starting, either a ground power supply or aircraft battery is used. Ground power supplies generally are of the constant-current type, and provide the best power source available for engine starting. In the majority of helicopter applications, where starter/generators are used, aircraft batteries are employed for starting. The batteries are rated 24 V and are either silver-zinc cr nickel-cadmium (par. 7-3). If multiple batteries are used, they may be connected parallel or in series to provide the desired starting characteristics, but consideration must be given to the applied torque vs generator shaft shear section and the engine gearing limitations. A typical startcr/gcnerator used in helicopter ap-

18. Maximum allowable time to cutoff speed 19. Starter/generator pad rpm at engine idle

plications is shown in Fig. 7-7.

20. Starter/generator pad rpm at minimum cruising

7-2.3.2

DC Generators

speed 21. Starter/generator pad rpm at maximum engine

The helicopter DC generator is identical electrically and mechanically to the DC starter/generator,

speed 22. Required generator output and vo!ftge under all speeds in range of regulation

with the -xception that, generally, no series turn is ciployAd ou thc main ficid winding. The preceding discussion relative to mechanical

I

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4

DC Starter/Generator

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K.

AMCP 706-202 design, speed rang, beatings, hrushts and cooling also are applicable here. It is suggested that reference be made to MIL-G-6162 as a guide for specification preparation. A typical blast-cooled DC generato., is shown in Fig. 7-8. 7-.3.3 DC Starters With the advent of the DC starter/generator, the DC starter now has limited application. In the past, the majority of applications were found on heli. copiers employing reciprocating engines, and, therefore, these starters were designed for crankink sCr. vice rathcr than for the type of starting service required by the turbine engine. For cranking. the starter operates at a nearly fixed speed until engine ligh,-off. Present-day starters are designed for starting the turbine engine, and the starter operates over a speed range of 0-20,000 rpm. The DC starter/motor is quite similar, electromechanically, to both the DC generator and the DC starter/generator. Ordinarily, often three windings are used, as opposed to the five windings often

empioyed in gcncriurb. Th, 5tt v ii, ai ",IioIIof't"'

') -

-

~,

lease rate of storage batteries is a matter of common practical knowledge. The engine.startilSg cApability of a fully chariged battery in a 70°F environ.ntnt is reduced to 27% after the battery has been stabilized in a commonly experienced -25°F environment. Experience has shown that the probability of a successful 0*F start capability decreases rapidly as the level of the battery charge decreases. Gas turbine starting problems are not restricted solely to the low end of the temperature spectrum. A less obvious, but very real, problem is encountered at the high end of tha temperature range. The hot start, in which the temperature within the turbine exceeds the safe operating rantije for turbine materials, is encountered all too frequently in high ambient temperatures during the starting sequence of gas turbinepowered helicopters. The hot start can be caused by a number of conditions. A battery that is not charged sufficiently and improper fuel control adjustments are two conditions that can lead to critical hot-start problems. Another cause of hot starts is the timelapse needed to accelerate the engine from light-off to ille speed. The net result of hot starts, whatever the cause, is premature engine failurc. Evmi thiigh a a1', '

main field with the armature and interpole winding provides the highest torque per ampere of current, and lends itself to use with battery power supplies, As in the starter/generator and generator, starter/motors employ grease-lubricated ball bearings. Brushes usually are .of the low-contact drop type, and contain a high percentage of metal so as to minimize voltage drop at the high currents required by the starting cycle. Because of the short duty cycle requirements, an integral fan provides required cooling, Thc DC starter/motor serves no function after starting the engine; therefore, design provisions to disconnect it from the engine accessory drive must be made. This usually isaccomplished by a mechanica!ly or electrically operated jaw engaging and disengaging mechanism. A turbine engine starter is shown in Fig. 7-9.

start might not result in an imanediate and catastrophic engine failure, it will shorten the times between engine overhauls, thereby increasing helicopter operating costs. The characteristics of electrical starting systems (and starter/generator systems) for turbine engines generally are such that maximum torque is delivered from the starter/motor to the turbine upon the initiation of the starting sequence. Starter/motor output torque decreases approximately as a straight-line function at increased turbine speeds, with the starter/mr3tor torque reaching a very low value at starter/motor cutoff. The starter/motor is designed to have a stall torque capability exceeding the torque limit of !he engine accessory drive system. Adding a boaot feature to the electrical starting system provides an additional supply of starting

7-22.4 Boost Starting System Modern military practices dictate the use. of turbine-powered helicopters for a variety of reasons, one of which is their ability to use completely unprepared terrain for takeoff and landing. Therefore, deployment of helicopters away from airfield support functions is common. Deployment in hcO or cold temperature environments having no APU or ground power facilities requires t.Oat batteries not lose their gas turbine engine-start (apabilities. The effect of cold temperature on the energy re-

c-ptually, the boost power is obtained from a combination of gas generator and gas motor, and by coupling the gas motor with the electrical starter/ motor. This boost feature is a supplement to, and in no way is intended as an alternative for, the electrical starting system. Even though the added feature is used only occasionally, the components added to the electrical starter, or starter/generator system, nevertheless must be taken into account, and their effect must be established upon such elements as overhand moment, vibration characteristics, operation of the system in a generating mode, and the overall physical

energy for extraordinary starting situations. Con-

7-Il

Figure 7-.

Biast-aemled DC Guemmfor

Figure 7-9. DC Starter Motor With Solemold-opersted Switch 7-12

CA.-

Flgure 7-10. Prototype Cartridge-boosted EPeiurkal Starter System constraints of specific engine installation. Prototype *hardware is shown in Fig. 7-10. MOORSvironmental 7-2. ELETRICL 7-2.

.-

7) ..

ELETRICL MOORSuniversally

Electrical motors for use in helicopter el.ectrical systems may be either AC or DC, depending upon the primary power source sekocted. AC motors in use are almost universally of the squirrel-cage induction type. They may be three-phase or single-phase, wvith the number of phase to be used being governed prin. cipally by motor size and by the characteristic* of typical load requiremenfts. The squirrel-cae induction motor compfnss a laminated statoi - identical in configuration to that employed in the AC generator - and a'laminted rotor. The rotor has slots on the outer peri~hery containing either copper or cas-. aluminum bars - shortcircuited on both ends - with end rings of the same material so as to form, ultimately, what resemble3 a cage. The input winding is contained ini the stator slots and maty be wound threve-phaze or two-

'squirrel

\ phase.

Housing materials maylbe either aluminum or magnesium, depending ~upon weight and enconsideratimis, and bearings are almost of tho groaeiawbricated ball variety. For the specification of ti~c AC motor, the best guide is contained in MIL-NI-799, The factors involved in construction of the DC generator generally apply also to the DC motor, with the ex~ception that seldom, if ever, arc pole face compensating windings employed. Depending upon tbhe load, the DC motors may be used in series, shunt, or compcund winding configuraiions. For very small motors, pcrrnanent-magnet fields are employed in place of the shunt or series field windings. MIL-M8609 is recommended as a guide for specifiution preparation. In general, the electrical motor may be termed a torque device, inasmuch as a certain volume of iron and copper is required to produce a given torque. The sizz ot the motor is dictated by the torque requirements. For helicopter applications, where size and operate weight are at a premium, it is normal to 7-13

S400

motors at high spoeds in order to obtain a high power output (power equals the product of torque and speed) pet unit volume. This, fundamentally, is why the majority of aircraft AC power systems operate on rather than W0Ha power usually em40- Hit forahr tndhsnrial thj 60nHzcowerial usea. Hmplyayd for industrial and commercial uses. Highspeed operation poses certain problems in bearing and/or brush life, which, when couplea with the fact that the airframe itself has a relatively short life, resuit in overhaul rates measured in hundreds of hours, rather than in the 10-20-yzar life rates considered normal in industrial and commercial ground applications. High-speed, high-power-per-unit-volume operalion points up the design parameter that usually determines the size of the helicopter electrical motor. Increased losses follow increased power output, and the size of the motor must be adequate to dissipate these losses without exceeding motor material .temperature limits. It follows that the availability of effective cooling directly affects motor size. High altitudes, with low air density, will decrease the cooling available from a fail. High ambient temperatures rethe he.. urm,':fer and add to the mofOr tot temperature. Also, there is a significant amouat of beat generated in tlýc rotating member of the motor during acceleration from standstill; and if repeated starts are made, motor temperatures rise 5ignificantly. Another significant item is the effect of voltage and/or frequency variation on motor life and size. These items must be accounted for in the design so that, under the worst conditions, the motor continues to produce the required speed at the spv Tfied torque. This means that, for all other conditions, the motor will be operating at high-speed, high-power output

7-2.5.1 AC to DC Convetters Devices for converting AC into MCare of two basic types: rotary and static. Rotary systems may be either AC-driven motor-generators or synchronous econverters. The latter are essentially DC generators in which slip rings have been connected to the armature winding by equidistant taps. The synchronous converter, in effect, combine the functions of an AC drive motor input with those of a DC-generator output, although with less flexibility in voltage and power-factor control than the motor-generator combination. Converters typically are cheaper, more efficient, and more compact than corresponding motor-gcnerat-'rs. It is important, however, that they operate as near to unity power-factor as possible since their "rating", i.e., relative output, decreases rapidly wiih decrease in power-factor. Table 7-1 displays this relative output relationship for various power factors and number of phases. Converters also must be synchronized with the input AC supply. The relationships between the AC and DC voltages and currents are functions both of the power-factor and the number of phases (hence, number of Elip rings). Converters may bc singic-phnasc, il whiuia nb

there are two slip-rings und two slip-ring taps per pole pair; three phase, in which there are three sliprings and three taps per pole-pair; and so on. However, because of the sensitivity of the output to the total number of phases, converters usually arc operated with six phases. With a sine-wave input voltage, the DC voltage is the peak of the diametrical AC voltage, the latter being the voltage between any two diametrically opposed taps.

At unity power factor and a typical 95% efficiency, the DC and AC currents are equal with three slip of input power, the larger the motor that is required. rings; with six slip rings, the DC current is twice the The electrical motor also is Atorque device in the AC current. sense that andimposed power by input, etc., This are direct results of thespeed conversion firom units torque AC to(TRU). DC is accomplished the load. leads byStatic transformer-rectifier The rectifier to use of motor speed torque curves. These curves concept essentially acts to block conduction during arue,smplyothr sequilibrium opetiorquecu . these cone-half of the reversing alternating current. Thus, are current is unidirectional, but only during one-half of motor. The a peed, current, power input, etc, the sine-wave cycle. B3 combinift, two units and a those values that occur when the motor operates ata ii center tap from a transformer, rectification throughgiven torque. out the full sine-wave can be achieved. A smoothing inductance connected in series with the load in such a device effectively smooths out the wave peaks to a relatively small pulsating rippt e. 7-2.5 ELECTRICAL SYSTEM CONVERSION In practice, most rectifiers c.mploy bridge circuits The typical helicopter requires both AC and DC to achieve such ful-wave rectification; however, the power. Therefore, the ability to volnvert one to the active components of which may be diodes or other as needed also is required. Various types of dethyristors (silicone-controlled rectifiers). The output vices are available by which this conversion can be rulti are a series of square current pulses, which effected, together produce a continuous current output. Ripwin, high•" e.

