Encyclopedia of Space Science and Technology

ENCYCLOPEDIA OF

SPACE SCIENCE AND TECHNOLOGY VOLUME 1

ENCYCLOPEDIA OF SPACE SCIENCE AND TECHNOLOGY

Editor Hans Mark The University of Texas at Austin

Associate Editors Milton A. Silveira Principal Engineer, Aerospace Corp. University of Vermont Michael Yarymovych President International Academy of Astronautics

Editorial Board Vyacheslav M. Balebanov Russian Academy of Sciences William F. Ballhaus, Jr. The Aerospace Corporation Robert H. Bishop University of Texas at Austin Aaron Cohen Texas A & M University Wallace T. Fowler University of Texas at Austin F. Andrew Gaffney Vanderbilt University Medical Center Owen K. Garriott University of Alabama Tom Gehrels University of Arizona at Tucson Gerry Griffin GDG Consulting Milton Halem NASA-Goddard Space Flight Center

John S. Lewis University of Arizona at Tucson Thomas S. Moorman Booz Allen & Hamilton Norman F. Ness University of Delaware Robert E. Smylie National Aeronautics Space Administration Richard H. Truly National Renewable Energy Laboratory Albert D. Wheelon Hughes Aircraft Co. Peter G. Wilhelm U.S. Naval Research Laboratory Laurence R. Young Massachusetts Institute of Technology Alexander Zakharov Russian Academy of Sciences Managing Editor Maureen Salkin Editorial Staff Vice President, STM Books: Janet Bailey Executive Editor: Jacqueline I. Kroschwitz Director, Book Production and Manufacturing: Camille P. Carter Managing Editor: Shirley Thomas Illustrations Manager: Dean Gonzalez Assistant Managing Editor: Kristen Parrish Editorial Assistant: Surlan Murrell

ENCYCLOPEDIA OF

SPACE SCIENCE AND

TECHNOLOGY VOLUME 1 Hans Mark Editor

Milton Silveira Associate Editor

Michael I. Yarymovych Associate Editor

Maureen Salkin Managing Editor

The Encyclopedia of Space Science and Technology is available Online in full color at www.interscience.wiley.com/esst

A John Wiley & Sons, Inc., Publication

Copyright r 2003 by John Wiley & Sons, Inc. All rights reserved. Published by John Wiley & Sons, Inc., Hoboken, New Jersey. Published simultaneously in Canada. No part of this publication may be reproduced, stored in a retrieval system, or transmitted in any form or by any means, electronic, mechanical, photocopying, recording, scanning, or otherwise, except as permitted under Section 107 or 108 of the 1976 United States Copyright Act, without either the prior written permission of the Publisher, or authorization through payment of the appropriate per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923, 978-750-8400, fax 978-750-4470, or on the web at www.copyright.com. Requests to the Publisher for permission should be addressed to the Permissions Department, John Wiley & Sons, Inc., 111 River Street, Hoboken, NJ 07030, (201) 748-6011, fax (201) 748-6008, e-mail: [email protected]. Limit of Liability/Disclaimer of Warranty: While the publisher and author have used their best efforts in preparing this book, they make no representations or warranties with respect to the accuracy or completeness of the contents of this book and specifically disclaim any implied warranties of merchantability or fitness for a particular purpose. No warranty may be created or extended by sales representatives or written sales materials. This advice and strategies contained herein may not be suitable for your situation. You should consult with a professional where appropriate. Neither the publisher nor author shall be liable for any loss of profit or any other commercial damages, including but not limited to special, incidental, consequential, or other damages.

For general information on our other products and services please contact our Customer Care Department within the U.S. at 877-762-2974, outside the U.S. at 317-572-3993 or fax 317-572-4002. Wiley also publishes its books in a variety of electronic formats. Some content that appears in print, however, may not be available in electronic format. Library of Congress Cataloging-in Publication Data: Encyclopedia of Space Science & Technology / Hans Mark [editor]. p. cm. Includes index. ISBN 0-471-32408-6 (set: acid-free paper) 1. Space Science–Encyclopedias. I. Title: Encyclopedia of Space Science and Technology. II. Mark, Hans, 1929QB497.E53 2003 2002028867 500.50 03—dc21 Printed in the United States of America. 10 9 8 7 6 5 4 3 2 1

High Flight By Pilot Officer John G. Magee, Jr., RCAF Oh, I have slipped the surly bonds of earth And danced the skies on laughter-silvered wings; Sunward I’ve climbed, and joined the tumbling mirth Of Sun-split clouds – and done a hundred things You have not dreamed of – wheeled and soared and swung High in the sunlit silence. Hov’ring there. I’ve chased the shouting wind along, and flung My eager craft through footless halls of air. Up, up the long, delirious, burning blue I’ve topped the windswept heights with easy grace Where never lark, or even eagle flew. And, while with silent, lifting mind I’ve trod The high untrespassed sanctity of space Put out my hand, and touched the face of God. Pilot Officer John Gillespie Magee, Jr., an American serving with the Royal Canadian Air Force, composed ‘‘High Flight.’’ He was born in Shanghai, China in 1922, the son of missionary parents, Reverend and Mrs. John Gillespie Magee; his father was an American and his mother was originally a British citizen. He came to the U.S. in 1939 and earned a scholarship to Yale, but in September 1940 he enlisted in the RCAF and graduated as a pilot. He was sent to England for combat duty in July 1941. In August or September 1941, Pilot Officer Magee composed ‘‘High Flight’’ and sent a copy to his parents. Several months later, on December 11, 1941 his ‘‘Spitfire’’ airplane collided with another plane over England and Magee, only 19 years of age, crashed to his death. His remains are buried in the churchyard cemetery at Scopwick, Lincolnshire. This can be found on the website: http://www.wpafb.af.mil/museum/history/prewwii/jgm.htm

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PREFACE Nicolaus Copernicus and Galileo Galilei developed the scientific knowledge that became the underpinning of spaceflight. Edward Everett Hale in ‘‘The Brick Moon’’ and Jules Verne in ‘‘From the Earth to the Moon’’ dreamed and wrote about it. But finally in the last half of the twentieth century, it was the Americans and the Soviet Russians, locked in the throes of the Cold War, who accomplished it. A good case can be made that when historians look back at the twentieth century, the initial efforts of humankind to slip ‘‘the surly bonds of Earth’’ will play a dominant role. Today, we call the sixteenth century the ‘‘Age of Exploration’’ because by combining the fore-and-aft sail rig of Arab dhows with the study hull of the Baltic cog, the ‘‘caravel’’ was created that could safely sail all the oceans of the world. Thus, in the final years of the fifteenth century, Bartolomeo Diaz, Christopher Columbus, and Vasco da Gama opened astonishing new vistas using the caravels. In less than a century after their epochal voyages, the geography of the Earth was essentially understood and things were forever changed. Today, because of the advent of rocket technology, we stand at the threshold of sending humans to Mars as well as to other places in the Solar System. We are within a decade of sending people back to our own Moon to establish permanent stations to exploit lunar resources and to create staging bases for the large-scale exploration of the Solar System. As was the case half a millennium ago, things will change forever when this is done. We have both been involved in this initial exploratory effort in an intimate way. One of us (Richard H. Truly) has actually flown in space and both of us have participated in and led the organizations established in the United States to conduct space exploration. Both of us have also been touched by the brutal wars of the twentieth century, and we therefore know how these have influenced the lives of people all over the world as well. The idea of this Encyclopedia of Space Science and Technology was conceived late in 1997 when one of us (Hans Mark) had a conversation with Dr. Edmund H. Immergut, who has had a long and distinguished career in scientific publishing and in the production of encyclopedias. He believed that the enterprise of space exploration was far enough along – 40 years after the first orbital flight of Sputnik I – that a good technical encyclopedia on the subject would be timely and appropriate. In developing the ideas for the encyclopedia, the following principles were established. *

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The encyclopedia would be written at a high technical level, i.e., for an audience of technically literate people who were not experts in space science or technology. The encyclopedia would contain articles that would describe the technology of space exploration as well as the scientific results and their applications. The authors who would be selected to write articles would be people who are, or have been, active participants in enterprise of space exploration. vii

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The encyclopedia would be international and would attempt to capture the spirit that animated the enterprise for the past half century. The encyclopedia would have a broadly based editorial board whose members would help to select authors and assist in passing judgment on the quality of the work.

It is our hope that we have largely adhered to these principles. The Encyclopedia of Space Science and Technology consists of nearly 80 articles organized under eight separate categories. There is an appropriate index and a table of contents that should make it easy for readers to find the topic of interest for which they are searching. Throughout this work, both of us have enjoyed working with old and new colleagues. We would like to extend our appreciation to everyone who participated in this effort, first and foremost, our authors for their contributions, our Associate Editors, Drs. Milton A. Silveira and Michael I. Yarymovych, and all the members of our Editorial Board for their participation and advice. Finally, special thanks are due to our Managing Editor, Ms. Maureen A. Salkin, for her tireless and highly diplomatic efforts to keep things rolling so that we can now all see the final result. Richard H. Truly Golden, Colorado Hans Mark Austin, Texas

CONTRIBUTORS Brian Allen, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets R.C. Anderson, California Institute of Technology, Pasadena, California, Pathfinder Mission to Mars Kenneth M. Baldwin, University of California, Irvine, California, Muscle Loss in Space: Physiological Consequences Vyacheslav M. Balebanov, Russian Academy of Sciences, Institute of Space Research, Russia, Plasma Thrusters V.A. Bartenev, Scientific-Production Association of Applied Mechanics, Russia, Communication Satellite Development in Russia Alexander T. Basilevsky, Vernadsky Institute of Geochemistry and Analytical Chemistry, Russian Academy of Sciences, Moscow, Russia, Exploration of the Moon by Soviet Spacecraft; Venus Missions J.R. Beattie, Westlake Village, California, Rockets, Ion Propulsion Robert R. Bennet, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets George A. Berkley, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Jon H. Brown, Fort Worth, Texas, Spacecraft Guidance, Navigation and Control Systems Boris Chertok, ENERGIA Space Association, Russia, Sputnik 1: The First Artificial Earth Satellite Edward L. Chupp, University of New Hampshire, Durham, New Hampshire, Sun Anita L. Cochran, The University of Texas McDonald Observatory, Austin, Texas, Comets Aaron Cohen, NASA – Lyndon B. Johnson Space Center, Space Shuttle Orbiter Project Office, Houston, Texas, Space Shuttle Orbiter Richard J. Cohen, Harvard University—Massachusetts Institute of Technology, Cambridge, Massachusetts, Cardiovascular System in Space Glenn D. Considine, Westfield, Massachusetts, Mars Douglass B. Cook, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Robert L. Crippen, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets F.A. Cucinotta, NASA Johnson Space Center, Houston, Texas, Space Radiation Alexander F. Dedus, Russian Aviation and Space Agency, Russia, Russian Spaceports J.F. Dicello, Johns Hopkins University School of Medicine, Baltimore, Maryland, Space Radiation Steven D. Dorfman, Hughes Electronics Corporation, Los Angeles, California, Commercial Applications of Communications Satellite Technology; Communications Satellites, Technology of Timothy E. Dowling, University of Louisville, Louisville, Kentucky, Jupiter V. Reggie Edgerton, University of California, Los Angeles, California, Muscle Loss in Space: Physiological Consequences Alexander N. Egorov, Yu.A. Gagarin Cosmonaut Training Center, Russia, Cosmonauts Selection and Preparation Gabriel Elkaim, Stanford University, Stanford, California, Global Positioning System (GPS) Maxime A. Faget, NASA-Johnson Space Center, Houston, Texas, U.S. Manned Spaceflight: Mercury to the Shuttle Dale Fenn, Orbital Sciences Corporation, Dulles, Virginia, Air and Ship-Based Space Launch Vehicles Harold B. Finger, National Aeronautics and Space Administration and Atomic Energy Commission, Washington, D.C., Nuclear Rockets and Ramjets Uwe Fink, Lunar and Planetary Lab University of Arizona, Tucson, Arizona, Saturn System Charles T. Force, Tracy’s Landing, Maryland, Earth-Orbiting Satellites, Data Receiving and Handling Facilities Marvin Glickstein, Pratt & Whitney, Palm Beach, Florida, Liquid-Fueled Rockets Teresa Gomez, NASA Johnson Space Center, Houston, Texas, Astronauts and the People who Selected Them: A Compendium L. Gorshkov, ENERGIA RSC, Russia, Russian Space Stations