7-14

hg., -'-

.

9|a

AMCP 7

--

-_

_

_

-202 _

TABLE 7-1. OUTPUTS OF CONVERTERS RELATIVE TO CONTINUOUS-CURRENT GENERATOR CONTINUOUS-

SINGLE-PHASE

THRLE-PHASE

FOUR-PHASE

SIX-PHASE

CURRENT GENERATOR

CONVERTER

CONVERILR

CONVERTER

CONVERTER

100

85

132

161

194

95.5

100

78

120

145

170

90

100

74

109

128

145

POWIR-FACTOR, %

100

ole voltages in the DC output can be avoided by the use of polyphase input supplies.

7-3 BATTERIES In general, battery selection is based upon battery

MIL-STD-704 specifies that utilization equipment requiring ar. AC input of 500 volt-amperes (VA) or more and a 28 V DC output of 5 A or less shall use static conversion, unless it is designed sp-cifically for use with DC generators.

characteristics, electrical characteristics of generators and associated controls, utilization loads, and cartain assumptions in unrcraft operations.

DC to AC Converters Devices for converting DC into AC also fails into

Nickel-cadmium and silver-zinc storage batteries preserntly :r used in aircraft eiectricai sybicvis.a.-

two fundamental classes: rotary and static. The rotary class are basically the same types of devices used in AC-DC conversion - they may be either motor-generator sets or synchronous con¶verters; the synchronous converter having the capability of operating on DC and converting into AC. In this condition they are said to be operating inverted and therefore are known commonly as inverters. In general, what has been said regarding motor-generators versus synchronous converters remains true here the inverteirs are more cmlqcrat, cheaer, and more compact than comparable motor-generator sets. In this DC-AC mode, however, some suitable electrical or mechanical speed control must be used since converters tend to "run away" in this condition. (The highly inductive load weakens the field through armature reaction and allows the speed to increase.) For most ordinary, i.e., relatively low-load, appli. cations the most common DC-AC conversion device is the static inverter. This is a circuit that alternately connects the output lines to opposite side of the DC supply typically via the use of such solid-state components as thyristors. Such semiconductor inverters, using both thyristors and transistors in three-phase, full-wave bridge circuits, are in conventional aircraft inverter in rotaryupon and static use. specific The choice castbetween must depend such parameters Sany

electrochemical system has particular service characteristics. Nickel-cadmium exhibits excellent cycle life and output over a wide range of discharge rates, and is preferred to other systems for starting turbine engines. Silver-zinc gives the highest electrical output per unit weight and volume, but is the most expensive of batteries and has the shortest cycle life. Lead-acid is the oldest of the systems, and is used in helicopter design not requiring main engine starts. It is well-adapted to most conventional electricai circults requiring mcierate discharge rates. lypical comparative characteristics are listed in Table 7-2.

S2,5.'.

. as the available input sources, the required output - chnasacteristics o; the system, and the relative costs and weights. •

•

°"'•

7-3.1 BATTERY CHARACTERISTICS

GENERATOR CONTROL BATTERY CHARGING Engine-driven generators are subject to variations in speed in the approximate ratio of 3:1. Thus, a voltage regulator must be provided in order to maintain constant voltage at high engine speeds. The generator is dropped off the bus by the reverse current relay at all lower engine speeds, and the battery must assume all utilization loads at engine speeds below the cutoff value. Various generator and voltage regulctor combinations have been designed that maintain svrctem voltage down to idling speed. in manner limited, currentwill If thethegenerator dependtheupon the utiwhicih battery isischarged lization load current during a specific time interval. and upon the state o :barge of the battery in that period. 7-3.2

-7-15

TABLE 7-2.

U

"

TYPICAL CHARACTERICS OF 24 V, 34 AH BATERY SYSTEMS .. E IG H T.