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Robert P. Graham, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Anatoly I. Grigoriev, Institute of Biomedical Problems, Russian Academy of Sciences, Moscow, Russia, Biomedical Support of Piloted Spaceflight; Space Life Sciences Herbert Gursky, Naval Research Laboratory, Washington, DC, Science from Sounding Rockets Martin Harwit, Cornell University, Ithaca, New York, Astronomy–Infrared W. Michael Hawes, NASA, Washington, District of Columbia, International Space Station Clark W. Hawk, Madison, Alabama, Rocket Propulsion Theory Steven A. Hawley, NASA Johnson Space Center, Houston, Texas, Human Operations in Space During the Space Shuttle Era Hans E.W. Hoffmann, ORBCOMM LLC, Dulles, Virginia, Spacelab Stephen Horan, New Mexico State University, Las Cruces, New Mexico, Earth-Orbiting Satellites, Data Receiving and Handling Facilities Ross M. Jones, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, Planetary Exploration Spacecraft Design Russell Joyner, Pratt & Whitney, Palm Beach, Florida, Liquid-Fueled Rockets Joseph Kerwin, Houston, Texas, Skylab Joseph J. Kliger, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Petr I. Klimuk, Yu.A. Gagarin Cosmonaut Training Center, Russia, First Flight of Man in Space Stanislav Nikolaevich Konyukhov, M.K. Yangel’ Yuzhnoye State Design Office, Dniepropetrovsk, Ukraine, Conversion of Missiles into Space Launch Vehicles Jean Kovalevsky, Cerga-Observatoire de la Coˆte d’Azur, Grasse, France, Optical Astrometry from Space A.G. Kozlov, Scientific-Production Association of Applied Mechanics, Russia, Communication Satellite Development in Russia Alexander N. Kuznetsov, Russian Aviation and Space Agency, Russia, Russia’s Launch Vehicles; Russian Spaceports James W. Layland, California Institute of Technology, Pasadena, California, Deep Space Network, Evolution of Technology David S. Leckrone, NASA, Goddard Space Flight Center, Greenbelt, Maryland, Hubble Space Telescope James R. Lesh, California Institute of Technology, Pasadena, California, Deep Space Network, Evolution of Technology John S. Lewis, University of Arizona, Tucson, Arizona, Space Resources, Occurrence and Uses Wah L. Lim, Hughes Electronics Corporation, Los Angeles, California, Commercial Applications of Communications Satellite Technology Glynn S. Lunney, Houston, Texas, NASA Mission Operation Control Center at Johnson Space Center Ronald W. Lyman, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Dmitry K. Malashenkov, Institute of Biomedical Problems, Russian Academy of Sciences, Moscow, Russia, Biomedical Support of Piloted Spaceflight; Space Life Sciences Jerry W. Manweiler, Fundamental Technologies LLC, Lawrence, Kansas, Interplanetary Medium Hans Mark, Austin, Texas, Evolution of U.S. Expendable Launch Vehicles Ian R. McNab, The University of Texas at Austin, The Institute for Advanced Technology, Austin, Texas, Electromagnetic Propulsion Valeriy A. Menshikov, Khrunichev Space Center, Moscow, Russia, Global Navigation Satellite System; Military Use of Space Jerome H. Molitor, Westlake Village, California, Rockets, Ion Propulsion Vasily I. Moroz, Space Research Institute, Russian Academy of Sciences, Moscow, Russia, Exploration of Mars by the USSR; Venus Missions Alexey I. Morozov, Russian Science Center, Kurchatov Institute, Russia, Plasma Thrusters David Morrison, NASA Ames Research Center, Moffett Field, California, Asteroids; Astrobiology

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Adam L. Mortensen, USAF, USSPACE/SIOE-r, Colorado Springs, Colorado, Military Ground Control Centers, United States Douglas J. Mudgway, California Institute of Technology, Pasadena, California, Deep Space Network, Evolution of Technology F. Robert Naka, CERA, Incorporated, Concord, Massachusetts, Space Programs Related to National Security John J. Neilon, Cocoa Beach, Florida, Eastern Launch Facilities, Kennedy Space Center Robert M. Nelson, Jet Propulsion Laboratory, Pasadena, California, Mercury R. Steven Nerem, University of Colorado, Colorado Center for Astrodynamics Research, Boulder, Colorado, Earth Orbiting Satellite Theory Arleigh P. Neunzert, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Arnauld Nicogossian, NASA Headquarters, Washington, D.C., Biological Responses and Adaptation to Spaceflight: Living in Space—an International Enterprise Bradford Parkinson, Stanford University, Stanford, California, Global Positioning System (GPS) Billy H. Prescott, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets C. Paul Pulver, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Craig D. Ramsdell, Beaumont Hospital, Royal Oak, Michigan, Cardiovascular System in Space P. Krishna Rao, National Oceanic and Atmospheric Administration, Silver Spring, Maryland, Weather Satellites Lawrence L. Rauch, California Institute of Technology, Pasadena, California, Deep Space Network, Evolution of Technology John C. Ries, The University of Texas at Austin, Center for Space Research, Austin, Texas, Precision Orbit Determination for Earth Observation Systems Robert Rosen, NASA Ames Research Center, Moffett Field, California, Liquid-Fueled Rockets Duane L. Ross, NASA Johnson Space Center, Houston, Texas, Astronauts and the People who Selected Them: A Compendium Roland R. Roy, Brain Research Institute, University of California, Los Angeles, California, Muscle Loss in Space: Physiological Consequences Roald Sagdeev, University of Maryland, College Park, Maryland, Vega Project Donald R. Sauvageau, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets H.H. Schmitt, University of Wisconsin—Madison, Wisconsin, Apollo 17 and the Moon B.E. Schutz, University of Texas at Austin, Center for Space Research, Austin, Texas, Size and Shape of Earth from Satellites Yuri P. Semyonov, ENERGIA RSC, Russia, Russian Space Stations William T. Shearer, Texas Children’s Hospital, Houston, Texas, Immunology and Infection in Space Milton A. Silveira, NASA Johnson Space Center, Houston, Texas, Space Shuttle Orbiter; U.S. Manned Space Flight: Mercury to the Shuttle S. Fred Singer, The Science & Environmental Policy Project (SEPP), Arlington, Virginia, Weather Satellites G.M. Solovyev, Khrunichev Space Center, Russia, Global Navigation Satellite System Gerald Sonnenfeld, Morehouse School of Medicine, Atlanta, Georgia, Immunology and Infection in Space Yu.B. Sosyurka, Yu.A. Gagarin Cosmonaut Training Center, Russia, Cosmonauts Selection and Preparation James Spilker, Stanford University, Stanford, California, Global Positioning System (GPS) Paul D. Spudis, Lunar and Planetary Institute, Houston, Texas, Moon Lawrence A. Sromovsky, University of Wisconsin, Madison, Wisconsin, Uranus and Neptune William Stoney, Mitretek Corporation, Reston, Virginia, Civil Land Observation Satellites Byron D. Tapley, The University of Texas at Austin, Center for Space Research, Austin, Texas, Precision Orbit Determination for Earth Observation Systems Jill Tarter, SETI Institute, Mountain View, California, Extraterrestrial Life, Searching for

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Roger Vignelles, Corbeil-Essonnes, France, Ariane Rocket Program G.I. Vorobyov, Yu.A. Gagarin Cosmonaut Training Center, Russia, First Flight of Man in Space Steven R. Wassom, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Martin C. Weisskopf, NASA – Marshall Spaceflight Center, Huntville, Alabama, Chandra X-ray Observatory Michael Werner, Jet Propulsion Laboratory, Pasadena, California, Astronomy-Infrared Nicholas J. Whitehead, Thiokol Propulsion, Inc., Brigham City, Utah, Solid Fuel Rockets Simon P. Worden, USAF, USSPACE/SIOE-r, Colorado Springs, Colorado, Military Ground Control Centers, United States V.I. Yaropolov, Yu.A. Gagarin Cosmonaut Training Center, Russia, Cosmonauts Selection and Preparation Michael I. Yarymovych, Boeing Space and Communications (Retired), Seal Beach, California, Evolution of U.S. Expendable Launch Vehicles Laurence R. Young, Massachusetts Institute of Technology, Cambridge, Massachusetts, Artificial Gravity Eliot Young, Southwest Research Institute, Boulder, Colorado, Pluto and Charon Leslie Young, Southwest Research Institute, Boulder, Colorado, Pluto and Charon Alexander V. Zakharov, Space Research Institute, Russian Academy of Sciences, Moscow, Russia, Exploration of Mars by the USSR

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A AIR AND SHIP-BASED SPACE LAUNCH VEHICLES Introduction In 1957, the Soviet Union placed the first man-made object in orbit around the earth. Since then, numerous launch vehicles have been developed to improve the performance, reliability, and cost of placing objects in orbit. By one estimate, roughly 75 active space launch vehicles either have established flight records or are planning an inaugural launch within the year. This does not include the numerous launch vehicles from around the world that are no longer operational such as the Jupiter, Redstone, Juno, Saturn, Scout, Thor, Vanguard, and Conestoga family of rockets from the United States or the N-1 from the former Soviet Union, to name just a few. Despite the many differences among all of these launch vehicles from both past and present, one common element can be found in all but four of them: they are ground-launched. Of the four exceptions, two are air-launched (NOTSNIK and Pegasus), one is ship-launched (Sea Launch), and one is submarine-launched (Shtil). It is important to keep in mind that numerous air-launched and ship-launched suborbital launch systems are in use by militaries, commercial entities, and educational institutions. However, the four mentioned are the only mobile launch systems that can place objects into a sustainable Earth orbit.