'T ' P• EW

- hr / b W

I

~~~RETE•NT ION C U RREN• T

10-13

50 x CAPACITY

9

7

93%

19..;

78

10-13

20 x CAPACITY

5

3

90%

12

1.9

34

2530

6 x CAPACITY

17

16

88%

17

1.4

NICKEL-CADMIUM

75

LEAD-ACID SILVER-ZINC

When the gtnerator output current does not reach

"thecurrent-regulator limit, the battery will charje at constant voltage. The charging current will be determined by the state of charge of the battery during the time .intervals considered. When the generator output current tends to exceed the current-regulation limit, the regulator automatically reduces the " generator voltage so as to lim;t the current to the these current conditions, value.atUnder regulated equalthe to battery thc oafa constant will be cha.rae,• ference between the regulated current and the utilization-load current. The increase in battery stateof-charge 6, in this ".ase can be computed by the followinm formula: A-hrSuch A-hr

(7-2)

where Ic c

- charging current, A - charging time, min -

.1,2.

j

overcharging at constant voltage can result in a condition called *'thermal runaway". This is an un: controllable rise in battery temperature that ultimatcly will destroy the battery. As the temperature increases, the effective internal resistance decreases, permitting ever-higher currents to be drawn from the constant-voltdge source. This in turn decreases the resistance still further, in an ever-increasing spiral. the battery the over-all In general, cha-o"n-- ofSOMA" durinscondition should be monritor,,d lations incorporate control systems (battery conditioner/analyzers) which monitor tewiiperature and state-of-charge, constantly analyze the geneial battery status, and cut the charging proess on and off ar conditioas dictate. systems typically us,; other approaches to battery-charging, some of which are itemized in Table 7-3, along with tweir rincipol operational characteristics. 7-3.3

0.8 - charging efficiency including battery When total system loading charging - exceecs the continuous rating of the generator, a current regulator limits the output current to a safe value. This means that the battery m ust. supply the difference between the utilization load current and the maximum generator current. In any case, when the system is operating, the batte~y is always either charging or discharging at some rate determined by the demands of the overall utilization system and the output of the generator. "The actual conditions of the charging state at any time are controlled by the generator voltage and current regulator. Battery temperature, especially in nickel-cadmium batteries, should be monitored to prevent overheating, which reduces capacity. A combination of high battery temperature (i.e., in excess of 150•°F) and 7-16

" ,

CEtL VOLTAGE

AT200A, 0 of

AT 2hr IuIRATE

14DAYS AT 80"F

CELLS

AT200A, 800F

lb

0.8I

'AE• O F C H A RG •. N o . O F

CAPABILITY

TYPE

S-

h /

- h / b-

UTILIZATION LOAD ANALYSIS

I he utilization load assignments should be based upon the most dcmand;ng conditions likely to be cncountered during operation. For example, it shoud be assumed that the aircraft is oper, ting at night, with landing lights used during takeoff, climb, and landing. Approximate data conoerning the duration of each load should be known or assumed. To obtain conservative results from the load analysis, the intermi:tent load peaks usually are considered to occur concurrently. Despite the short duration, heavy loading, such as during engine starting, can reduce the battery capacity sharply and should be considered. The load analysis is prepared for an arbitrary set of operating conditions. Too svere a set of conditions would overburdcn the power system during Fpst operations. An overly optimistic choce ot conahtions would limit the usefulness of the aircraft. The

.

)

AMCP7M ALTERNATIVE CHARGING METHODS

TABLE 7-3. METHOD

r

IF EECTONBA-TFRY PERFORMANCE OPERATIONAL

MAINTENANCE

PRINCIPLE OF OPERATION

CHARGERCONSIDERATIONS

SYSTEMCONSIDERATIONS

CHARGED AT CONSTANT BATTERY TRUIECONSTANT POTENTIAL

MAX ELECTROLYTE FAST RECHARGE 055 15LE LOSSEACHCYCLE

VOLTAGE. INITIAL CHARGE CURRENTLIMITED ONLYBy SOURCEANO LINE IMPEDANCE

DOESNOT REGUIRE CE ES MATCHED CLORELY

BATTIRY TEMIPERATU4RE REQUIRES SENSOR

BATTERY CHARGED AT TEMPERTATURE.COMPENSATED VOLTAGE, CURRENTLIMITED INITALLY By CONSIDERATIONS BATTERY

CON;TANI POTEN'IAL/ CONSTANT CURRENT

SIMPLEDESIGN

REQUIRES W1IE RANGEOE: POWERHANJILING CAPABILITIES DOESNOT AM~IRE FACII [TATE$ ZOOO MATCHING OF CELLS MATCHED CLOSELY QUICK RECHARGE REOUIRES CHARGERVOLTAGE TO B1tTTERY CYIARACTERISTICS TO AVOID THERMALRUNAWAYOF NICKEL CADMIUM CELLS

ELECTROLYTE LOSSEACHiCYCIL

REQUIRESBATTERY LOGEPTEMPERATU1TE SENSOR REQ~'IE MULTI-LEVEL. CONSTANT CURPENI

~

TR CHARGE REDUCEOAS STT-CAG NRA;T STTPFCAG NRAE

Ek- IARGE TINE THNONAT OENSTIAL T

EECRLYE ELOSECITROCYCLE YETA OSE~

YIDUCED PUWERCAPABILITY AS OPPOSEDTO C NWANT POTENTIALMETHOD)

CELLS MATCHED REQUIRAES CLOSELYFORCAPACITY

DTIFFICULT TI. DETERMINE OliMuM CHARGE TERMINATION CONDITION BATTERY EPFOIIIRYS SENSCR TEMPE[RAYLIRE

CHARGING CONTROLLEDOR PROGIIAMVIEDSO'HAT EATTLYT RETURNSTO FULLY CHARGED ELECTROL YTE PEARCHNAGE 1PPC)

PULFIE CHARGING

I AS' RECHARGI EORSINEATG

PT AC- BATTERYRfCIURNS WHEN BATTERY FREQUENCYOETERMINED CURACY OF STATE-OF-CHARGE TO FULLY CHARGEE NOT!YIJILY SENSING CZVICE. WHICH MUST CONDIitON CHAkGED; BE RESETPERIOOICALI.Y TO NULL OUT ERIRORS

CHARGERPROVIDES REVERISE CURRENTPULSEAFTEREACH CHARGING PULSETO CAUSE DEPOLARIZATION OF BATTERY ASPLATESTO ALuOWBETTER

ELECTROLYTE LOSSEACHCYCLE

SOF-TION OP CHARGING CURRENT

ORAhINUDVC Q01LIT MONITORING

RQUESCL!MTHD LSEYORAAIT

REQUIIRES COMPENSATION FOR CHARGE EFPFICIENCY ANDI POSSIBLYSTANDRYLUJSSES

SHOULD REDUICE CHARGFTIME SINCEHIGHER POWFll HIGYH CHARGERATES REQUIRES C.YPABILITY CAN BEUSEDAS

REQUIRESCELLSMATCHED CLOSELYFORCAPACITY

COMPAREDTO CONSTANT EQ.-

POSSIBLYAEGUInES BATTERY EPPEC FOREMOST REDESIGN,

POSSIBLE EM, PROBLEM

YTlE__ITII

jSHOULD REDUCE C'HARGETIME

PULSED CONSTANT

CHARGE CONSISTSOF CURRENT ,TJL5ES WHOSEPEAKVALUES MAY ELECTROLYTE BE AS HIGH At BOOA. WHILE CONTROLLING THE DUTY CYCLE LOSSEACH CYCLE TO OBTAINITHEDESIREDAVERI-

AGEVALUE OPCURRENT

I

______

I

choice is, necessarily, a compromise. For most aircraft, the following is sugge~sted: 1. Bittery opcrating temperature, 00 F 2. Duration of flight 3. Night operation. See Fig. 7-11 for a sample set of utilization loads roic. n CUR basd b Es tpialmisin VY EN STAiclmsso R TIN 7-3.4 HE V U RN T R IGinsure REQUIREMENTS s\ome aircraft crigine, have starting characteristics

AS COMPARE TO CON STA:NT) EYETNTiAL

HIGH POWER REQUIRES CAPABILITY

IZATI)N

RQIE EL ACE ACE EL RQIE CLOSELYFOR1CAPACITY PSIL PSIL

M M

RBE RBE -

.

that place severe loads upon batteries, particularly in extreme low-temperature crnvironimenu, or where the battery is not used for long periods. Under such cond i~ons, not only is it more difficult to start the engine, but the battery itself is less active electrochemically, causing the internal resistance to increase greatly. Additional energy must be incorporated in order to a reliable system. This can be accomplished by thec use of parallel-series connected batteries, to sa to provide a marked increase in the available voltage 7-17

AMP 706-202 •- -. f ioq

I

I~

CURRENT

TAKIING

Ln c

'IIIPMN

4

62

62

SIARtLH

1

150

0.5

150

RELAY-GAIlIER

1

0.7

-.C

0

INDICATOR LIGHTS

1

0.5

C

0.5

INSTRUMENTS

1

0.5

C

0.5

0.2

8.4

POSITION LIGHTS

0.7

0.7

0.7

0.7

0.7

0.7

0.7

0.7

().7

0.7

cn -Inacj 0,

0.5

0.5

0.5

0. 5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5b.

0.5

0.5

0.5

0.5

0.b

0.5

0.5

0.5

0.5

0.1.

0.5

0.5

5.0

5.0

5.0

5.0

5.0

5.0

5.0

5.0

5.0

5.0

5.0

5.0

5.0)

3.2

0.2

0.2

0.2

0.2

0.2

0.2

0.2

0.2

0.2

0.2

0.2.

0.2

I(.4

8.4

8.4

8.4

8.4

8.4

8.4

4.51 4.5

4.3

LANDING LIGHTS

1

8.4

MISC. ELECTRICAL CHECK OUTI RADIO RECEIVER

1 Al.0 4.5

C

4.5

4.5

RADIO TRANSMITTERI

1 10.5b

2.0

10.b

10.5

20.0

20.0

Figure 7-1M. Sa.mENpCe -CSIGNIFILtS CONTINUOUS

,n

0.5

C

i

I ''=:

0.5

C

TOTAL UTILIZATION

uCE5

rý

0.7

1.7

1

c:!

o

0.7

0.2

(.l,

unI~C. oE

, .71 0.

3

5.0

cE

-

10

INSTRUMENT LIGHTS

LAIN

TAKLOQEF AND CLIMB In E o~ 'n I InnE1,n5

z

ACtJL

4.5

4.5

4.5

20.0

4.5

4.5

11).5

10.5

4.5

10.5 10.5

.

7.of Ut

VlOzat LoaT AL

o

and current, resulting in a higher power ongreater torque at thease c or necrod initially in parallel and then switched to series afterstart. a predetermined time delay for completion of the

7/4

7-3.5

7-4.1

Battery maintenance can pose a serious o~pc-

4,5

10.5 lo.b

Igme.I.~m~

MAINTENANCE

20.0

,

ormv

Dc should hos provide for cas) acci. to remove o a startell batteri undwt eriesational conditions. Battery installation is de-scribed in par, 7-7.8.

VOLTAGE REGULATION AND REVERSE CURRENT RELAY,: DC VOLTAGE REGULATION

DC aencratina .sy.titmn in raG!! h .elireonle con.•i.t

rational problem, Becausn of the necessity for periodic addition of water to the cell electrolyte, the battery generally is removed from the aircraft on a

of torwith a serias-field starter, volragea starter/gen regulator, reverse-current relay, overvoltage reluy, field relay, starter relay, and start-control relay.""

scheduled basis, and operational delays thus are cncountered, Operating water loss results from the two natural functions of evaporation and electrolytic dissociation. Except at extremely high temperatures, water loss by evaporation can be considered unimportant. Dissociation occurs at a relatively constant rate, and is a function of voltage, current, and temperature. Higher voltage will increase the overcharge current and gaising rate (Fig. 7-12). Higher temperatures will increase losses by evaporation and will lower the potential at which electrolysis occurs. Higher temperatures also will increase the overcharge current when charging at a constant voltage, The requirements for battery maintenance must be considered in locating the battery compartment. The 7-18

One or more of these components may be supplied to the contractor as Government-furnished equipment (GFE), and thus will establish some arc.% of the design. Military Specifications for a generating system using carbon pile regulators and reverse current relays- include MIL-C-5026, MIL-R-6106, MIL-G6162, MIL-R-6809, MIL-R-9221, MIL-R-25078, and MIL-R-26126. 74.1.1 Voltage Regulaior A voltage regulator designed to incorporate all but the line contactor functions, and having additional functions such as feeder fault protection and field weakeiing for shunt starters, is available. The Military Specification for voltage regulators of the

' '"

AMCP 7065202 Wu

j•

-

-

1

-.

,,

"',

S".

• _-

,

' 0 O

~

-

r.'

4 G

vide, as an integral pArt of the startcr/generator. a tachometer generator that enables the iegu!ator to sense speed, thereby terminating the start at a preis weight •dctcrmined speed. economy. The primary advantage of sht-nt I A line contactor. designed in accordance with I

' I "starting I I

So•

•, : .- 4• --Sf

~

.1 VA

0L

CUNkNI

0F D

......

Io i... ,•

t2

14

1b

IRa

__-i-.-j'o

...... , A

Figure 7.12. Gaem Emitted from Nickel-Cadmium Slitered Plate Cell During Owercarge at 70 0-73F

)

-

*

--

static. type is MIL-R-23761, However, as thi% spcification deals only with voltage regulation and paralleling, the system designer must consider the ftiritionlng of voltage regulation. generator paralicling, field weakening for shunt star'ing. line contactor control, engine stirt control, and protection against reverse current, overvoltagc, overc•citation. startup into shorted bus, and fceder fault. It is recommended that, whenevcr possible, the static type voltage regulator be used. Uhe regulator procurement specification should include all of the foregoing funi.tions. This will economize on weight and installation time since separate componcrts will not be required. The switching action of som- static voltage regulWtrs has " app!rc.ti.n probl'0:s. Sw-chinfrequencies that are kept constant, and at values above 1000 Hz, generally will be above any engine or generator resonant frequencies. This switching action also can produce some radio frequency noise; but if proper switching speeds ard filtering are used, radio noise can be held to a minimum. Locating the reguiators close to the generator also will serve in keeping down radiated and conducted interferences. The use of shunt starters with field weakening is a recent approach to turbine engine starting. The regulators sense the voltage on the starter/generator at the equalizer terminal, and use this variablecurrent voltage by varying the shur, field current in the starter/generator so as to provide a predetermined armature current. Starter/generators with interpole 'indings can develop a shunt field current that can result in no-loaa, overspeed self-destruction, In case of shaft failure, a means must be provided to limit the no-load speed. Some manufacturtrs pro-

MIL-R-6l06 and with the proper rating, can be used to connect the ge .crator to the bus. This contactor also would be used fLr the starter armature current during starting, and by tht; reverse current ovcrvoltage, ovtrexcitation, startup into short, and feeder fault functions in disconnecting the generator from the bus.