Mobile Space-Launched Vehicles Project Pilot (NOTSNIK). NOTSNIK is the oldest and, until recently, the least well known of the four mobile space-launched systems. Following the launch of Sputnik by the Soviet Union, President Eisenhower’s administration elicited proposals to launch a satellite into orbit. The Naval Ordinance Test 1

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Station (NOTS) located at China Lake in California proposed launching a rocket from a jet fighter (1). The idea is the same as that of the current Pegasus vehicle: reduce the amount of energy needed to place a payload into orbit by launching it above the denser portion of the atmosphere. In this fashion, the engineers at NOTS designed a vehicle from existing rocket motors that could place a 2-pound satellite in a 1500-mile-high orbit. The engineers recognized the energy savings from such a launch concept and also the utility of such a flexible platform. Launching from a jet fighter could, theoretically, place a satellite into any orbit from anywhere in the world at any time. The U.S. Navy accepted the proposal from NOTS in 1958, by some accounts as a safety net in the event that the ongoing Vanguard project was unsuccessful. The program was officially called Project Pilot, but the engineers at NOTS preferred the name NOTSNIK in direct reference to the Soviet satellite that was currently orbiting above them and the rest of the world. A Douglas Aircraft F4D1 Skyray was the carrier aircraft for the rocket and consequently was considered the first stage. The second and third stages were modified antisubmarine missiles. The final stage was taken from a Vanguard rocket. The entire launch vehicle measured a mere 14 feet in length and had four fins at the aft end that provided a span of 5 feet. The NOTSNIK was launched six times from an altitude of about 41,000 ft. Four of those launches ended in known failures. However, the results of two have never been verified. Some in the program insist that they achieved their goal of placing the small payload of diagnostic instruments in orbit. At least one ground station in New Zealand picked up a signal in the right place at the right time. However, confirmation that the signal was from the NOTSNIK payload was never established. Even the possibility of a success was veiled in secrecy for more than 40 years for, by all accounts, two critical reasons. The first was that in the days following the early embarrassments of Vanguard, the Eisenhower administration did not want to claim success unless it was absolutely certain. The second reason was that a mobile air-launched system that could reach orbit had extremely appealing military applications. However, the tactical advantages of such a system were far outweighed by the strategic consequences, as stated in the Antiballistic Missile (ABM) Treaty between the United States and the former Soviet Union that was concluded in 1972 (2): Further, to decrease the pressures of technological change and its unsettling impact on the strategic balance, both sides agree to prohibit development, testing, or deployment of sea-based, air-based, or space-based ABM systems and their components, along with mobile land-based ABM systems. Should future technology bring forth new ABM systems ‘based on other physical principles’ than those employed in current systems, it was agreed that limiting such systems would be discussed, in accordance with the Treaty’s provisions for consultation and amendment.

Pegasus. Roughly 30 years later, while NOTSNIK remained an official government secret, the idea of launching payloads into space from an airborne platform was revisited in the form of the Pegasus launch vehicle. The driving forces behind NOTSNIK and Pegasus were essentially the same. An air-launched space vehicle provides several advantages compared with ground-based counterparts.

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As an example, Pegasus is launched at an altitude of 39,000 ft, which is above a significant portion of the atmosphere. As mentioned, with NOTSNIK, this eliminates the need for extra performance that would otherwise be needed to overcome atmospheric forces. This also implies that the structural components of the vehicle can be lighter, which improves the efficiency of the rocket as a whole. The energy required from the launch vehicle is also reduced by the speed already achieved by the carrier aircraft. An air-launched system also allows applying more of the impulse of the first stage along the velocity vector. This is a more efficient use of the vehicle’s energy than that of ground-launched vehicles that must first apply the thrust almost perpendicular to the velocity vector already imparted by Earth’s rotation. These factors combine to produce a requirement for a velocity increment that is on the order of 10% less than a comparable groundlaunched rocket. The Pegasus vehicle is a winged, three-stage, solid rocket booster (Fig. 1). It is the first space-launched vehicle developed solely with commercial funding. Three versions have been developed and flown over the years: Standard, Hybrid, and XL. The XL is the only vehicle within the Pegasus family currently in production. The XL is roughly 10,000 lbm heavier than the Standard or Hybrid models and is roughly 6 ft longer. Because the XL extends farther aft beneath the L-1011 carrier aircraft, the port and starboard fins become an obstacle to the landing gear doors. To correct this problem, the port and starboard fins were Payload separation system

Avionics structure Stage 2 motor Wing

Payload fairing Stage 3 motor Fin Interstage

Aft skirt assembly Stage 1 motor

Figure 1. Disassembled version of standard Pegasus launch vehicle.

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modified to include an anhedral of 231. To maintain commonality between the various members of the Pegasus family of vehicles, the same anhedral was introduced into the Standard vehicle, which was then given the designation Pegasus Hybrid. Other than the anhedral of the fins, the Standard and Hybrid vehicles are exactly the same. The Standard, the first Pegasus vehicle built, was flown on six missions. The Hybrid vehicle has flown four times. The XL vehicle has flown 21 times. Of 31 Pegasus launches, only three missions failed to reach orbit. The Pegasus XL was designed and developed to provide increased performance above and beyond that provided by the Standard and Hybrid vehicles. A typical Pegasus XL vehicle weighs roughly 51,000 lbm at launch, is 55.4 ft long and 50 inches in diameter, and the wingspan is 22 ft (3). At launch, the Pegasus XL is carried aloft by the company’s carrier aircraft, a modified L-1011, which originally saw commercial service with Air Canada. The vehicle is dropped from an altitude of 39,000 ft at Mach 0.8. Five seconds after release from the L-1011, the first stage ignites and the vehicle’s on-board flight computer continues the sequence of events that eventually lead to orbital insertion. The brief coast period between drop and stage one ignition is designed to provide a safe distance between the L-1011 and the launch vehicle. The Pegasus Standard vehicle was originally dropped from a NASA-owned and operated B-52. The Pegasus vehicle was attached to one of the pylons underneath the starboard wing much in the same manner as the early supersonic and hypersonic test vehicles such as the X-15. For a variety of reasons, Orbital purchased and modified the L-1011 to facilitate all future launches. Unlike the B-52 that supported initial Pegasus launches, the L-1011 carries the Pegasus vehicle underneath the fuselage rather than underneath the wing. Once Pegasus is ready to be mated to the carrier aircraft, it is towed from Orbital’s integration facility at VAFB to the plane on the Assembly and Integration Trailer (AIT). Regardless of where the launch is to take place, the Pegasus is always integrated and mated to the L-1011 at VAFB. From there, the launch system can travel to any location in the world for launch. There is enough ground clearance for the L-1011 to take off and land with Pegasus attached underneath. However, the added height of the AIT underneath Pegasus requires raising the L-1011 off the ground slightly by hydraulic jacks to mate Pegasus to the carrier aircraft (Fig. 2). While mated to the L-1011, the vertical rudder actually protrudes into the plane’s fuselage in a compartment specifically designed for this purpose. When mating the Pegasus to the L-1011, the rudder is usually detached from the Pegasus vehicle and placed inside the housing first. Then the Pegasus is rolled underneath the L-1011 and attached to the rudder and then to the plane. Removing the rudder first minimizes the height to which the L-1011 needs to be raised for the mating process. The entire mating process from rollout to mating takes about 6 hours. Pegasus is attached to the L-1011 using four hooks on the center box of the wing and a fifth hook on the forward portion of the vehicle. The inside of the airplane has been stripped of all unnecessary equipment and hardware. Up front in what would normally be the first class cabin are eight seats for personnel during ferry flights from VAFB to the launch site of interest and two computer stations from which personnel can monitor the health of the vehicle and the payload. The rest of the interior of the cabin has been completely gutted. Access to the rear portion of the aircraft cabin is obtained through a galley door.

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Figure 2. Fully assembled Pegasus launch vehicle being mated to the L-1011 aircraft. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

Unlike most other launch vehicles in the U.S. fleet, the Pegasus launch vehicle is integrated horizontally on the AIT (Fig. 3). Horizontal integration facilitates easy access to the vehicle and eliminates the need for high bays and large cranes. Components are received as needed either from groups within Orbital Sciences or from outside vendors. To ensure that all of the major flight hardware and software is thoroughly tested before flight, Pegasus, like many other vehicles, is subjected to a series of ‘‘fly to orbit’’ simulations at various stages of the integration process. Four flight tests are normally performed. The first tests the three stages individually. The second test is conducted after the three stages are electrically mated together. The third test is performed after the three stages are electrically and mechanically mated and the stack is electrically mated to the payload. The fourth and final flight test is performed once the payload has been mechanically mated to the rest of the vehicle and the half of the fairing that includes the pyro devices necessary for jettisoning the shroud is

Figure 3. Horizontal integration of Pegasus launch vehicle. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

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electrically mated. These tests are intended to verify that various systems function and also respond as expected to known disturbances. If the inertial measurement unit (IMU) onboard receives data to indicate that an unexpected attitude change has occurred, will the fins or thrust vector control systems respond accordingly? Are all the commands to the various subsystems appropriate, and do those subsystems respond appropriately? Once the Pegasus vehicle has been mated to the L-1011 carrier aircraft, one last test is performed, called the Combined Systems Test (CST). This test verifies that the launch vehicle and the carrier aircraft are communicating as expected. This is particularly important since the vehicle’s health can be monitored both from telemetry that is broadcast from the vehicle to the ground via antennas on Pegasus and also by the computer stations inside the L-1011 via hardwired electrical connections. More importantly, some data and commands are sent to the Pegasus vehicle before launch. The only method currently available for accomplishing this transfer of data is through the electrical connections between the Pegasus vehicle and the carrier aircraft. To be fully mobile, the Pegasus launch system must also be fully self-contained. Except for those services provided by the range (such as radar coverage), the L-1011 can transport all of the equipment required to support a launch of Pegasus, including, of course, Pegasus itself (Fig. 4). Some launches take place off the coast of California where the Western Range (based at VAFB) is the lead range. In these instances, no ferry flight is required. The L-1011 simply takes off from VAFB and flies to the designated drop point roughly 100 nmi out to sea. The checklist that is processed in the control room on the day of launch requires about 4 to 5 hours to complete. The L-1011 usually takes off an hour before the scheduled launch time. If all systems are ‘‘go,’’ as determined by the mission team members in the control room, the launch conductor on the ground commands the pilot of the L-1011 to drop the Pegasus from the carrier aircraft. Shtil. In a classic example of turning swords into plowshares, the Russian Navy developed a satellite delivery system for nonmilitary applications that uses a submarine-launched. The SS-N-23 (NATO’s designation) is a three-stage

Figure 4. L-1011 aircraft taking off with Pegasus. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

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liquid-fueled vehicle that can deliver small satellites to low Earth orbit. Very little is known about this launch vehicle service including performance to various altitudes and inclinations. What is known is that two satellites belonging to the Technical University of Berlin were successfully launched in 1998 from a Russian submarine for the stunningly low price of $150,000 (4). Some sources indicate that the typical commercial price for a Shtil launch is actually in the neighborhood of $500,000. There are two possible reasons for the low cost of a Shtil launch. The first is that more than 200 missiles have already been produced by the Russian military. There is also speculation that offering commercial launch services provides a way to maintain proficiency in launching missiles without using precious military funding. One disadvantage of this system is that the Shitl vehicle likely does not have enough performance to achieve circular orbits in the medium to high Low Earth Orbit (LEO) altitudes (4). This is a direct result of the Shtil’s heritage as a ballistic missile first and foremost. Sea Launch. The most recent mobile launch system is the Sea Launch vehicle which is launched from a converted oil-drilling platform along the equator (Fig. 5). Sea Launch is both the name of the launch vehicle and the name of the international joint venture that provides the launch services. The partnership is comprised of Boeing, KB Yuzhnoye of Ukraine, which provides the two Zenit stages, and RSC Energia of Russia, which provides the Block DM-SL upper stage. The launch vehicle and payload integration takes place at the vehicle’s home port of Long Beach, California. Once integration is complete, the launch vehicle is loaded onto the converted oil-drilling platform and towed to a predetermined

Figure 5. Computer simulation of Sea Launch. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

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launch location at the equator, specifically 1541 West. Once on site, the Zenit 3SL is raised into its launch attitude (vertical) and launched. A second ship that houses mission personnel and the control room monitors the launch from nearby. The vehicle itself is a little less than 200 ft long and roughly 13 ft in diameter. The performance to Geosynchronous Transfer Orbit (GTO) is approximately 5250 kg (4). ‘‘In terms of spacecraft mass in final orbit, this would be equivalent to approximately 6000 kg of payload capability if launched from Cape Canaveral, because the spacecraft does not need to perform a plane change maneuver during the Geosynchronous Earth Orbit (GEO) circularization burn’’ (5). There are three key phases in the integration of a Sea Launch vehicle (5). Phase I takes place in the Payload Processing Facility (PPF). This phase includes receipt of the spacecraft, processing of the spacecraft, testing, and enclosure within the payload fairing. Phase II takes place on the Assembly and Command Ship (ACS). This entails mating the encapsulated spacecraft to the launch vehicle and testing the integrated stack. Phase III takes place on the Launch Platform (LP) once the vehicle has been transferred from the ACS. While still in port, the integrated launch vehicle is raised to its vertical launch attitude so that a series of tests can be conducted. The launch vehicle is then lowered back into a horizontal position, stored in an environmentally controlled room, and transported to the equator while on board the launch platform. At the launch site, the launch vehicle is rolled out to the launch pad, raised to a vertical attitude again, and fueled. The launch is performed by an automated system and monitored by the Assembly and Command Ship which is moved for launch to a distance 6.5 km away (Fig. 6). The Assembly and Command Ship for Sea Launch serves as the launch vehicle integration and testing facility. In addition to acting as the temporary home for launch crews, the ship also houses the Launch Control Center (LCC) and the equipment necessary to track the initial ascent of the rocket. Unlike the

Figure 6. Sea Launch successfully lifts DIRECTV 1-R satellite into orbit. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

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Pegasus carrier aircraft that was modified after serving in a different capacity, the ACS was designed and constructed specifically to suit the unique requirements of Sea Launch. The ship is roughly 660 ft long and 110 ft in beam and has an overall displacement of approximately 30,830 tonnes. The rocket assembly facility is on the main deck of the ACS where the launch vehicle integration takes place. This activity is conducted before setting sail for the equator and simultaneously with spacecraft processing. After the spacecraft has been satisfactorily processed, it is encapsulated and transferred to the rocket assembly compartment, where it is mated to the launch vehicle. Following integration and preliminary testing, the integrated launch vehicle is transferred to the launch platform. Then both ships begin the journey to the equator, which takes roughly 12 days. The launch platform has all of the necessary systems for positioning and fueling the launch vehicle, as well as conducting the launch operations. Once the launch vehicle has been erected and all tests are complete, personnel are evacuated from the launch platform to the ACS using a link bridge between the vessels or a helicopter. Redundant radio-frequency links between the vessels permit personnel on the ACS to control all aspects of the launch, even when the command ship has retreated to a safe distance before launch. The launch platform, which was converted from an oil drilling platform, is very stable. It is supported by a pair of large pontoons and is propelled by a four-screw propulsion system (two in each aft lower hull). Once at the launch location, the pontoons are submerged to a depth of 70.5 ft to achieve a more stable attitude for launch, level to within approximately 11.

Advantages of Mobile Space-Launched Systems NOTSNIK, Pegasus, Sea Launch, and Shtil were never intended to replace the existing fleet of ground-launched rockets. Rather, they effectively supplement the existing worldwide capability by providing additional services to a targeted market of payloads that benefit greatly from the mobility and flexibility of these unique space-launch systems. These vehicles can provide services similar to ground-launched vehicles for payloads within their weightclass. In fact, all four vehicles have fixed launch locations for standard services. For example, Pegasus uses the launch location of 361 N, 2371 E for all high-inclination missions that originate from VAFB. In this regard, the mobile launch systems are no different from ground-launched vehicles in that they repeatedly launch from a fixed location, albeit a location that is not on land. However, they can also offer services and performance that avoid many of the restrictions inherent in being constrained to a particular launch site. Few of those restrictions are trivial. They include inclination restrictions, large plane changes required to achieve lowinclination orbits from high-latitude launch sites, large plane changes required to transfer from GTO to GEO when launching from certain ranges, and lowfrequency launch opportunities for missions that require phasing such as those involving a rendezvous with another spacecraft already in orbit. Inclination Restrictions. Inclination restrictions stem from range safety considerations. To understand these restrictions fully, it is first necessary to

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understand two concepts: (1) transfer orbits and (2) instantaneous impact-point tracks. Transfer Orbits. Transfer orbits are intermediate orbits established by the various stages of a launch vehicle that provide a path to the final desired orbit. The transfer orbits for early stages are mostly suborbital, meaning that some portion of the orbit intersects Earth’s surface. The most efficient way to transfer between two orbits is to apply thrust at opposite apses. An application of thrust in the right direction at the perigee of the initial orbit will raise the apogee. Coasting to the new apogee and applying thrust (again in the appropriate direction) at this apsis will then raise the perigee. This provides a stair-step approach to raising the altitude of a vehicle’s orbit. The ascent of a launch vehicle from launch to orbit follows a similar trend with one critical caveat. The impulse of initial stages is usually not sufficient, individually, to raise the perigee above Earth’s surface. This means that using the optimal Hohmann transfer approach would bring the launch vehicle back to Earth before another transfer burn could be made. As a result, initial launch vehicle stages usually apply their thrust at places within a transfer orbit other than the apses and usually always on the ascending side of the orbit. Consider a modest three stage, ground-launched rocket launching into a circular low Earth orbit as an example. Before launch, the vehicle is effectively sitting at the apogee of an orbit (Fig. 7). If the surface of Earth were not present to support the rocket, it would be drawn downward along a path that would take it closer and closer to Earth’s center before swinging back to an apogee altitude equal to the radius of Earth. This is essentially the first of several transfer orbits and the rocket has not even been launched. When the rocket lifts off, it applies its thrust at an apsis, but in a direction that is perpendicular to the initial velocity vector of the rocket, which itself is in the direction of Earth’s rotation. During the first burn, the vehicle slowly tilts over so that the thrust is applied in a direction that is increasingly parallel to Earth (Fig. 8). This has the effect of increasing both the apogee and perigee. The perigee will most likely still be suborbital at the end of the burn. The apogee will be increased sufficiently that the launch vehicle Path of rocket without Earth's surface

Center of the Earth Earth

Figure 7. Path of rocket without Earth’s surface. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

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Eventual impact point of Stage 1

Jettison of 1st Stage

New transfer orbit after Stage 1 burn

Figure 8. Path of rocket after launch. This figure is available in full color at http:// www.mrw.interscience.wiley.com/esst.

can coast up to a location near the new apogee, following the first stage burnout, and ignite the second stage. The key consideration here is that the second stage will be ignited near but not at the apogee. Again, this is not the most energyefficient way to transfer orbits, but it is necessary because the opposite apsis is still below Earth’s surface, and the second stage may not have sufficient impulse to raise it above the atmosphere. Igniting the second stage at a location other than the apogee again has the effect of raising both the perigee and the apogee. In this case, because only one stage is left, the burn is designed to raise the apogee to the desired altitude of the final orbit. After the second stage burns out, the vehicle coasts up to the new apogee and ignites the third stage. This will raise the perigee up to the final orbit altitude without changing the altitude of the apogee. Impact-Point Tracks. By always burning on the ascending side of the trajectory and iteratively raising the apogee while the transfer orbit remains suborbital, anything jettisoned before the final burn will reenter the atmosphere and either burn up or impact Earth’s surface. As the burn of each stage progresses, the point at which the transfer orbit intersects the Earth extends farther and further downrange until, at some point late in the final burn, there is no longer a point of intersection. These points of intersection comprise the instantaneous impact-point track. Clearly, as the vehicle is coasting, the instantaneous impact point does not change. Conversely, during a motor burn, it is constantly changing and each point represents the location of impact on Earth if, in fact, the thrust were to be instantly terminated either by design or due to some sort of failure. It is this impact-point track and the need for it to avoid populated areas that is a primary source of inclination restrictions from various ranges. For any rocket launch, whether it be space-based, suborbital, groundlaunched, ship-launched, or air-launched, the public-safety considerations that must be satisfied are very stringent. Those stages of a rocket that are jettisoned before reaching orbit should avoid land. And no launch vehicle whose impactpoint track nominally crosses land can risk a casualty among the public with a probability of greater than 30 in a million. Calculating the expectation of a

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casualty depends on many factors, including the reliability of the launch vehicle (e.g., how many failures it has had in the past), the density of the population being overflown, and the speed with which the instantaneous impact-point track crosses over a populated region. Late in flight, the distance between successive impact points increases dramatically and reduces the risk to the population below. This is why it is generally more permissible to overfly populated regions far downrange than it is early in flight. For instance, the risk to a populated region in Africa from a rocket launched at the Eastern Range would, in general, be less than the risk posed to an area with the same population density overflown in the Caribbean. This is not to say that overflight of any part of Africa is acceptable. There are some extremely high population densities in Africa, especially along the west coast of northern Africa, which are avoided at all costs. And it is this very consideration that constrains the paths of many launch vehicles from the existing ranges. The key land masses that must be avoided early in flight for vehicles launching from the Eastern Range include the entire eastern seaboard of the United States when launching on an ascending pass (northerly direction) and the Caribbean and South America when launching on a descending pass (southerly direction). For maximum performance from any given launch vehicle, this restricts the range of inclinations achievable from the Eastern Range to between roughly 28.51 and 511 for ascending passes and between 28.51 and 401 for descending passes. Clearly, inclinations outside this range would be achievable if plane changes were instituted, but that has the disadvantage of reducing the maximum available performance for any given launch vehicle. Higher inclinations are available from the Western Range but restrictions still exist there due to Hawaii, islands in the South Pacific, and the western coasts of both North and South America. When the inclinations from both the Eastern and Western Ranges are combined (assuming direct injection), a block of inclinations is unavailable without plane changes and subsequent reductions in weight-to-orbit capabilities. For small payloads with limited budgets that require an inclination outside what is directly available from the existing ranges, the cost of launching on the heavy-lift launchers that can execute the necessary plane changes can be prohibitive. And reducing launch costs by flying as a secondary or even tertiary payload is advantageous only in the rare event that a primary payload can be found that requires the same final orbit. For these customers, Pegasus and Shtil provide an alternative due to their relatively low cost, mobility, and self-contained launch infrastructure. Sea Launch provides a similar alternative for the heaviest satellites that are intended for either GEO or low Earth orbits. Plane Changes Required to Achieve Low Inclinations. The inclination of an orbit represents the angle between the equatorial plane and the orbital plane around Earth. This also happens to be similar to the definition of lines of latitude. It is no coincidence then that the maximum latitude of the ground track for any object in space is roughly equivalent to the inclination of the object’s orbit. The only reason that the maximum latitude is not exactly equal to the inclination is because Earth is not a perfect sphere. Conversely, this implies that the minimum inclination attainable by a launch vehicle is roughly equivalent to the latitude of the location from which it is launched. The maximum is 1801 minus