A relay in the regulator should be used to de-excite the generator in the event of overvoltage, overexcitation. startup into short, and feeder fault conditions. 7-4.1.2 Reverse Current Pelays The reverse current (cutout) relay is designed to connect and disconnect a generator automatically from the bus in a 28 V DC system. The reverse current relay will close when the generator is producinR 18 to 28 V and is at leest 0.5 V above the bus potential. Depending upon the unit-rating, when the generator voltage drops below bus voltage, the relay will open %ith a given reverse current. These units are available in 100-, 300-. and 600-A continuous ratings. Dependin 3 upon gen•erator -apacity, teversc current relays shall be sized to match the maximum continuous generator output. 7-4.1.3 Overvoltoe Relays Overvoltage relays are used to remove the generator from the bus by tripping the field relay if the gornerotor voltaae exceeds a specified limit. 7-4.2 AC VOLTAGE REGULATION MIL-G-21480 is a teprcscntative Military Specification for AC systems. Highly rclipble control units - which provide voltage regulation, field relay control, contactor control, and ovcrvoltagc, under. voltage, feeder fault, and underfrequency protection are available in solid-state versions. AC generator manufacturers design and build static AC voltage regulators to match their generators. I[he designer must consider the electromagnetic interference requirement, regulator operation environmental conditions, and the qualifitation data before choosing a rtgulator. 7-5

OVERLOAD PROTECTION

7-5.1 GENERAL The primary objectives of overload protection are 7-19

4L

to limit malfun~ction automaidcially toa1 single circuit, and to minimize the danger of &moke and fire not only in the components, but also in the wiring. Overload protection of the equipment should be considered separately 'rem circuit overload protaction. In order to obtain maximum safe use of the equipment, any protection required shall be integral. If the equipment is not requked in order to maintan controlled flight, and maximum equipment use is not necessary, the equipment and circuit protection may be accomplished by the same dzvike, provided that this dual function does not conflict with the basic requirement of protecting the wiring bringing powcr to the equipment. The primary intent of circuit protection is to protect the interconnect wiring and the eq~uipment. All load measuring be oroied itha soft form ofe bruitoproe tltoa. sbepropied wileti onefrtheui of protetv-dvc tecton.Proer elecionof he rotetiv deice should result in the lowest rating that will not openl the circuit inadvertently. A circuit-protection device shouldj be umed at any point in the circuit where the wire size cI cges, un-

that responds to a 1mr~aietic effect rather thar. to the heating effect of the current carried by the bicaker. Magnetic circuit breaktrs normally incorporate time delay so as to avoid nuisanca tripping from current surges ef short durat'on. Although the magnetic circuit breakers ame less affected by adverse environmont, they are not used to the extent that thermal circuit breakers are hecause the trip characteristics of magnetic circuit breatkcra may be affected by their mounting position and vibration.

7-.. ktmm Coto Urcailt Brerkers A rtmote control circuit breaker consists of a contactor whose solenoid circuit is coaitrolled by a current-scrnsitive element, plus a manual-switching adtrpiicin We.Teltrui oenosists of a mntia.ally operated circuit breaker arranged so as to trip wlienev,;r the rrmote sensor trips. The remote ciicuit brcaker can be utilized best for bus feedrs and wiring connected to a single load. Although an approved remote control circuit breaker isntailbMltrySefcton ILC833 I--38 pcrcto i o viaiMltr ý# is being developed for a family of remote control cirth ,nara~m ,n,,L~l f Ii. 'h.,y.,i~~tin wire. Where moethan on .rcuit isfed from a sig cuib~~ circuit-protection device, the protection should be 7-51.3 Currtt Ses sizzd to provide adequate protection for the inA current sensor is used in conjuinction with a condividual circuit. The circuit prolection should be lotactor and a manual-switching or trip-indicating decated as close to the power tsource as is practicabic in vice in order to obtain the actuation of a remute conard .r to minimize unprot-ucted wiring, trol circuit breaker. Tht sensor c'lrrent-sensitive cdo7-5.2 OVERL9AF! PROTECTION DEvICES ment controls the solenoid of the contactor. The tripindicating dt.vii.e often consists of a manually Overload protection devices fall into three. cateope'.-ated circuit breaker arranged so as to trip whengoriess circuit breakers, including remote Orcuit ever the current limit of the sensor is exceeded. When breakers; current scnowrs; and fuses. abeaer crcut t ~n~t acurrint 0nmso W 7-5.21 reilr, ~ *r'p-indicating device, the lowest possible rating should be used in order to obtain an immediate iuidictutedeithr termlly Circit ~eaer2n~aybe Circit myeakes beactatedeiter termlly cation of when the sensor has tripped. The current or magnetiwlly. Both typr4 are covertJ by MIL-C sensor can be, utilized best when there is a need to 5809. control a high-currcrit loai, such as in; motor with a low-current ccntrol circuit, and to keep t1,e high7-51.1A1 ThermaW Circuit Breakers current loads to a minimum length. The actuation of thermal ircuit breakers is depcetdent upon a temperature increase in the sensing 752. Fse 75.4Fu cirmtnt which is produced principelly from the load A fuse relies upon the melting of the cureent. current heatiopg. The thermal element will be affected carrying element in order to open the circuit when an by externtl heating or cooling, and must be derated overload occurs. The four basic fuise types art: noror uprated fihom calibration temperature to allow for mal time delay, very fast-acting, and currentfluctuations in am' ient temjlrature. Tb:.majority of limitiag. the circuit breakers used at tiie present li-'ie are of the tingniai type. Each type of fuse is available in a variety of characteristics so as to meet various circuit require7-5.21.2 MixukascCrcult Broakers ments. For a complete listing of characteristics, mdor Magneti circuit broakers use a trip mechanism to MIL-F-23419 and MIL-F-5372. 7-20

______________AMCP

.7-5Z, OVERLOAD PROTECI'ION APPLICATION

possible methods of compatibility correction or alleviation are disacussed.

Circ,.-it breakers arc preferred to fuses. A fuse must be replaced once its current limit has betrn exceeded, and replazcement with an improper size or type is possible. Circuit breakers shou~d hi. grouped in order 'of function or usage, and should be labeled by function for rapid ieifctn.They should be located in a protective panel, or covereii so as to eliminate the possibility of hazard to personnel or contamination by forcign objects. The placement of circuit breakers in the crew area should be avoided. Or.:y those necessary in order to maintain safe flight should be accessible to the flight crew, as any malfunction must be corrected prior to reinstating the circuit. The installation -.~quirements foi fuses and circuit breakers are detailed in MIL- E-7080.

7-6.2 ACCEPTABILITY REQUIREMENTS Unacceptable equipment responses to EMI levels are exhibited as aural, video, or equipmenit malfunctions. In sonme cases, negative aural response can be acceptable if testing indicates that it does not affeet overall mission capability or ilight safety. EMC tests are required to demonstrate control of the electronic intesrference environment. The detailed requirements for these tests shall be specified in the contracto. s control and test plan. See par. 9-li1 AMCP 706.203, for a discussion of the helicopter system NIMC demionstrationi requirements. ý'n t~sting certoin equipment - for example, ordriaoice - f!.ýr u7 efss`rabic response, it is neccasary to itwwuý that 'the systemr functions within a wide safety mo*'gin. Mib~t.ýfry requirements state that an interferev.-e sigrei impressed upon the most critical point of a subsy'ptem must be at least 6 dB (20 -dB for explosives) below !he level that would cause an undcsirif~c resnonee. Items of equipment that directly flight safety, or thut cause or lead to a uisa:-abort or to failure tc~accomplish a mission, arc determining factcis for the safety margin tests iss indicatcd in MIL-E-6051. 7-.INEFR CES CFCAOS 7-. INEFRNEPCFCTIN Military Specifications require that stfficient tests be made of equipment or weapon systems to insure that they are compatible Spocifications and standards applicable to the design reqaiiremenr~s and test procedures necessary to controll th.- electroniý ;nt~rfPr~nrCC onvironmpnt of sk helicopter are Mll.-B-5087, MILLE-6031, NhilL-I 16165, MIL-STD-454, MIL-STD-461, and hilLSTD-462. Iii gentral, the most current spocification in force will be the controlling factor for EMC quulification.

7-6

ELECT'ROMAGNETIC INTERFERENCE (EMI/YMC:)

GEENALaffc-;t 7-61 Electromagnetic compati')iity (EMC) describes the abliity of aircraft electronic/electrical equipment to perform in its intended c;-erational environracnits without suffering or causing unacceptable d'fgr~kdation as a result of unintentional electromiagnaetic radiation or response, i.e., electromagnetic interferenice (EMI1). EMI is generated by a varying electrical or magnetic field. As a result, almost any device carrying electrical current is a possible source of interference, Likewise, within a weapon sysiemn, cach bubsysterii is a potential victimn of a generated interference. In thec course of EMIC qualification of a weapon system, electrical equipment victim response; to interference sources is defined and evaluated. The solution is to control the EMI by reducing the magnitude of interference, isolating the source, or designing the receptor to be 'ess susceptible to the EMI. To achieve a compatible weapon system, the entire environment, fromn intcrcircwit and intersystem to intrasystem. must be considered by following interference specifications and state-of-the-art engineering designs. The samte results can be achieved by several mear.s; and the best solution depends upon the judgment of the cognizant angir'eer, and upon the budget and time allowance of the particular ap-

?

/

706-202

"~plication.

JThis paragraph outlines the design procedures for

the determinatioai of acceptable EMI levels. In, addi. tion, the identification of sources of interference and

746.4 INTER4FERENCE SOURCES Electromagnetic interference originates from either natural or rnantrade~ sources. Natural sources include atmrospheric, precipitation, corona, and lightning dischargec noise. Natural EMI varies randomly with time, geographical area of operations, and seasonal conditions. Thi- type of interference generally affects a broad frequency range in the low-frequiency band. Manmade sources of EMI are either broadband or narrowband generators, and they must be evaluated and bandied separately. Broadband interference diatributes energy over a wide frequency spectrum, and can be either random 7-21

t

AMCP 706-202 or constant in time and amplitude. Typical broadband generators of EMI are motors, switches, power distribution lines, ground currents, pulse circuits, transistors, and capacitors, Narrowband interference is produced by an oscillatory circuit that contains energy only at the frequency of oscillation or its multiples. The output barmonics of a communication transmitter or its internal oscillators arc typical of narrowband EMi. Spurious outputs of a transmitter or receiver can cover a wide range of frequencies and exhibit the characteristics of broadband noise; however, the energy distribution is

defined sharply.

4. Inierference time coincidence, i.e., signal presentation during timea of receptor susceptibility. The complexity of the subsystem, and the number and magnitude of the internal interfcrence sources, determine the choice of protoctive design approachcs. Basic appwoaches to interference reduction within tne helicopter or subsystem include: i. Dftiagn of inherently interference-free cornponents 2. Equipment isolation 3. Cable routing 4. Source suppression

5. Signal point containment and suppression.

Inherent interferences unique to the helicopter can

arise from sources such as the rotating members of the engine, drive shaft, and main and tail rotors.

t

In small- and medium-sized helicopters, radio/ radar operation frequently is hampered seriously by a phenomenon called rotor modulation which creates problems especially in VOR/ILS, ADF, and some communication systems. Rotor modulation interferences arise due to the chopping or reflection of the RF signal by the main rotor. The rotor speed and the nurmber of rotor blades combine to pass a givcn point

7-.5.1 interferencC-free Compoeats All electrical systems shall meet the limits imposed

by the applicable equipment spec-ification, such as MIL-STD-461. These speci!i'zations primarily are concerned with radiation, and with susceptibility to radiation- or conduction-propagated broadband and narrowband interference. Compliance with these specifications represents maximum state-of-the-art interference control. However, the specifications are oroad and do not necemsariiV soive ihe inic.i-eren:ce

resulting in the modulation of the arriving RF signal These distortion perturbat,,ns (amplitudes, cancellations, or harmonics) can set up interference patterns that create navigation system noise, error, and needle oscillation. The interference caa become critical when integrated flight control systems are used, resulting in helicopter oscillation. The expanding use of helicopters in a variety of ei-

problems arising in all systems. If individual borderline component interference sources are not eliminated, compliance with specification limits does not insure that EMC problems will not develop wben the total system degrades from specification limits. 4-6.5.2 Equipment Isolation and Cable Routing Many EMC problems arr. oolved by positioniing eetrncqup ntoruigcbesshthth) pick up or radiate minimal interfereacc. Lccation

considered previously. These interference effects can

and orientation are two importan'.t parameters in pre-

downgrade seriously, or even prievent, a particular

mission capability. Some interference probles arise

__ ._ .. i:ola.i. attnuates with distance. antenna location and oricn-

from atmospheric field charging potentials, precipi(electation charging, corona discharge phenomenon trons accelerated by a strong electrical field around a sharp point), or triboelectric charging potentials (frictional charging as a result of dissimilar material contact)Of these sources, probably the most noticeable effect for EMC qualification will be produced by the tniboelectric charging of helicopter rotating members (engine, transmission, drive shaft, and rotors).

tetion can prevent or reduce EMI. Simple shielding of cables is not always effective, due to the magnitude of interfering signals. In such instances, isolation of equipment cables is necessary. Scparation of high-level from low-leve, cables may be required, depending upon design and space allowances. Signal wires and primary power cabIls may rnquirm separate routing even when terminating at a single connector. If interference is a result of equipmea.t location or cable routing, the following areas should be investi-

7-6.5 INTERFERENCE SUPPRESSION EMI within a subsystem may be divided into four

gated: I. Power and control wiring run separately from

categories: 1. Device signal interference emi3sions 2. Device susceptibility to such signels 3. Transmission path of interfering signals (solid or wave)

signal-carrying wires 2. Audio frequency wir-s run separately from wirec of higher frequency 3. Provisions madt for the right-angle crossing of sensitive circuit cables

7-22

.

S'qk 4 Pro.: wit types used 5. itaxim.urn -patiai separation of antennas or intet fetrencproduc'iog cables 6. Cr]tOaling of nonintcrferinig equipment away

)

ferrous materials will provide shielding above audio

frequencies (electrical fields).

Shielding used to contain interference is dependent primarily upon the attenuation (absorption)

from ý;nowvn interference sources.

of the shield. Reflection loss becomes an properties important consideration for exclusion of interfering

Satirce Suppresslon and SRsceptibity Reduction After using physical isolation aid cable routing to the maximum extent, additional techniques for EMI source and susceptibility reduction include: L. Grounding and bonding 2. Cable and cquipment shielding 3. Filtering, Source suppression is the application of appropriate bypassing, decoupling, or filtering at the source of interference or at a point of maximum susceptibility.

signals. Discontinuities in a shielded enclosure can: provide an entry/exit path for EMI radiation. Ventilation openings, panel meters, access c*vers, dial shafts, or switches are possible EMI containment problem areas. Interference coupling of electronic subsystems can be reduced by careful selection of interconnecting cables. Types of interconnecting cables available to the designer include unshielded wire, twisted pair, shielded wire (single or double), twisted shielded pair, aid coaxial (single or multiple shield). The selection of interconnecting cables to reduce interference coupling and audio crosstalk will be, dependent upon physical isolation of the operating frequency range, and the power and susceptibility level&. In general. a shielded wire provides protection against eietricatl fiellds, whilz the twisted pair reduc-s susceptibility to magnetic fields. To achieve maximum EMI shielding from cnclosures and shielded cables, it is necessary to terminatc them cffectively to the helicopter unipotential ground plane. Both multipoint and single-point ground systems provide certain design features. Single-point grounding (floating shield) may provide the best approach where the possibility of interference coupling with sensitive low-frequency circuits is a matter of concern. When a shielded cable, in

74.5.3

.

MqP 706-202

7-6.5.3.1 Groundlig and Bonding A fundamental requirement for helicopters is the establishment of a well-bonded, low-impedance t all ,,r-...ll! A %n;t, ground pOmni exte•n•d• potential ground plane prevents EMC problems resuiting from unequal ground potentials and ground loop currents, and reduces the possibility of equipment transmitting or rmceiving undesired energy while insuring that shield and filter applications are effective, Bonding refers to the method in which various subsystems or structures are conneced or integrated electrically and mechanically. Bonding avoids the development of electrical potentials between adjacent metallic parts, and provides hl,mogenous flow of radio frequency currents between subsysiteis and structures. MIL-B-5087 provides detail requirements for all bonding aspects of airbornt systems. 7.6.5.3.2 Stdelding A major area of practic.l EMI suppression involves the application of component or cable shielding. Effective use of shielding requires investigation of the interference signals, and of the nature of metallic sl.ielding. The question of whether the source or rt ceptor is prvented from radiating or receiving undesited signals deserves equel attention. Metallic shielding is dependent upon the ;nterfering sgnal component, e.g., the electrical or magnctic field. The lowest frequency for which a desired "shieldingis required normally determines the type of shielding material. High-permeability materials can be used to improve shielding effcctiveness for low-frequency, lowimpedance magnttic fields. Aluminum, copper, or

a sensitive orircit, is ground-d at both ends for the

return circuits, power frequencies in the ground plane can induce audio frequency interference in the signal wires. When electronic and electrical equipment is distributed over large areas, experience has shown that multipoint grounding is superior for RF frequenries. Multipoint grounding involves shield grounding at both ends of all cables, and at all immediate points where the cable runs through equipment. A ,plikation of proper shielding techniques for interference alleviation should be performed in the following areas: I. The radiation source or sensitive component should be installed in a properly bonded metallic housing with limited openings. 2. The magnetic field should be directed away from sensitive components or wiring by use of lowreluctance, high-permeability matarial. 3. Twisted, shielded, or shielded and twisted cable 7-23

should be used for AC and DC power ci,-:uits in order to prevent coupling of super-imposed EMI noise and transients, 4. Two conductor-twisted and -shielded cables should be used for DC signal, control, and audio circuits. Single-point grounding is required. 5. Single- or mustiplc-shield coaxial uable shoull be used for RF circuits. Multipoint grounding is required. 6. Continuity of shielded enclosures is necessary. 7. Shields should be routed through connectors. 8. Minimum-length grourtd returns should be used, and shield insulation from structural members should be insured.

stallation) in the airframe. This includes electronic ccmponents, clectrical relays, electrical power generators, wires, coaxial cabh:s, junction boxes, test connectors, etc., but does not include aircrew control panels and instrument panels. Electrical system installation should be in accordance with MIL-E25499, MIL-E-7080, and as described subsequently.

..

7-.5.3.3 Filtess Filters are used at the outputs of EMI generating sources in order to prevent EMI signal (broadband or narrowband) interference coupling paths. Types of filters utilized for EMI containment ani attenuation include low-pass, high-pass, and band-pass filters, as well as bypass and feedthrough capacitors. Basic filter pao ameters include capacitance, induciMC.x, and mb-aiax. F•cit paraainte, u,.pI.mhwas filtering action by a differant method; i.e., capacitance by short-circuiting, -inductance by opencircuiting, and resistance by dissipation. i-ltcrs should suppress only the :interfering signals.; However, the filter may have an effect upon desired currents necessary to the operation of the equipment. Therefore, an understanding of insertion loss is important to filter applications. In the application of bypass capacitors, the lead length from the capacitor to ground becomes an important factor. Self resonance nullifies the effectiveness of the filter for signals at 1rejueneieS equaito, or .. . greater than, the resonant frequency. Filter containment of EMI can be effectiv; only if the source can he. shielded and isolated from olher internul circuitry, thus preventing the interference from being coupled into other wiring or circuitry within a subsystem. Such coupling may conduct spurious energy to external wiring, or radiate directly from other parts of the unit. Proper bonding must be used in order to prevent interference currents in the ground circuit from shunting the filter element.

7-7 7-7.1

ELECTRICAL SYSTEM INSTALLATION GENERAL

Electrical system instaliatior refers to the installatior. of electrical and electronic oquipmelt (equipment installation) and wire bandle (electri~cal in7-24

7-7.2 EQUIPMENT INSTALLATION During the design of equipment installations, maitainability, reliability, and producibility must be considered from design concept to the production hardware phase. Close attention should be given to the servicing problems tha: might arise with each particular installation, It is not likcly that all electronic components can be made immediately accessible. The service reliability of each must be considered during design of the installation. Factors such as electronic alignment after installation and acccsibility to test points must be considered. If equipment is installed in rows, front row con-ip'nents must be capable of being removed quicHy to provide accessibility tc rear mounted components. Equipment-mounting hardware should consist of not less 'han Number 10 screws, except where vibration isolators are used, in which case the box mounting screws should be no smaller than Namber 10, with the isolator multiple mounting screws no smaller than Number 8. Care must be taken to insure that mounting screws are not hidden behind flanges and protruding portions of neighboring boxes. For easy accessibility, the straight-in approach should be provided for all mounting hardware. Equipment installations involving the placement of electrical receptacles facing bulkheads or other obstructions must allow sufficient room for installation of the wire bundle with a bend radius in accordance .with MAL-W-5088. as well as room tc engage and disengage electrical connectors without darmaging the wires. If possible, electrical terminals on boxes should permit the use of a ratchet-drive socket wrench for wiring installation and removal. Junction boxes must be designed so as to facilitate maintenance and troubleshooting. Access to internal compoaicrts must be such as to permit easy replacement. Th,. locations of internal components must be ider.tified by permanently attached decals. Foreignobjct protective covers must be provided on all junction boxes. On all nonsealed boxes, drain h,'les must be incorporated at the lowest point. Wiring musi be installed neatly, and numbered or color-coded for ease of maintenance, Relays, resistors, small transformers, etc., must be

.

/

%A

-

45 dog

'(A)

(B) Figure 7-13.

-

Permihsbl CIamP Ddumsmiiea

grouped functionally in panels similar to relay panels, tubing, but shoule .ot be covered by the braid or cxtruded outer jacket of the bundle. All components must be located and i4entified by The primary wirv bundle clamps should be of an means of a decal permanently attached to the panel. environmentally compatible type. Nylon clamps are Power contactors must be installed so that the conpermissible in low-tempsature, low-vibration, easily tactor case or box is isolated from the airframe strucaccessible arm. Plastic clamps are not to br use for ture. wire bundle support in arse where a damp failure "Alljunction boxes and panels should have power 91 ,tsam ^ ,har 6 w X;i. k, Air eho.f ci .... ., c,-iuOf and c.tassis grounds emanaung from ui or to interfere with controls. The preferred orienthe electrical connectors, tation of all wire bundle clamps is with the bell (loop) down. The bell should not be turned upward if the 7-73 ELECTRICAL WIRE BUNDLES wire bundle weight threat 'ns to deform the damp. Clamps of the MS 21919 type may be deformed in Basically, there are two types of wire harnesses ordwr to meet special Installation problems by flattenallowable: ing the dlamp bell, as in Fig. 7-13, to a height no 1kw 1. Open-wire bundles, where individual wires are than 3/4 of the original bell height. The mounting tied in bundles and routed through the airframe ears may be bent, but not more than 45 d4, as shown 2. High-density bundles, where an abrasionin Fig. 7-13. resistarm szovering ic braided, extruded, etc., over the entire bundle.