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the latitude of the launch point. This leads to the important conclusion that the only latitude from which all inclinations are directly accessible is 01 (the equator). The Eastern Range is at a latitude of roughly 28.51. Therefore, the minimum inclination attainable without plane changes is roughly 28.51. Lower inclinations can be achieved by launching into any available inclination, achieving a preliminary orbit, and then making an inclination correction burn when the satellite is over the equator or at any latitude that is numerically less than the desired inclination. The significant disadvantage of this process is that inclination changes while in orbit require a great deal of energy. The larger the change in inclination required, the more energy must be expended. Depending on the final orbit desired, this usually requires an additional stage to correct the inclination and achieve the final orbit. The most common recipient of this type of orbit maneuver is a satellite headed to geosynchronous orbit. However, there are low Earth orbit payloads that require low inclinations as well. The ability of Pegasus and Shtil to move the drop point to a latitude from which such energy-intensive plane changes would not be required permits smaller launch vehicles to achieve the same orbit from lower latitudes that larger vehicles can achieve from higher latitudes. The difference in cost, complexity, and performance can often mean the difference for some customers between launching or not. Some launch locations maintained by other countries are at significantly lower latitudes than those in the United States. For some customers, such ranges can provide the necessary services. However, many satellites in the United States, especially government sponsored, are required to contract with a U.S. launch service provider and use a U.S. controlled range. Phasing. An object’s orbit is essentially a locus of points that defines the path of the satellite. Those points define a plane that goes through the center of Earth. To define an object’s precise position within an orbit, that plane and every position in it is defined with respect to both Earth and a coordinate system, one of whose axes always points toward the vernal equinox. Every position of a satellite as it orbits Earth is defined in terms of an epoch (time), the semimajor axis, and eccentricity, measured from Earth’s center, inclination and argument of perigee, which are both referenced to Earth’s equator, and the right ascension of the ascending node, which is referenced to the vernal equinox frame. A rendezvous between two objects in space involves a series of maneuvers designed to make the orbital elements of both objects the same, hence confirming the fact that they have, in fact, become a single object orbiting Earth. Just as motor burns can raise or lower the perigee or apogee of an orbit or change the inclination, so too can motor burns be used to change every orbital element that defines a satellite’s motion. However, changing some of those elements, especially those that require plane changes, requires large amounts of energy, and they are considered ‘‘expensive’’ in the parlance of orbital mechanics. One way to avoid paying the high price of actively changing the orbit of a satellite with a motor burn is to do it passively through the aid of various external forces. Several naturally occurring forces cause every orbital element to change over time. These include atmospheric drag, solar radiative pressure, the gravitational attraction of the Moon, Sun, and planets, and the nonuniform gravitational forces due to Earth’s oblateness. These forces can be used to one’s advantage when planning a rendezvous mission. However, some changes resulting from these forces can take

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a very long time to reach significant levels. This means that the initial differences between the rendezvousing satellite and the target must be initially small to avoid spending too much unproductive time in orbit. This can be accomplished by simply timing the launch appropriately so that at the time of orbital insertion, the satellite that has newly arrived in orbit is very close to the orbital plane of the target satellite. To accomplish this maneuver, the launch must occur when the target satellite passes almost directly overhead. It also must be passing in the same direction as the intended launch. In other words, if the satellite being launched is to head off in a southerly direction (along the descending pass), the target satellite must be overhead and also on its descending pass as well. Otherwise, the two satellites will end up with right ascensions that are 1801 apart which would be excessively expensive (either in terms of time or energy) to correct once in orbit. For ground-launched vehicles, the wait between successive passes of the target satellite could be as much as several days, depending on the target orbit because the distance between ground tracks on successive passes depends on the period of the orbit, which depends on the orbit’s semimajor axis. Clearly, the ground track of an object that requires only 90 minutes to orbit Earth will be more closely spaced than the ground track of an object whose period is several hours. These ground tracks will pass to the east and west of the given launch site on a daily basis, but the distance between the ground track and the launch site will only be minimized by a periodicity of the order of days. Mobile assets, however, can eliminate the wait by essentially choosing a launch point that is ideally suited for a rendezvous. Instead of waiting for the ground track to come to the launch point, the launch point is moved to the ground track. In this way, the launch opportunities can be reduced from one every two to three days to at least once a day if not twice a day, if the launch vehicles have the flexibility to launch on both ascending and descending passes. Consider an example of a satellite being launched by a Pegasus XL to rendezvous with a satellite currently in orbit at an altitude of 400 km circular. A normal ground-launched vehicle would require a wait of about 2 days between successive launch attempts. However, the mobility of Pegasus permits two launch opportunities every day, which is graphically represented in Figs. 9 and 10. Two key assumptions need to be kept in mind when viewing these figures. The first is that the maximum range of the Pegasus carrier aircraft is roughly 1000 nmi. This includes a captive carry to the launch site, an aborted launch, and a return to base with Pegasus still attached. The second assumption is that for launches that do not require the full advantage of Pegasus’ mobility, the standard launch point for Pegasus out of the Eastern Range is 281 N, 281.51 E. The vertical axes in Figs. 9 and 10 represent the difference in argument of latitude between the two satellites (the angular separation within the same orbital plane). The horizontal axis represents the launch point as the difference in degrees from the nominal point listed before. The diagonal lines represent the difference in argument of latitude for each day in the first week of October, which was chosen simply as an example. Figure 9 represents the difference in argument of latitude for northerly launches (launch along the ascending pass). Figure 10 represents the difference in argument of latitude for southerly launches (launch along the

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300

Difference in argument of latitude (deg)

*** Note: Labels reference day of month*** 3

200

5

100 2

7 0

4 1

_100

6

_200 _300 _360 _400

2

0

Figure 9.

3

10 6 8 12 14 16 Change in longitude from drop point 1 (deg)

18

20

Graph of difference in argument of latitude for northerly launches.

descending pass). The horizontal lines simply demarcate zero angular separation between the two satellites. The intersection of a diagonal line with a horizontal line defines a drop point within the range of the Pegasus carrier aircraft from which Pegasus can be 300 *** Note: Labels reference day of month*** Difference in argument of latitude (deg)

200 2 7

100 4 1

0

6 3

_100

5

_200 _300 _360 _ 400 0

Figure 10.

2

10 6 12 3 16 8 14 Change in longitude from drop point 1 (deg)

18

20

Graph of difference in argument of latitude for southerly launches.

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launched and effectively deliver its satellite to the front door of the target satellite at the time of orbital insertion. Realistically, this is not how a rendezvous would normally be achieved. Ideally, the satellite being launched would be placed in a temporary parking orbit slightly below and behind the target satellite. Over the course of several orbits the distance separating the two objects would be slowly decreased using several controlled burns of the satellite just placed in orbit. This would imply that a drop point is needed not to achieve 01 difference in argument of latitude but some finite value. The example is still valid. Simply shift the horizontal lines up or down until the desired difference in argument of latitude is matched. Again, an intersection between a diagonal line and the horizontal line defines a launch point within the range of the Pegasus carrier aircraft that would result in the desired difference in argument of latitude. As can be seen from Figs. 9 and 10, every day except two in the first week of October provides two launch opportunities. A southerly launch on the 3rd does not provide a drop point within the range of the carrier aircraft that will achieve the desired result. However, a drop point can be found on that day if the launch is along the ascending pass instead. Likewise, a suitable drop point cannot be found within the range of the carrier aircraft on the 6th of October when launching along the ascending pass, but one can be found if launching in a southerly direction. The same qualitative results would be obtained for any other time frame. The quantitative results might be slightly different. For instance, instead of having only one launch opportunity on the 3rd and 6th it may be the 4th and the 7th. But the end result is the same. The mobility of Pegasus and, by definition, Sea Launch, and Shtil provides ideal rendezvous launch opportunities at least once a day and in most cases twice a day. Clearly there are disadvantages with all of these mobile assets. Pegasus is limited in its size due to the restrictions of the L-1011 and, more importantly, the mechanical limitations of the hooks that hold the vehicle to the plane. Sea Launch has somewhat of a temporal disadvantage in that it requires almost 2 weeks to travel to the launch site. Those problems are exacerbated for Shtil because its home port is farther north. Nonetheless, for some specific missions, the mobility and flexibility that are provided by these unique space-launched assets provide valuable supplemental services to the fleet of existing groundlaunched vehicles.

BIBLIOGRAPHY 1. 2. 3. 4.

Powell, J. The China Lake Launches. Air and Space, pp. 367–378, Feb/Mar 1997. http://www.state.gov/www/global/arms/treaties/abm/abm2.html. Pegasus Users Guide, Release 5.0, Orbital Sciences Corporation, August 2000. Isakowitz, S.J. International Reference Guide to Space Launch Systems, 3rd ed. AIAA, Washington, DC, 1999. 5. Sea Launch User’s Guide, rev. B, Boeing Commercial Space Company, July 2000.

DALE FENN Orbital Sciences Corporation Dulles, Virginia

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APOLLO 17 AND THE MOON Apollo 17 was not the last flight of humans to the Moon. This writer was not the last human being to step on the lunar surface. More lunar exploration and even lunar settlement will occur, baring the future stagnation or disappearance of our civilization. Exploration and scientific investigations in the earth sciences are rarely complete, particularly for studies related to a specific field site. A long hiatus between field investigations may occur, but other forms of investigation, directly or indirectly related, continue. Apollo 17’s field study of the Valley of Taurus-Littrow on the Moon in 1972 and subsequent examination of its significance to our understanding of the origin and evolution of that small planet and of our own constitute a good example of these facts of scientific life. As the third of the specifically ‘‘science’’ missions to the Moon in the twentieth century, Apollo 17 actually became the last lunar landing of the Apollo Program in September 1970 (1) rather than on 11 December 1972 when the mission reached TaurusLittrow. The National Aeronautics and Space Administration (NASA) and the Administration of President Richard M. Nixon, with the acquiescence of the Congress, had concluded that no further planned amortization of the American taxpayer’s investment in deep space exploration would be undertaken. As historically naive a political decision as this may seem today to some, it did not prevent the achievement of one of the Program’s major goals—gaining a firstorder understanding of the Moon and its relationships to the terrestrial planets. This became one of the primary historical legacies of the post-World War II generation. Apollo had evolved quickly and radically toward increased scientific emphasis after Neil Armstrong first stepped on the Moon on 20 July 1969. Its purpose changed from the completed goal of meeting President John F. Kennedy’s challenge (2) to land ‘‘men on the Moon and return them safely to Earth,’’ to an objective of increasing human knowledge about the Moon and space. This would be done to the maximum extent possible using the technological and operational systems in hand and reasonable extensions of that capability. This shift in emphasis occurred smoothly and rapidly thanks to the foresight of senior NASA managers such as George M. Low, Apollo Spacecraft Program Manager; Robert Gilruth, Director of the Manned Spacecraft Center; Eugene Kranz, Chief of the Flight Control Division; Maxime Faget, Chief Engineer of the Manned Spacecraft Center; and General Samuel Phillips, Director of the Apollo Program. As early as the spring of 1969 (3), scientific packages were being enhanced, adding new experiments and improving old ones. Astronaut training in field geology, overseen by the author for the Astronaut Office, was altered to consist of field simulations (4) at geologically relevant sites using mission-specific equipment and procedures. These scientific training exercises taught pilots the art of geologic observation, sampling, and documentation and also put that learning in the context of real geologic problems related directly or indirectly to those they would encounter on the Moon. In addition to the U.S. Geological Survey’s Principal Investigators for field geology, Eugene M. Shoemaker, Gordon A. Swann, and William R. Muehlberger (also of the University of Texas), world-renowned Earth scientists who doubled as outstanding teachers were given increased