In either case the wire best suited for the particular application must be used; and, when open-wire bundles are used, the wires shall have markings in accordance with MIL-W-5088 and the bundles shall be tied at 3- to 8-in. intervals. Lacing shall be comsatible with the operating environment of the helicopter, Where high density bunweas are used, the bundles must be taped at 8-in. intervals with a thin layer of Teflon tape. An outer abrasion-resistant covering must be braided or extruded over the wire bundle. Tape is not acceptable as an abrasion-res•atant covering except on repair areas or at the ends of a bundle. Tape must never be used as primary insulation. Repairs to high-density bundles should be made by routing a wire external to the abrasionresistant covering. The external wire must have an abrasion resistant covering. Splices are to be coverd ith an ab.asion-resistant material, such as Teflon

! . zaI'q; -7. ZME V NAE • Terminal strips should bN MS 27212 or MIL-T81714 with MS 18029 covers. Torminal strips shall be installed as shown in Fig. 7-14, with the mounting holes isolated, for example, by filling with MIL-A46146 Type I sealant to prevent short circuits to ground. MS 25227 insulating strips may be used in lieu of potting; however, an additional nut must be installed between the insulating strip and the bottom terminal to that there is no resilient material in cornpression with the terminals. A maximum of four terminals shall be used on one atud. MS '5266 boa bars may be uscd between studs to interconnect terminals. When terminals are exposed to the weather - such as in wheel wells - terminals and studs shall be brushed with phenolic resin varnish. Thf wire bundle *Wl be tied to a terminal at each breakout. Ther sell be at least one wire identi7-25

SELF-LOCKING NUT LOCK WASHER FLAT WASHER -.

A---

_44A

MIL-A-46146 TYPE I SEALANT Figure 7-14.

Terminal Strip Installation

fication number visible on each wire without cuttin3 ties. Lacing (or tying) shall be done with single ties. Continuous lacing shall be permitted only in junction boxes and panels. The end studs used for attaching the MS 18029 terminal covers cannot be used for electrical purposes. If two electrical terminals with mounting hardware arc m placed on the end studs, the self-locking feature will not engage in the terminal cover nuts. See par. 7-8.2.1 for a further discussion of terminal blocki;

7-7-6 DOOR HINGE WIRE BUNDLE ROUTING Electrical components mounted on access doors v,ill require routing the wirr? bundles over the door hinges. The wire bundles shall be routed so that they twist instead of bend, i.e., the bundle shall be routed parallel to the hinge for a distance sufficient to allow he bundle to twist. Consideration should be given to using Teflon-cushioned clamps at the twist points to provide added bundle mobility. Added abrasion resistance at the hinge, in the form of vinyl . , T.n tubing mnay be required.

7-7.5

Wire bundles that are exposed to weather and when doors are opened during flight, or abrasion during ground servicing, shall be protected by extra cove.ring (such as braiding or tubing). Weatherexposed braiding shall extend into the connector back shell clamp, but, because of the wattr-wicking properties of the braid, should not extend into potting or connector waterproofing.

ENGINE COMPARTMENT WIRING

The two major installation hazards encountered in 4ngine compartment environments are heat and viI ration. Special attention should be paid to the highvibration environments of engine enclosures. Wire gage shall be a miinimum of 20 in order to reduce strand fatigue breakage. Wire bundle clamps shall be spaced in close proximity so as to prevent wire vibration between clamps and possibie resultant breakage. Crimp-type contacts shall be used in order to Sclumninate strand vibration breakage due to solder capillary action, Wire bundles in low-temperature areas (200°C or lower) of the engine compartment may be in accordance with par. 7-7.3; in higher-temperature areas and on the engine itself, open wire bundles of wire rated at 260*C shall be used. Particular care shall be taken to route all wire bundles away from sharp edges, and around equipment in the engine area to allow extra room for vibration and for structural expansion and contraction due to ambient temperatures and engine thrust. Wire bundles shall be routed and clamped well out of the way for engine change, and design shall take into consideration the use of any necessary installation/removal ground-handling tools. Fire detector elements ghllbe routed, and securely clamped into position, to eliminate crush possibilities during engine change. 7-26

7-7.7 WIRING TO MO I, WNG COMPONENTS Special attention is required when it is necessary t route wiring bundles to components such as actuotors, missile launchets, or electronic components that move during use or storage. These bundles usually flex a number of times and are critical in their operation. The installation should be designed as follows: I. The wire bundle shel bZ clamped firmly to the moving component so that no movement of the wire takes place at the connector or terminal. 2. The wire bundle shall not be under tension at any point in the movement of the equipment. 3. The wire bundle shall be clamped firmly ti the fixed structure at a position whek e if there is any motion, the wires will twist and not bend. 4. The attach point of the fixed structure must be, whenever possiblen at the center of the arc formed by the moving equipment. 5. If the fiued point car not be at th: center of the

GRýOUJNDING PAD

PRIMARY STRUCTURE LOCK WASHER NOTE: BOND ALL PARTS PER MIL-B-5087 Figure 7-15.

) ...

Typical Connection to Grondilag Pad

74

may be required on the slack wire bundle. Vinyl

fined in only one appropriate specification. Environ-

sleeving is not to be used as a substitute for good engineering. Protect:ye tubing should not ride on sharp edges of structure, 7-7.8

BATIFERY INSTALLATION

Batteries shall be installed so that they are readily accessible from the outside of the helicopter. The aircraft connector shall be of the quick-disconnect MNI 25182 type in accordance with MIL-C-18148, and shall be accessible without moving any equipment or reaching around any obstruction. The battery compartment must be located in such an area that battery gas and fumes will not enter the

cockpit or cabin. The battery compartment shall be

)

COMPONEI TS

moving arc, a loop must be made to take up the slack in the wiring. This loop :nust be of sufficient length to insure that the wire bundle is never under noticeable tension. This loop must be self-supporting and selfforming. The seit-supporting feature can he assisted by a preformed spring steel wire woven in, or attached to. the wire bundle. 6. Attention shall be paid to chafing of the wiring, Added protection, such as vinyl or Teflon tubing,

painted with a material resistant to the electrolyte used in the battery. There shall be no oxygen, hydraulic, or flammable lines in the battery compartment. The batter, cables shall be clamped and protected against chafing during installation and removal of the battery. The battery ground cable shall be attached to primary structure that is heavy enough to carry shortcircuit current without damage. A grounding pad, as cshOwn in Fig. 7-15, may be used to increase electrical current capacity.

7-8.1

WIRE

The choice of wire should take into consideration .. cl'z--. -h . .. f t. i,.

not O,,t ,

also the environment in which the wire must operate. The electrical requirements cap be satisfied by the wire current capability; however, the environmental requirement may be compatitzie with the wire *.'mental compatibility will vary depending upon the type of insulating material used. The designer s&Wll assure that the finished diameter of the wire eected

is compatible with the wire scaling ranges of the connector used and compatible with the connector insertion/extraction tool.

.

7-81.1 Wire Imulatihg Materials 7-8.1.1.1 Polyethylew Polyethylene is a commonly used dielectrical material. It is excellent for high-froquency applications. Howev,:r, because of its physical properties, it has definite limitations s an Insulating

material. Polyethylene pgssesis

low abrasion rm-

sistance; the maximum 3afe operating temperature is only 80*C, and it will burn freely in the preence of an open flame. 7-8.1.1.2 Polynylclilorlide Polyvinylchloride (PVC) has physical properties that surpass those of the basic polyethylene. It poswsess greattx abras-an resistance, higher operating temperature limitations, and increased resistance to flame. However, the molecular imbalance of PVC precludes its use at high frequencies, although it 7-27

"

AMCP 022( is excellent in low-frequency applications where reI"

*

sistance to moisture, Rlame, oil, and many acids and alkainesis ~numerous mporan 74.1.3 ~ ~ ,,., ~scription Fluorin~ated ethylene propylene (FEP) demonstrates, excellent electrica: stability over a temperature range of -65* to +230*C, and is suitable for ultrahigh-frequency applications, 74.1.1.4 Polychlorotrlfluonsethyliea eolychloiotrifluoroethylcne, more commonly known as KEL-F, combines many of the advantages of Teflon with a superior resistance to abrasion, thus enabingit b use oasa tin-alle inulaion without any outer covering or mechanical pro. tection. This material is rated for continuous operationthrughthetemeratre ang of-65 to +15thog0tetmprtuerngCf-.*t + 1500C.draulic 7-8.1.1.5 Polyliexamethylene-adipaisldc Polyhexamethylene. adipamide is a readily exru dabe btte plyiid, kownas y is amiy nme dabof noylo.Bimid, betrkof n assbc~ y poor eamlectical of nlon 11"uscof ts clafvel por elctrcal characteristics, it rarely is used as a primary in' asulation on wire. However, it makes an excellent outer coverings wt icn applied over vinyl insulation. Ex truded nylon jackets are tough cnd resistant to &brasion and oil, and have a tendency to increase the temperature sawbility of the piimary insulation,

The Military Specifications for aimrcrft wire a--e too to cover in detail. However, a brief doof some of the more commonly used typos of wire and of the specifications defining tham is giver. to assist in selecting the specification that satsfie& the general requirements. MIL-W-5086 covers PVC-insulated, singleconductor hookup and interconnecting electrical wires made with tin-coated or silver-coated conductors of copper or copper alloy. PVC insulation may be used alone or in combination with outer insulating or protective materials. It is a good general purpose wire, and is available in voltage ratings from 600 to 3000 V and a temperature rmpg of -55* to + I 109C. The wire construction of this specification contains nylon jackets for increased mechanical toughness and resistance to fuels, solvents, and hyfluids. MIL-C-7078 covers single-conductor and multiconductor shielded wire. The basic wire in this specification is MIL-W-508 and MIL-W-81381. MI1L-iW 1818 covers wifVe designed for internal wiring of meters, panels, and electrical and electronic equipment, and requires that such wins, have mini. mum size and weight consistent with service requiremerits. The temperature rating of wire included in this specification ranges to 260*C, with potential ratings of 250 ^,o 3000 V. This wire is primarily a hookup wire, but it may be uted for wiring elcetronic Teraflorathyeneequipment in protected areas of the aircraft. 7-8..1. 74.11.6 etrahaorcthyeeM IL-C-22759 covers fluorocarbon-insulated. Tetrafluoroethylene (TFE), better known as single-conductor electric wire made with tin-coated,

Teflon, is an excellent electrical balance ar~d. applications. TFE offers exceptional electrical, chemcal proertes an thrma ot vaiablein ny othe wiemaeria. isultio TE isultio is rate fo cotinousopertio at200C. ut rmais feilatcygnctemperatures. 74.1.1.7 Dimediyl-slloxane Polymer Better known as silicone rubber, dimethyl-siloxane polymer is finding widespread application as a wire insulation because of its good high-temperature characteristics and low-temperature flexibility. It will withstand 2000C continuously, and can withstand as much as 300"'C for short intervals. However, iii the presence of flame, silicone rubber will burn to a nonconductive ash, which, if held in place, could function as an emergency insulator. Its abrasion resistance is *improved greatly by the addition of a saturated glass braid. Unlike vinyls, polyethylene, and nylon, silicone rubber is a therinosetting plastic. *

74.1.2 Military Wire Spelfkcadioa

7-28

silver-coated, or nickel-coated conductors of copper wires may be polytetrafiuoro~ethyloen, fluorinated thyenepropylene (FEP), or polyvinylidene fluoride. Thefluorocarbon may be used alone, on in cornbnaton ithother insulation materials. This wire is available in a temperature range of 2000C to 2600C, arnd vlaeratings of 600 to 1000 V. MIL-W-7072 covers low-tension, insulated, singleconductor, aluminum wire for aircraft electrical power distribution systems. Aluminum wire usually is used where an appreciable weight saving can be re~alized. MIL-W-81044 covers a variety of construction suitable for airframe and electronic hook-up wire, ineluding flght, medium, and heavy wall insulation thickness and tin- and silverpl'ted-copper conductors. These wires are rated to 5W0 V over a temperature range Of - 53 to +I150C. The insulation consists of crosslinked polyvinylidene fluoride. Improvvd thermal stability is realized through mole-

'r

AMCP 706-202 cular crosslinking of both materi.Js by the high,nergy electronic beam process. These consiructions provide significant space and weight savings while retSining excellent abrasion resistance, MIL-W-25038 covers single wire for electrical use under short-time emergency conditions involving exposure to flame and temperatures of up to 2000*F. This wire is intended for use in circuits where it is ,necessary to maintain the electrical irtegrity of the insulated conductor for 5 min in a 2000*F flame with the operating potential not exceeding 125 V.

Fittings cover a broad area, and include any fixture attaching to a wire. Two basic fittings are terminal strips and connectors. 7-8.2.1 Termhna Strips strips

requirements. Thus, the selection of a connector for a specific application will involve a compromise. M IL-C-5015 covers circular electrical connectors with solder or removable crimp contacts, and accessorics such as protective covers, storage receptacles, strain relief clamps, and potting molds. These connectors arc for use in electronic, electrical power, and control c-ircuits. They have threaded couplings, and may require safety wiring in order to eliminate inadvertent decoupling in high-vibration areas. MIL-C-26482 covers environmental-resistance, quick-disconnect, miniature electrical connectors with solder or removable crimp contacts and accessories. Thes connectors ha'e bayonet couplings and do not require safety wire. MIL-C-83723 covers an environmental-re3isting family of miniature, circular, electrical connectors. These ;:onnectors may have threaded or bayonet

wires. Terminal of two or more ment for a junction ued aya diconcct inapstris e aso strips also may be used as disconrects in ap-

gd inr c MIL-C-28748 covers rectangular rack and panel and electrical connectors with nonremovable solder

"7-8.2 FITTINGS

Termina! strips are used where there is a require-

plications where :'t is impractical to use a connector, or to simplify assembly and maintenance pro-

couplings.

cnat n removable eoal rm contacts. otcs crimp contacts and

or, ts. ly sb nMIL-C-39012 covers the general requirements for connectors used with fiexibie cofrequency radio or 27212 MS is the etrip terminal The standard The temiablrpesth.S ~ ~ ~ ~ ~ ~712 ~ ~ xastndr MIL-T-81714 which consists of a series of threaded axial RF cable. studs retained in a plastic insulating strip. Each terThe designer shall make every effort to select only connectors that provide common termination minal stud will accommodate a maximum of four termethods; i.e., common contacts, common back hardminals; however, a bus bar may be used between ware, and commor, assembly methods and tools wiles four studs in order to allow for more than using MIL-STD-1353 as a guide. having a common junction. The new NAS standard terminal strip, which con-

sists of series of modules rctainei between mounting rails, offers maity advantages over the old style MS terminal strip. MIL-T-81714 covers environmental feedthrough and noni'mr.dilzrough tef ii,,l srifi ips. Fror

7-9

new designs qualified parts shall be in accordance with MIL-T-81714. This type of unit is similar to an •loctrical connector in concept in that it uses a crimp pin, and in insertion-extraction tool for installing the wires, Each terminal strip requirement must be evaluated individually in order to determine which of the type can be used best.

The proper functioning of electronic systems is taking on increased importance in mission effective. ness and flight saf.ty with the development of electronically controlled, automatic flight and engine controls. Thus, the common occurrence of total electrical system failure ft'om lightning strikes is no longer acceptable and a higher degree of static electricity and lightning prottction must be provided for the helicopter in order to ussure reliable, safe, and effective operation over its operational lifetime. One lightning strike can ft expected to occur on a helicopter approximately e-.cry 2500 flight hr (Ref.l), depending upon aircraf zone of operation, mission. normal flight altitudes, susceptibility, etc. Minor to serious structural damage cav result in cases where protection is not provided. New materials, such as pbolyurethane paints. have many advantages relative to corrosion protection;

"74.2.2 Conmetors

)

The ideal situation, as far as reliability is concerned, is to have continuous conductors throughout the entire circuit. However, this usually is not pussi•': interconnects must be added to facilitate assembly and maintenance. The designer must select the connector that best combines high-performance factors with capabilities for meeting env'ronmental

LIGHTNINC( AND STATIC ELECTRICI'TY 7-9.1 GENERAL

7-29

AMCP 706-202 but their exeleknt dielvctri: characteristics also can introduce serious static electricity problems. The high dielectric strength of the painted surface permits the buildup of 5000 to 50,000 V from friction charging of the surface, which may be followed by puncture of the base metal and accompanied by an energy releast in tens of joules. This can cause precipitat'on static or streamer radio interference, and - if the paint is covering an elcctiical component. such as an engine inlet heating grid - also can result in a short circuit of the element. This often is followed by burnup, as a result of energizing of the initial spark by the power system, with resultant major damage. Possible internal problems with high-quality dielectrics include the charging of fluid lines from the liquid flow and the charging of painted internal fuel tank walls from spray electrification or sloshing. 7-9.2 LIGHTNING F OTECTION FOR ELECTRONIC SUBSYSTEMS

The designer of lightning protection for helicopter electronic subsystems should make maximum use of the metallic frame and skin for shielding purposes. Specific lightning protection, or lightning-resistant designs, should be provided at the major lightning entry points. These include main rotor and tail rotor blades, antennas, nrvigation lights, pitot-static tubes, active electrical discharger probe heads, and any other electrical -omponents exposed on the exterior of the helicoptcr. In addition, because of the generally reduced shielding of helicopter frames and skins (compared with fixed-wing aircraft), greater considerations must be givan to magnetic and electrical field penetrations into the vehicle interior. Where all other factors are roughly equal, it i' advisable to use mechanical primary flight controls as engine and rotor controls and to use the electronic systems prin-arily for trim or management controls. Electronic surge suppressors of various types, such as gas or zener diodes and simple capacitors, may be used on critical circuits for suppressing the residual voltage surgc (which can penetrate despite the external lightning protection design), particularly if the electronic systems require very-low-voltage protection. In summary, the preventive design approaches are: I. Principal lightning protection efforts should be directed toward blocking electromagnetic energy entrance through electroma.gnetic windows such as navigation lights and antennas, S. Use of electronic systems for primary flight controls should be avoided. Use should extend only to trim or management. 3. Surge suppressors should be used where 7-30

required, eith(r because of large surge voltages that cannot be reduced at the entry point or for lowsignal-level circuits that require low-level protective devices. 4. Simple lightning test facilities should be used to permit quick evaluation of component performance. Untested lightning protection designs often have proved to be not only ineffective, but sometimcs more dangerous than the components they were intended to protect. Lightning protection through geometrical configuration control of external components, such as antennas and navigation lights, has proven to be one of the most effective methods of preventing lightning penetration into the aircraft. For example, tests of navigation and cofision light designs have shown that a 1-in.change in a cover screw position can reduce the resultant lightning damage from total destruction of the element, with major energy penetrations into the vehicle interior, to negligible physi-

cal damage resulting in voltage pulse amplitude re4;uctions to a few hundred volts. Thus, geometrical control of all external components for lightning proivt;6iun purpubcz gcncraiiv is itc most economicai approach, in terms of weight and cost. Typical entry points requiring protection design effort are shown in Fig. 7-16. Earlier HF and UHF antennas of the voltage-fed type constituted one of "he princip~t electromagne',windows through which lightning energy could enter the vehicle interior. To offset a possible total electrical system loss, these units often can be replaced with shunt-fed antennas, which are inherentl) grounded designs in which the lightning energy essentially is channeled into the external vehicle skin, with only residual high-,oltage, low-energy pulses entering the electronic systems. HF lightning arresters are available commercially for HF antennas. ttid their effectiveness in preventing bothi structural and radio equipment damage has been demonstrated in their use on commercial jet airliners during millions of flight hours. Other external components, such as pitot-static heads and active discharger probe heads, require typical electronic system protection approaches. The pitot-static heads can be protected effectively by conventional electrical system protective devices such as zener diodes or gas diodes; however, the high-voltage active discharge probe heads require more extensive protection development because of high operating voltage levels. For electrical surge suppressi n, mac:y types of devices are available commercially - including zener diodes, gas tubes, simple capacitors, spark gaps, and

AMM 7W

LIGHTNING DISCHARGE

PITOT TUBE

( NCOLLISION LIGHT

(

,

•-

....

•"

I •,q-•UGHr(~ COLLISION

7.

EM FIELD -

I

I-

'

\

\

\ANTENNA

AFT NAVIGATIO(SI LIH

, • ANTENN4A

ELECTROMAGNETIC FIELD PENETRATION THIOUGH PLASTIC COVERS

1

) Figure 7-16.

Typical Lightlng Electrical Circidt Entry Points

silicon controlled rectifiers (SCR). The prinCipal problem in their application ib the selection of the right device, or combination of devices, for the particular equipment being protected. As an example, for antenna front ends, semiconductor devices have a major sh,'wtcorming, th.;y introduce cross-modulation through their inherently noniinear transfer characterutici. Simple gas tubes present only a light additional capacitive load on the front end, and thus provide more suitable protection for this application. For other types of comnponents - such as ciectronic contro! systems, where nonlinearity may not be as important as is obtaining sufficient lowvoltage protection levels - zener diode protection devices may be more suitable. STATIC ELECTRICITY The requirements for control precipitation charging are much mcre severe for htelicopters than for fixed-wing aircraft because of the cargo-handling requirements. Potentials that would be acceptable on fixed-wing aircraft -- 20,000 to 30,000 V, which is well below the radio noise threshold of the vehicle can represent a serious shock hazard to ground personnel unloading cargo from helicopters, and posbly carl cause ignition of ordnance or fuels. Several

7-9.3

general approaches have been suggetd and carried to various degrees of development, including use of active dischargers in which the clectrical field from the aircraft is measured and an opposite charge is ap plied to the vehicle, the use of passive wick-type discharge devices at the blade tips, and the use of conductors hanging from the helicopter to the ground to discharge the vehicle before the ground crew contacts the load. The active dischargers suffer from several disadvantages, including indicating a charge in external electrical crosaficids when using single-head field meters when no charge actually is present on the helicopter and thereby charging the vehicle with the protection device. This can be prevented by using dual field meters, one above and one below the vehicle. However, space-chargc shielding of the field meter sensing head can occur from a recirculating charge during hover. It generally is acknowledged that the use of active dischariers, in spite of the shortcomnings, is advisable, rarticularly when ground handling is frequent. The passive wick dischargers located on the blade tips have the advantage of simplicity, but suffer from the fact that substantial potential is required an the vehicle before they begpn discharging, i.e., they do not 7-31