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access to mission planning, mission operations, and astronaut training. These new participants included Richard H. Jahns (Stanford University), Robert P. Sharp and Leon T. Silver (California Institute of Technology), James B. Thompson and James Hays (Harvard University), and Gene Simmons and William Brace (Massachusetts Institute of Technology). Also important to science was an increase in the capability of the Lunar Module (5) so later missions could include heavier scientific payloads, a lunar roving vehicle, and greater consumables, thus, longer time on the Moon. These augmentations meant a large increase both in time for exploration (22 hours for Apollo 17 versus a total of 19.4 hours for Apollo 11, 12, and 14, combined) and in distance traveled (35 km for Apollo 17 versus B5 km combined walking distance for Apollo 11, 12, and 14). The Apollo 11, 12, and 14 missions were flown largely under the original payload, training, and operational constraints imposed by the race to the Moon and the conservatism necessary for success in that race. These missions still managed, however, to produce remarkable suites of samples, photographs, and observations in addition to giving the Apollo team the operational confidence to land at more challenging but scientifically more interesting locations away from the lunar equator. In spite of the operational limitations, the analysis of samples and other information returned from the first three landing sites rapidly increased the understanding of the Moon and its history. Ironically, the Apollo 13 mission, which failed to land on the Moon, set the stage for the even more spectacular scientific returns from the last three landings, Apollo 15, 16, and 17. The crew and backup crew of Apollo 13 had embraced the new training emphasis on field geology and encouraged the Apollo 15 crew to follow suit. Apollo 13’s backup crew, already convinced that a science focus was important, was assigned to fly the Apollo 16 mission. Finally, the designation of a scientist and geologist as the Lunar Module Pilot on Apollo 17 assured that all of the last three missions truly would be ‘‘The Great Voyages of Exploration’’ (6). Due to the foreknowledge that Apollo 17 would be the last of the Apollo series, selection of its landing site became a contentious issue (7) among lunar scientists and between lunar scientists and operational planners. The usual candidates for landing sites reappeared: crater floor and central peak opportunities for deep sampling of the lunar crust, like the impact crater Copernicus; possible volcanic features, like the Davy Crater Chain and dark material in Alphonsus; and highland areas such as the rim of the crater Tycho and an area ‘‘Southwest of Crisium.’’ Even a farside landing in the basin Tsiolkovskiy was given brief consideration due to the efforts of the author (8). Eventually, however, the scientists became increasingly interested in an unnamed, 2300-m deep, 50-km long valley, radial to the 740-km diameter circular basin, Serenitatis, that cut through the Taurus Mountain ring near the crater Littrow. This Valley of ‘‘Taurus-Littrow,’’ however, was not a favorite of the operational mission planners. In spite of the pinpoint landing accuracy they had demonstrated on all previous missions since Apollo 11, the narrow valley, the mountainous approach, and the high valley walls gave the planners pause. Their legitimate concerns were compounded by the relatively short time, only 14 minutes, for navigational updates after acquisition of communications from the lunar module, Challenger, as its last orbit before landing carried it around the Moon from the farside. Initially, trajectory calculations indicated that three-sigma errors, the normal

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Figure 1. Night launch of the Saturn V rocket carrying the Apollo 17 Mission to the Moon at 12:40 A.M., 7 December 1972 (courtesy of NASA). This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

extremely conservative planning limit, might result in hitting the side of the northern mountain wall. Gradually, however, refinements in navigational techniques for the mission and the inevitable synergistic give and take that so characterized Apollo interactions narrowed the three-sigma errors to about 1-km, the limit where all agreed that Taurus-Littrow could be the selected site. Thus, in late February 1972, only 9 months from launch, Taurus-Littrow was approved as the exploration site for Apollo 17 (9) (Fig. 1).

The Apollo 17 Mission The Valley of Taurus-Littrow (Fig. 2) offered four major benefits as the last Apollo landing site, taken in the context of a final test of then current hypotheses related to the origin and evolution of the Moon. First, photogeologic analysis indicated that Taurus-Littrow provided access to a three-dimensional window into a mountain ring created by the Serenitatis large basin-forming event, by now well established as the result of a giant impact of an asteroid or comet. Second, major units of mare basalt and older nonmare rocks would be within easy reach of roving vehicle traverses. Third, a mantle of dark, possibly young volcanic debris partially covered the region as well as portions of the valley, and craters of a range of depths

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Figure 2. The Valley of Taurus-Littrow as seen from the Commander’s window on the left side of the Lunar Module Challenger on the orbit of the Moon before landing. The view is approximately west northwest, looking toward the Serenitatis Basin (courtesy of NASA). This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

penetrated this debris and the underlying basalt. And, fourth, the valley lies about 600-km north and 200-km east of the Apollo 11 and Apollo 15 sample areas, respectively, adding significantly to our exploration coverage of the Moon’s nearside. The Lunar Module crew of Apollo 17, Commander Eugene A. Cernan, and the writer as geologist and Lunar Module Pilot, conducted 22 hours of field exploration and experiment deployment in the Valley of Taurus-Littrow between 11 and 14 December 1972. During this period, the crew investigated, photographed, and sampled 11 major field locations. We traversed, observed, and sampled more than 35-km of the valley floor and obtained and documented 120-kg of samples from 97 major boulders and 75 other lunar materials. We took 2200 documentation photographs and deployed the 11 experiments of the Advanced Lunar Science Experiment Package (10). The crew had trained together for 15 months before launch, several days a month consisted of simulated traverses at field sites illustrating one or more of the types of geologic problems expected on the Moon and specifically at Taurus-Littrow. Combined with the geologic experience of the author, the organization and flexibility of the exploration plans (11–13), and the close cooperation of the science team in direct support on Earth, this training gave a stronger scientific foundation to Apollo 17’s exploration that had been possible during previous Apollo missions (14,15).

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The North and South Massifs constitute the major structural boundaries of the Valley of Taurus-Littrow. Slopes of an approximately constant 251 flank the Massifs, rising 2000 and 2300 m, respectively, above the valley floor. Discontinuous, but roughly horizontal exposures of thick sections of crustal rocks that predate the major mare basalt eruptions exist on the steepest upper slopes. These outcrops or near outcrops create numerous fields of exposed rock from which tracks lead downward to some of the sampled boulders at the base of the Massifs. Interlocking domes called the Sculptured Hills constitute Taurus-Littrow’s northeastern wall and have concentrations of boulders apparent only on the inaccessible upper slopes of these hills. The valley floor consists of an undulating, highly cratered, relatively flat surface, covered largely by broken and pulverized basalt. One group of the cluster of craters surrounding the spot where Challenger landed, lies on a ray of secondary ejecta from the crater Tycho (16) 2000-km to the southwest. The largest of these craters is about 600-m in diameter. An older cluster of craters of about the same range of sizes cut into the floor northeast and to the west of the landing point, near the base of the North Massif. An irregular fan of material, the light mantle, projects northeast from the base of the South Massif. Finger-like projections of this fan reach out as much as 6-km from the Massif. Premission photographs suggested that a mantle of dark material covers the valley as a whole, including portions of the surrounding mountains. All surfaces are composed as pulverized debris called ‘‘regolith’’ (17), consisting largely of fragments of the bedrock below mixed with dark mantling material and other materials thrown into the area by more distant meteor impacts or introduced by volcanic eruptions younger than the bedrock. Scientific activities in Taurus-Littrow (Fig. 3) began with the deployment of the experiments constituting the sixth and final Apollo Lunar Surface Experiments Package (ALSEP). This package had been enhanced to have a design life of 2 years rather than one (18). In connection with drilling holes for the heat flow experiment, two cores through the upper 3.2-m of the central valley regolith were obtained. Despite minor interruptions to work on technical problems with the ALSEP, about a third of the first excursion (extravehicular activity or EVA) and most of the second and third excursions concentrated on the planned traverses and exploration. Actual traverses followed this plan closely (19–22) except for a curtailed first traverse that only reached a point on the rim of Steno crater rather than reaching the original objective of Powell and deletion of the third traverse’s Station 10 at Sherlock. Investigations of the basaltic mare materials south of the landing point began on the first excursion. On the second day, the traverse went west to sample premare materials at the base of the South Massif (Station 2). We then worked back over the light mantle deposit (Stations 2A and 3); to the dark, possibly volcanic crater, Shorty (Station 4); over the contact between light mantle and mare; and finally into the basalt boulder field surrounding the 100–150 m deep crater, Camelot (Station 5). The third day started with a long study of large boulders at the base of the North Massif (Stations 6 and 7), followed by sampling in the regolith at the base of the Sculptured Hills (Station 8), and a final stop at another possible volcanic crater, Van Serg (Station 9). Along each lunar rover traverse, we periodically sampled the surface of the regolith across various geologic units, deployed explosive charges for the active seismic profiling experiment, monitored the receiver for the surface electrical properties experiment,

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Figure 3. Apollo 17 exploration area in the Valley of Taurus-Littrow showing the landing site, exploration stations (numbers), and general traverses (solid lines) (courtesy of NASA). This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

and obtained readings from the traverse gravimeter (23). The principal sampling tools used included a rock hammer, a pair of long handled tongs, 35 and 70 cm core tubes, a long handled scoop (also used for trenching), and a supply of prenumbered Teflon sample bags (Fig. 4).

Impact Cratering Almost everything we think we know about to the Moon must be viewed through the filter of impact cratering effects (Fig. 5) that have dominated lunar history from its origin to the present (24–29). The impact of comets, asteroids, meteors, micrometeors, dust, and energetic atomic and nuclear particles have modified the surface and near-surface expression of all of the internally generated processes that contributed to the present physical nature of the Moon. The secondary effects of each impact have magnified the importance of these impacts. Most comet, asteroid, meteor, and micrometeoroid impact velocities are between 13 and 18-km/s, and some are as high as 70-km/s, giving target pressures at the point of impact of several hundred Gpa (gigapascal). Extraordinary amounts of heat per unit mass are released as conversion of kinetic energy into forward and rearward shock waves takes place almost instantaneously. The amount of extralunar material that can be identified in regolith samples returned to Earth

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Figure 4. The author as a well-equipped astronaut on the Moon during the Apollo 17 mission in the Valley of Taurus-Littrow. He is using ‘‘the rake’’ sampling device to sift rock fragments from the finer portions of the regolith and has a 70-mm Hasselblad camera mounted on his chest (courtesy of NASA). This figure is available in full color at http:// www.mrw.interscience.wiley.com/esst.

indicates that about 98% (30) to 99.7% (31) of all but the larger projectiles is melted, vaporized, or ionized, and returned to space. The general characteristics of lunar impact craters as a function of diameter are summarized in Table 1. Processes associated with cratering and space radiation have created a well-defined zone of debris that covers essentially the entire Moon; its thickness depends on the length of exposure of a specific geologic unit or feature. This zone is called the ‘‘regolith,’’ a terrestrial term also used for the Moon. Essentially all the samples returned from the Moon by Apollo have come from the regolith or from rocks incorporated within it. It has been defined as ‘‘the layer or mantle of fragmental and unconsolidated rock material, whether residual or transported and of highly varied character, that nearly everywhere forms the surface of the land and overlies or covers bedrock. It includes rock debris of all kinds, including volcanic ashylunar regolith consists [largely] of particles o1-cm in size although larger cobbles and boulders, some as much as several meters across, are commonly foundy.much of the pulverized material is melted and welded together to produce breccias (fragmental rocks) and impact melt rocks, which make up a significant portion of the regolithy’’(32,33). A particularly important part of the lunar regolith consists of aggregates of rock, mineral, and glass fragments,

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Figure 5. The 95-km diameter impact crater Copernicus as seen from the Apollo 17 Lunar Module Challenger after its departure from the Moon on 14 December 1972. Data from the Apollo 12 mission indicate that Copernicus formed about 900 m.y. ago (courtesy of NASA). This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

called agglutinates, held together by impact melt glass. Recently, it has been shown that on the nanometer scale, iron metal particles accreted on and formed in the rims of regolith grains significantly affect optical and magnetic properties (34–36). Further, the lunar regolith contains embedded solar wind gases, meteoritic material, and isotopic products and crystal structure damage produced by solar and cosmic radiation. The average depth of the regolith in a given area reflects the age of the underlying bedrock. Lateral mixing of material derived from adjoining bedrock units is a function of the age of the separating contact.

Origin and Evolution A ‘‘standard’’ or ‘‘conventional’’ hypothesis for the origin and evolution of the Moon evolved significantly during the last three decades of the twentieth

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Table 1. General Characteristics of Lunar Craters as a Function of Sizea Diameter range examples o10-m

General characteristics

*

*

*

*

*

B10-m to B100-m Van Serg and Shorty Craters in Taurus-Littrow

*

*

*

*

*

*

B100-m to B10-km Taruntius; Camelot Crater in Taurus-Littrow

*

*

*

*

*

*

B10-km to o300-km Copernicus

*

*

*

*

*

*

*

Craters norm ally do not penetrate the regolith. Depth to diameter ratio variable. Glass discontinuously lines shallow pits in the center of fresh craters. Mineral grains shattered around small craters on solid rock (zap pits). Deep pits (B1/3 the crater diameter) in the center of some craters. Craters normally penetrate mare regolith if above 20-m diameter. Depth to diameter ratio about 1:3 to 1:4 for fresh craters. Inner benches common if target material stratified. Regolith breccias present inside and on the ejecta blankets of young craters. Ejecta blankets extend to about one crater diameter. Target strata are overturned, but original vertical sequence is preserved in ejecta blanket. Both transient and initial steady-state craters are hemispherical and have circular and raised rims. Depth to diameter ratio about 1:3 to 1:4 for fresh craters. Impact breccias present inside and on the ejecta blankets of young craters. Ejecta blankets extend to about one crater diameter. Secondary impact cratering significantly modifies surface features out to many crater diameters from the edge of continuous ejecta. Target strata are overturned, but their original vertical sequence is preserved in ejecta blanket. Transient crater approaches hemisphere and has a circular raised rim and probably is lined with a shell of impact melt. Initial steady-state crater has a flat floor and central mound or peak. Initial steady-state crater walls have many stepwise benches (slump landslides) on walls. Hummocky crater floors and the depressions on wall benches and near-rim ejecta blankets of larger craters have indications of pools and flows of impact melt. Ejecta blankets extend to about one crater diameter. Target strata are overturned, but their original vertical sequence is preserved in ejecta blanket. Secondary impact craters, crater clusters, crater chains, and herringbone crater chains extend several thousand kilometers beyond edge of continuous ejecta.

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Table 1. (Continued) Diameter range examples 4300-km (basin) Orientale

General characteristics

*

*

*

*

*

*

*

*

*

*

*

Transient crater depth to diameter ratio decreases with increasing size as lithostatic pressures compete with explosive pressures. Transient crater has an increasingly flat trapezoidal crosssection and increasing diameters. Transient crater has a flat floor, a circular raised rim, and probably is lined with a thick shell of impact melt. Initial steady-state crater has a fractured, flat floor and central ring or partial ring of peaks. Initial steady-state crater walls have many wide, stepwise benches (slump landslides) on walls. Floors and depressions on wall benches and near-rim ejecta blankets have indications of large pools, mantles, and flows of impact melt. Impact melt also injected into target materials. Ejecta and debris flow blankets extend beyond one crater diameter. Two to six rings of mountains outside transient crater rim around basins 4400-km in diameter. Target strata sequence is not well preserved in ejecta blanket due to extensive mixing of ejecta during flow. Within one crater diameter of the final steady-state rim, there is a continuous deposit of melt breccia, possibly several hundred meters thick at the rim of the larger basins. Secondary impact craters, crater clusters, crater chains, and herringbone crater chains extend beyond the edge of continuous ejecta and debris flows and reach thousands of kilometers and probably around the entire Moon to the basin antipode.

a

Ref. 620.

century. This hypothesis currently holds that the Moon formed about 4.57 billion years (b.y.) ago by the aggregation of material produced during a giant impact between the very young Earth and a Mars-sized asteroid; most metallic coreforming material remained with Earth (37–43). Such an origin could explain the high angular momentum of the Earth–Moon system and at least some of the lunar geochemical constraints related to iron, volatile, and alkali elements other than potassium (44). Soon after or during lunar aggregation, lunar core formation occurred (45,46), and a Magma Ocean developed on its surface (47,48). The lunar Magma Ocean largely crystallized within 50 million years (m.y.) of the creation of the solar nebula and, at the same time, differentiated due to contrasts in mineral densities into an olivine-pyroxene dominated mantle and a 60–70-km thick Caplagioclase-rich crust (49,50). Late in this differentiation process, potassium, rare-earth elements, phosphorous, and thorium-rich residual liquid (urKREEP) (51) accumulated beneath the crust, largely in the region beneath what is now the Procellarum basin (52–56). Late ilmenite-rich cumulates (57) sank toward

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the base of the less dense olivine and pyroxene cumulates carrying some urKREEP material with them (58). Intrusive and extrusive basaltic magmatic activity began soon after the Magma Ocean crystallized and, before the main sequence of mare basalt eruption began at about 3.8 b.y., produced the magnesium-rich suite of plutonic rocks (Mg-suite) (59–61), KREEP-rich basalts (62), and the cryptomaria (63–66). At about 3.85 b.y., a concentrated bombardment of the crust took place that produced most or all of the B50 basins greater than 300-km in diameter visible today as well as most other observed cratering effects in the lunar crust (67–71). The effect of this late lunar ‘‘cataclysm’’ was to reset the ages of all crustal impact glasses yet studied (72,73). South Pole-Aitken basin is probably the only basin 41000-km in diameter to form during this late bombardment (74); the Procellarum basin is an artifact of the superposition of several smaller basins (75). A global magnetic field and presumably a circulating fluid metallic core were present at least between 3.9 and 3.8 b.y. (76–78). The core is now between 300 and 400-km in radius (79–81). Between 3.9 b.y. and about 1.0 b.y., mare basalts and basaltic pyroclastic materials erupted, largely on the nearside of the Moon (82–84). Major features on the Moon have been little modified subsequently other than the development of several meters of impact-generated regolith on most surfaces (85). Although some aspects of this conventional hypothesis of lunar origin and evolution are attractive and probably correct, as will be discussed throughout this article, numerous difficulties exist in reconciling a number of its implications with everything we think we know about the Moon (86–89). Some of the major questions that can be raised with the conventional hypothesis are as follows: 1. Was the Moon formed as a result of a giant impact on Earth immediately after Earth’s accretion or was it formed independently and later captured? 2. Did core formation in the Moon and other terrestrial planets occur immediately after their accretion or was it delayed by the existence of a silicate protocore? 3. Did thermal convection and/or impact-induced splash cooling play a significant role in the crystallization and differentiation of the Magma Ocean? 4. Did the Moon’s Magma Ocean’s late ilmenite-rich cumulate sink near the base of the cumulate pile globally or only in response to local destabilization by the formation of a few extremely large impact basins? 5. Was the Moon’s Magma Ocean’s residual liquid (urKREEP) initially concentrated beneath the Procellarum Basin or distributed in a spherically uniform shell under the lunar crust? 6. Was the global thermal insulation effect of the impact-generated megaregolith of the lunar highlands critical to the later formation of the magmas that formed the basaltic maria? 7. Is the Procellarum Basin a consequence of the merging of several smaller basins or of a single extremely large impact? 8. Were the one or more extremely large impact basins and the B50 large basins on the Moon the result of a ‘‘cataclysm’’ of impacts at about 3.85 b.y. or of a sustained bombardment lasting about 400 m.y.?

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9. Was melting of the mantle due to pressure release (90) after large basin events significant in generating of magmas related to the Mg-suite of lunar rocks? The scientific results of the Apollo 17 mission can now be viewed in the context of the conventional hypothesis and questions about that hypothesis from the perspective of more than 40 years of modern study of all of the Apollo missions and other lunar investigations. These later investigations have included telescopic and photogeologic mapping of the lunar surface, Apollo sample analyses, automated missions that both preceded (91) and followed Apollo, and the remarkable thought and computer modeling that has been stimulated by the collected data. Broadened multidisciplinary discussions of lunar origin and evolution are assisted by a descriptive formulation of the formative stages of lunar evolution as an augmentation of the traditional time-stratigraphic approach (92). The term ‘‘stage’’ is not used below in the normal time-stratigraphic sense (93). Rather ‘‘stage’’ is used in a more general sense for overlapping periods of lunar history that have definable but somewhat arbitrary beginnings and endings due in large part to the current incompleteness of information about the absolute ages of lunar events. Thus, the evolution of the Moon as a small planet (94–98) can be descriptively summarized as follows (Plate 1):

Major stages of lunar evolution 1

Beginning (large earth impact or capture)

2

Magma ocean/crust and upper mantle form

3

Cratered highlands/very large basins

Stage

4

Large basins Old large basins /crustal strengthening

4a

Young large basins /core formation

4b

Cataclysm ? 5

Basaltic maria

?

?

Cryptomaria

5a

5b

Maria

Ti-rich 5.0

4.0

Ti-poor

3.0 2.0 Billions of years before present

? Ti-rich 1.0 Red = Major uncertainty

Plate 1. Major stages of lunar evolution. This figure is available in full color at http:// www.mrw.interscience.wiley.com/esst.

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Stage 1: Stage 2: Stage 3: Stage 4: Stage 4A: Stage Stage Stage Stage Stage

4B: 5: 5A: 5B: 6:

29

Beginning [Pre-Nectarian (99)]—4.57 b.y. before present Magma Ocean (Pre-Nectarian)—4.57–B4.2(?) b.y. Cratered Highlands/Very Large Basins (Pre-Nectarian)—B4.4– B4.2(?) b.y. Large Basins—(Pre-Nectarian–Upper Imbrium)—B4.2(?)–3.8 b.y. Old Large Basins/Crustal Strengthening (Pre-Nectarian)— B4.2(?)–3.92 b.y. Young Large Basins (Nectarian–Lower Imbrium)—3.92–3.80 b.y. Basaltic Maria (Pre-Nectarian–Copernican?)—B4.2(?)–1.0(?) b.y. Cryptomaria (Pre-Nectarian)—B4.2(?) –3.92 b.y. Maria (Upper Imbrium–Copernican?)—B3.9–1.0(?) b.y. Mature Surface (Pre-Nectarian–Copernican)—B3.9 b.y.–Present.

Each of these formative stages overlapped significantly. The Magma Ocean began to form before the end of lunar accretion and probably was not fully solidified until after the end of the formation of old large basins. The Cratered Highlands overlapped at least the beginning of the Large Basin Stage. The Basaltic Maria magmas probably began forming initially by pressure-release (decompression) melting and then by thermal remelting of the upper mantle. Basaltic maria lavas first appeared on the lunar surface during the Large Basin Stage as cryptomaria and then partially filled many later large basins and other depressions. The regolith that underlies mature surfaces began forming on exposed units at the beginning of the Cratered Highlands Stage and continues to form today. Graphical cartoons illustrating this formulation of lunar evolution are referred to by Plate number in the following discussion. Beginning (Stage 1). Discussion of the origin of the Moon (Plates 2 and 3) includes issues related to the origin of Earth and also to the origin of the solar system as well (100,101). One of the few undisputed scientific conclusions about the solar system as a whole is that it was formed from a concentration of interstellar dust and gas 4.567 b.y. ago. This conclusion is inferred from the radiometric ages of chondritic, eucritic, and iron meteorites (102–104) and from the initial isotopic ratios (radiometric model ages) of many lunar samples (105). Meteorites and lunar samples also preserve a record of extinct radionuclides. This record is consistent with the hypothesis that the formation of our solar system was initiated by an interstellar shock wave generated by a nearby supernova (106–108) that contributed the now extinct radionuclides and other materials to the solar nebula. The chemical similarity of carbonaceous chondrite (CI) meteorites to the composition of the Sun (109) and the current apparent abundance of such material in the solar system have led to the assumption that these meteorites closely represent the composition of primordial material that formed the Sun and the terrestrial planets. Computer models of processes in the early nebula have cast some light on what may have been happening after the shock wave during the first 10 million years or so (110–112). Once the initial angular momentum of the collapsing interstellar cloud had been dissipated and the rotating disk of the solar nebula had formed around the young Sun, particles began to stick together. This led gradually to the formation of planetesimals and then more rapidly to the

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Apollo model of lunar evolution Beginning ~4.567 plus b.y.