,

AMP706-202 bring the vehicle potential down to zero. This still permits sufficient potential to give shocks to the ground-handling crew. The technique of using a conducting cord from the vehicle to the ground, and permitting it to contact the ground before the ground crew handles the load, has the disadvantage of the cord being whipped by helicopter downwash. and will not nocesarily hold the vehicle pottntial down continuou.ly while the ground crew is in con tact with the load. The other major problem with external static clectricity on helicopters is radio interference. The complexity of the problem is caused by: the variety of charge-generating mechanisms, of nois-generating mechanisms, and of coupling modes into the communication systems; the difficulty in separating the effects from internally generated equipment interfcrence; and the differences of effects upon different types of equipment. The basic method of controlling radio interferenceincludes: I. Avoidance of all electrically floating external sections on the aircraft 2. Use of some type of active or passive discharge.- in order to reduce the potentials, on the vehicle under friction electrification conditions 3. Location of antennas in areas where the DC electrical fields are minimized under thunderstorm crossfield conditions 4. Use of radio-interite,-,cc-resi,"ant antennas 5. Coating of all external diclectric surfac-s subject to particle impingement with resistive paint&so as to prevent streamer interference, particularly over plastic sections where the interference coupling is most evere. In addition 'o the external problem, which is complicated by the difficulty of proper identification of

the interference source. internal static electricity

7-32

problems involve the fact that helicopiers often are engineered by designers who posse little knowledge of the hazards posed by electrical interference of fuel systems. As an example, plastic tubing often is considered for fuel jettison tubes. Friction electrification of the plastic surfaces of the= tubes can ignite the fuel vapors, particularly when the fuel tanks and jettison tubes are nearly empty. As a solution to this problem, it has been saggested that all dielectrics with a resistivity )f higher than 10' ohm-cm be carefully considered for aircr-,ft use. Thus, the use of such materials would be permitted, but freedom from static electricity hazards would have to be assured for each specific installation. LIGHTNING AND STATIC ELECTRICITY SPECIFICATIONS There are a number of Military Specifications containing reierences to surges and protections. MILSTD-704 defines the accep:able limits of transients on electrical power systems. MIL-A-9094 specifies the requiremerts for aircraft l~ghtning arresters for HIF antennas, and it probably will be extended to indlude all surge penetration into vehicles. MIL-E-6051 is the eiectroinagnetic compatibiiity specification, and refers to permissible EM pulse limits. MiL-B5087 is the standard military bonding specification and covers test current waveforms, bonding jumpcr sizes, protection of canopies, and lightming-induced surge penetration limits. There are other specifications with reference to lightning, but those listed herein are the principal ones with specific data on waveforms, test arrangements, and requirements. .

REFERENCE 1. Rotary Wilng Aircraft Susceptibility, DN 74A,

AFSC DH 1-4, 10 January 1972.

AMCP 706-202

CHAPTER 8 AVIONIC SUBSYSTEMS DESIGN &-I INTRODUCTION

I

8-1. 1 GENERAL clctrni")is efiW astheapAvioics(avatio Avonis (aiaton eectonicj i defnodas te aplication of elenronic techniques io accomplish such functions as communication, navigation, flight )ntrol. identificatio~n, sensivig, surveillance, and terget deaignation. The avionic subsystems will be defined by the detail specification. This chapter will discuss desin rquiemets o iterliic t'eses~asysems weith th uiemheniots toStractcs:sbytm withthe elioter.system. From an operational viewpoint, the helicopter avionic cowplrement can be subdivided into (1) the basic helicopter configuration, and (2) the specialmiwoneqwpent.8. The batsic helicopter configuration as discussed in thic handbook is limited to the space, weight, ad power requiremnents of the minimurn electronics necessary in order to provide the basic mission capability for a specific ciazz of heiicopter. The helicopter classes include light observatio'n, utility, tactical and heavy transport, and external heavy lift transpor-.. Special-misuion equipment is defincd as the additional e,'ectronics - beyond the basic communcation, navigation, and identification functions - requnired to accomplish specific missions such as IFFK flight, night operation under reduced visibility conditions, target detection and recognition, target dcsignation, and integrated fire control, such as is lound in gunships and tactical aircraft weapon systems.

magnetic compatibility/intcrfercnce (EMC/EM 1) must be considered. in general, the ioflowing dcsilbn sequencing must occur. 1. Determine the avionic requirements. 2.Dtrieheaonchrctiscs t 3. CDnteruineth ablonck diagramot:erinterfces heitefcet agam. the eonctrictal syste n h ar system.h 4.eeletricalbsclyoto ntear ai aototesser eeo 4 craft for mock-up purposes. . Develop a schematic wiring diagram for the 6. Develop an interconnect diagram. 7 eeo at it 7 eco at it Develop a wire list. .Dvlpaneetia la nlss 10. Complete a preliminery EMC/EMI analysis pa o h ytm A-i.!

Jmedium

Avionic procurement, installation, and quaiii-

)

cation, Along with bench, preflight and flight test requirements, are defined by Military Specifications such as MIL-STD-454, MIL-STD-461, MIL-STD462, MIL-STD-704, MIL-B-5087, MIL-W-5088, MIL-E.540, MIL-E-6051. and MIL-1-8700. Tbc first step in avionic systcm design is to determine the proper location for each individual system. Because avionic systems are made up of several subsystems and coimponents, it is mandatory that the total helicopter systcrr. and its environmental capability be known. Every avionic system component has temperature and vibration limE*tu!:ons. Before any placement or location is determined, the inter/intra-system compatibility of the location must bi, determined to insure that heat and vibration will not have a detrimental effect upon the performance of the equipment. in addition, electro-

ELECTROMAGNETIC CNPTBLF RGA RGA CMAIIIT

Interference generated by items of electrical/electronic equipment iirutalled in close proximity, as in a typical helicopter systcem, easily can result in an intolerable interference loe, that could reduce seriously the usefulness of airborne equipment, or might even render it ineperative. As defined in par. 9-11.2, AMCP 706-203, the prime contractor shall establish an overall integratcd EMI compatibility program for the helicopter. EMC is achieved by application of an optimum -

mhin,,,i.nn of miannapriai

-*-*----

*-*--

-

anti iephnicM -

-

r~norr-c-

from the earfiist design stage through the: final product or operational feasibility demonstration stage. Accordingly, an EMC program shall be es;ablished that will; i. Insure the efficiprn integration of engineering. management, and q, l1ity assurance tasks as thty relate to EMC. 2. 1isure the efficient integration of EMC withk all other systems an~d subsystems. The first requirement for acnieving EMC in an avionic system is that all major components and subsystems be designed, constructed, -,ndtested in comnpliance with MIL-STD-461l The second requirement is compliance with MIILE-6051 as an operating helicopter system, with all avionics and other equipme~nt installed and per. for-ning their normal functions. B-1

AMCP "&6202j 9-1.3 DEA2GN CONSIDERATIONS The design considerat'ons that follbw are applicable to EMI and should be ased to assist in keeping EMI to a minimum. The first design consideration iinvolves the creation of a good, basic grou-id plane. This is normally the avioic ompnentchasisor te arfrme srucure for the avionic system installation. An ideal ground zro-ipednce plan awileropotetia, povid reference b~se for all circuits, and a sink or trap for all sigalstha ndeire canoecrneintefernce sources.unifornn, desgnurcecsl.l Asecond deinconridecration, patclryat th lower communication frequencies, is the requirement for single-point 1grounding so as to avoid ground loops. The h-ige. circul~ting utirrents in ground loops are potential causes of interference. A third design consideradion concerns shielding practices for major components ard for the tntal aircraft installation.thdeinpaetocne A fourth design considcration calls for isolating, as fras possible, the power-carrying -vires and %cables from the high-impedance, low-lcvel signal wiring. The baý,ic p-inciple is to categorize conductors on the -l their Primary leaakange f basicc _r -h..eth......poinents are mnagnectic or clctrusiatic. Ail condu~ctors carlyin,3 power or signal eniergy have associated withý ar~ external or- leakage field that ran hiduce ut, wanted signals or noise in nearby conductors by 'i ductive or capacitive coupling. To minimize these undesirat~le field components, various techniques are usedI - such as electrostatic and magnetic shielding, spac searaiontwitin of irepais, cossver spaetistng spartio, f wre air, cossver wiring methods, use of field-absofbing materials, and netralzaton ethds.much sophstiate '~

Vthem

A fifth deinconsideration is to provide adequate troni-. eq.uipment, and for parts of tile vehicle structure that can contribute to the generation of clectri cal ncisi. All electrical and avionic equipment, subsystems, and systems that produce electromagnetic energy shall be installed to provide a continuous lowimpclanv'ý path from the equipment enclosure to the aircraft structure. The. designer mi.st demonstrate that thac proposed bon,' 4 methods result in a D-C resistance as specified Jh. ~.i various cl asses of bonding in MIL-B-5087. The design shall minimize the long-term effects of ope:ra' oal vibrotion, the effects of cerjrosion bz~ween adj..- nt surface and of' galvanic aztion, the diC!eCtric breakdown of insulating finishes, and the undzsii'nblc cifect of intermittent electrical contact. Bonding stiali be accomplished by direct mn..tal-to-metall contact whercvcw practicable. A bonding .imper shall be used where direct meta!-to-

~,8-2

mettal contact is impracticable. Such jumpers shall be pcfe nMLB58,o o h appropriate tnadt~Sa types, and shall be kept as short other

r addrctspoibeWhrpatcbltejm hjme addrc spsil.Weepatcbe shlr o xcd3i.i egh ufc rprto for bondý and grounds shall be accomplished by removing all anodic film, grease, paint and lacquer, or othecr high-resistance: materials from the imnmediate area of contact. Direct-to-basicý structure bonding "hlused wherever possible. For vehicles with metallic skin, the skin &Wal be designot. so that a low-impedance skir. is produced through inherent RF bonding during construction. RF bondint must oc accomplished bctwoer all structural comnportents. Hatches, access doors, and similar comnwirientshnot einhe roxuiy tointrfpermanenl soucn-o wiigsalbethrbnetorprantyisulated from the vehicle skin except for the protciesticdanbd.tishgldsrbedin thedstaign phasi, toanfer reishghlaly weithleauirngm eirlwthifam designers so as to resolve compatibility problems. For i;,uidelines to analysis and design, the design er.H14 gne hudcnutMLB58,AS Re.I. and-4cN6S1.douetsrfrecdi I..SD41 A sixth design consideration for rninimniziiie EN.1 is to separate :alid isolate pulse device and equipment from other devices that are high)) isusceptible to EMI. This is accomplished by attemptiall to separate use aas itroaortas sc tesa aas nergtrtas sc tm spit ponders, ani I H F transmitters from com~puters, data processors, and sliscrptible receivers. This is not always possible, -nasmuch as the physical locations of somie devicus are dictated by m.ission requirements. However, the designer should strive to achieve as physical and electrical isolation as is practicable. A fl-!inaldedrg conadialrabls tien cablvteuse o ie f doquilerinhieldescohaxil cables Ote cablesm or 9ire alhveamimmo90 reuin shld providg.Cnncosued withblc shilsfifselin cables shilllde. prvddwtblcshlsfiatengalehcd. -. ENIOM TA AS CS SET 8-. NIOMNA Ervironmental considerations are pertinent to the design of the b,ýsic avionic system, and to the airfrume-systcm interface. Susceptibility to rotor mnodulation must be consdte6e. The very high frequcncy omni-dircctional range (VOR), instrument landing system (ILS) localizer and glidescope, VHFFM %omer, and other cquipn~ent have been affected adversely by near-frequency rotor modulation. As rotor blades pass over the air-.raft, a modulation of the incoming wavefrorit is set up, with pronounced

K

Engineering Design Handbook Helicopter Engineering Part Two Detail Design AMCP 706-202 (Helicopter Engineering, Detail Design)

Engineering Design Handbook Helicopter Engineering Part One Preliminary Design (Helicopter Engineering AMCP 706-201, Preliminary Design)

Engineering Design Reliability Handbook

Engineering Design Reliability Handbook