Current lunar diameter Magma ocean

Note: Assumes a solar system origin independent of earth. • Core •Primitive pb •Chondritic w •Nonmantle volatiles • Mantle/core •V Discontinuity •Increase in Al & Mg • Timing • Hf/w gives 0

Figure 5. Vibrating bell gyro.

displacement was proportional to the rotational rate. This is shown in Fig. 5. An example of this type of gyro is the hemispherical resonator gyroscope (HRG) discussed in Ref. 7. In this example, the resonating element that is analogous to the wine glass is a 30-mm diameter bell made of fused silica. A surrounding housing induces vibration and also senses the nodal pattern shift through the use of capacitive pick-offs. This gyro has been used on satellites and on the Jet Propulsion Laboratory’s Cassini spacecraft. The main advantage of the HRG is that there are no moving parts other than the resonator bell. A disadvantage is that the case must be evacuated and vacuum-sealed to prevent air damping. The tuning fork gyro shown in Fig. 6 is another type of vibratory gyro. In this case, the tines are excited in the plane of the tines. As the tuning fork is rotated about an axis parallel to the tines, they tend to continue oscillating in the original plane, as shown in the vector diagram in the figure. The vector component perpendicular to the plane of the tines is proportional to the rotational rate and may be measured by capacitive or optical sensing. Materials used for the tuning fork include crystalline quartz and silicon. Crystalline quartz is a highly stable piezoelectric material suitable for micromachining (8). In the case of silicon, the fork may be part of an integrated circuit chip where the controlling and sensing electronics are designed into the chip (9). Accelerometers. Accelerometers are devices used to sense changes in velocity. They are made in a number of ways, but the common feature of most is a mass that moves in accordance with Newton’s second law. This sensing mass, sometimes called the proof mass, may be suspended in a number of ways and held in the neutral position by a magnetic field. As the acceleration is sensed by the mass and it begins to move, a pickoff detects the movement and sends a restoring signal through an amplifier to the restoring coil. Rather than hold the mass in a neutral position, some designs force the mass to swing back and forth

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>0


=0

Vibrating tines

Figure 6.

d=k


655

d

Tuning fork gyro.

on a pendulum using a series of back and forth pulses. This restoring circuit also sends the restoring pulses to a counter that adds the positive and negative pulses algebraically; the sum represents the sensed acceleration. If the counter is coupled with a digital computer and integrated over time, it can keep an ongoing status of vehicle velocity. This type of accelerometer, called a pulsed integrating pendulous accelerometer (PIPA), was used successfully in the NASA Apollo Lunar Landing Program (10). Advancements in the past decade in microelectromechanical systems (MEMS), also known as microsystems technology (MST), have produced accelerometers of continually decreasing size, mass, and power usage (11). Using the same principles of vibratory gyros discussed before, sensing elements now include flexing quartz seismic beams and ‘‘squeezed film.’’ Movement may be detected by measuring changes in capacitance between the flexing mass and the adjacent fixture. Nanotechnology, generally defined as the next order of size reduction, will no doubt reduce the size of accelerometers further. Inertial Measurement Units. The stable platform, variously referred to as the inertial platform, guidance platform, or inertial measurement unit (IMU) is a common application of gyros and accelerometers. In a typical approach, a group of three single-axis gyros or two dual-axis gyros is mounted on a rigid platform, and their input axes are aligned orthogonally. Three single-axis accelerometers, or two dual-axis accelerometers, are also mounted orthogonally on the platform. The platform is then mounted on two or three gimbals, and the restoring torque signals from the gyros are used to command the gimbal drive motors. The result is that, after initial erection and alignment, the platform is maintained inertially fixed in space. A platform designed in this manner provides an inertial attitude reference and measures accelerations along the inertially fixed axes of the platform. This information can be used by a flight computer to calculate and maintain the status of attitude, acceleration, velocity, and position of the platform.

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The GN&C system then essentially flies the platform through space, and the vehicle moves around the platform in the process. The three-gyro, three-accelerometer platform was used on the NASA Apollo Lunar Landing Program (10). The two-gyro, two-accelerometer IMU is currently used on the NASA Space Shuttle (12). The Space Shuttle IMU is shown in Fig. 7. For a further discussion of IMUs, see Ref. 13. Stap-Down IMU. The limitations of early onboard flight digital computers encouraged a marriage with gimballed IMUs because of their ability to erect themselves and maintain alignment for at least short times using self-contained torquer motors and dedicated electronics. Even if the flight computer failed, the IMU alignment data could be used in these short periods of time to drive cockpit displays for manual steering. Later on, dedicated local processors within IMUs off-loaded the flight computer even more by applying scale factors, correcting biases, and encoding data words. Over the years, gimballed IMUs have earned a reputation for reliability and have become cheaper to produce. As onboard digital computers have grown in capability there has been a trend to replace the gimballed IMU by an assembly of gyros and accelerometers rigidly mounted on the spacecraft structure. This approach is called the strapdown IMU or strap-down guidance system. In this approach, the flight computer must do all of the angle resolutions (body angles to inertial angles, or body angles to Euler angles) and continually maintain the inertial reference. Further, corrections must be made by the flight computer (or a dedicated local processor) to eliminate the effects of spacecraft rotation on the accelerometers. Strap-down systems are generally smaller than gimballed IMUs, require less electrical power, and are cheaper to build. A disadvantage is that they must be continually serviced by the flight computer, and if either the strap-down system or the flight computer should fail, inertial reference is instantly lost. It is interesting to note that the Apollo Program spacecraft used a gimballed platform for the primary system and a limited form of a strap-down IMU for backup. In the latter case, body-mounted gyros were used for backup angular

Azimuth gimbal Single-axis accelerometer

Pitch gimbal

Inner roll gimbal Y axis Torquer

X axis Resolver

Outer roll gimbal

Azimuth gyro Platform Dual-axis Vertical gyro Z axis accelerometer

Figure 7. Space Shuttle inertial measurement unit (courtesy NASA).

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657

rate and displacement information for both the flight computer for automatic steering and displays for manual steering. But there were no body-mounted accelerometers, and if the gimballed IMU accelerometers failed, thrust duration had to be manually timed. It seems likely that future trends in MEMS and nanotechnology will continue to reduce the size and power usage of strap-down IMUs, making them more and more attractive for spacecraft use. However, inherent time-dependent inaccuracies in both types of IMUs require realignment using noninertial sensor or manual updating. Rendezvous and Docking Sensors. In the hypothetical mission described before, there is a phase when the spacecraft must approach and dock with another spacecraft. Assuming that the target is passive and if the spacecraft is not manned so that manual control may be employed, the maneuver must be accomplished automatically by the onboard GN&C system. Candidate sensors for the rendezvous phase would include a Doppler radar and possibly the Global Positioning System for rendezvous in Earth orbit (14,15). A laser might be suitable for docking. Both translational and rotational commands to the spacecraft attitude control rockets are required and would be generated by the guidance function of the GN&C system. The guidance algorithm must be carefully scripted so that the spacecraft is slowed down enough to prevent impact damage but still has enough kinetic energy on impact to overcome the resistance of the latching mechanism.

Space Navigation The foundation of space navigation was laid in the seventeenth century by two major advances. Early in the century, Johann Kepler, using the observations of Tycho Brahe, empirically derived his laws of planetary motion. The first law, and the most important for celestial navigation, stated that the planets of the solar system move about the Sun in elliptical orbits and the Sun is at one focus of the ellipses. Later, Sir Isaac Newton stated his laws of motion and formulated the law of universal gravitation. His work confirmed Kepler’s findings and allowed extension to celestial bodies other than planets, for example, comets, and to motions described by conics other than ellipses. These trajectories include circles (a special case of the ellipse), parabolas, and hyperbolas. A discussion of the historical background of this development can be found in Ref. 16. (See also article Earth Orbiting Satellite Theory by S. Nerem in this Encyclopedia.) Newton’s law of universal gravitation may be stated generally mathematically as F¼

Gm1 m2 ; s2

ð3Þ

where F is the magnitude of the force of attraction, m1 and m2 are the masses of the two bodies, s is the distance between them, and G is the gravitational constant whose numerical value depends on the system of units used. The force F points in the same direction as the line s that joins the two masses.

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Equation 3 may be applied with reasonable accuracy to a spacecraft orbiting Earth if certain simplifying assumptions are made: 1. Let m1 represent the mass of Earth; Earth is of uniform density and is spherically symmetrical, that is, the oblateness of Earth is ignored. This allows the Earth to be treated as a point mass at its center. 2. Let m2 represent the mass of the spacecraft, so small relative to the mass of Earth that the center of mass of the system lies at the center of Earth. 3. Let m1 be fixed in inertial space with the origin of the reference axes at its center. 4. The spacecraft is in coasting flight with only gravity acting on it, that is, other forces such as aerodynamic drag, solar winds, and electromagnetic forces are ignored. These assumptions allow simplifying the analysis to what is generally termed the ‘‘restricted two-body problem,’’ and the approach may be used for spacecraft operating near other relatively large bodies, for example, where m1 represents the Sun, Moon, or one of the planets. Following the approach of Mueller in Ref. 17, the spacecraft equations of motion can be derived as follows. Using polar coordinates and vector notation, we show the Earth–spacecraft coordinate system in Fig. 8. Referring to Newton’s second law and restating Eq. 3 in vector form in the polar coordinate system, X X

2 r ~ ¼ m2 d ~ F ; 2 dt

ð4Þ

~ ¼ ðÞ Gm1 m2 r^ ¼ ðÞ Gm1 m2 ~ r; F r2 r3

ð5Þ

where r is the distance between the masses, r^ is a unit vector that points along the line joining the two masses. Combining Eqs. 4 and 5 and simplifying produces r m~ d2~ r þ 3 ¼ 0; dt2 r

ð6Þ

where m is defined as a constant m ¼ Gm1 . Equation 6 is the vector differential equation of motion for the restricted two-body problem. Note that it is independent of m2. Using the methods described elsewhere in this Encyclopedia (see article Earth Orbiting Satellite Theory by S. Nerem) and referring to Fig. 8, two equations of importance for an orbiting spacecraft can be derived from Eq. 6: v2 m  ; 2 r

ð7Þ

H ¼ rv cos g;

ð8Þ

E¼ and

where g is defined as the flight path angle.

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SPACECRAFT GUIDANCE AND CONTROL SYSTEMS ontal l horiz Loca

z

Flight path angle



659

al

rtic

V

ve al

c

Lo

m2 Spacecraft r

Earth m1

y

x

Figure 8. Earth–spacecraft coordinate system.

E is a constant in Eq. 7, termed the specific mechanical energy of the spacecraft, implying a continual exchange of kinetic energy and potential energy throughout the orbit. H in Eq. 8 is called the specific angular momentum and is also constant throughout the orbit. Notice that as r increases, v decreases, as might be expected intuitively. Finally, referring again to the article by S. Nerem, Eq. 6 can be solved, resulting in the following scalar equation termed the trajectory equation: r¼

H 2 =m ; 1 þ B=m cos v

ð9Þ

where, as before, H is the specific angular momentum and B is the magnitude of a vector from the focus to that point on the trajectory nearest the focus. This point is called the perifocus or periapsis, or in the case of Earth orbits, the ~ and ~ perigee. The angle n (nu) is the angle between B r, the radius vector. Equation 9 is in the form of the general equation for a conic section written in polar coordinates. If p is defined as p ¼ H 2 =m and e is defined as e ¼ B=m, the equation can be written as r¼

p : 1 þ e cos n

ð10Þ

In this form, p is called the ‘‘parameter ‘‘or ‘‘semilatus rectum’’ and is a measure of the size of the conic. The term e is called the ‘‘eccentricity.’’ The solution of Eq. 10 is a circle if e ¼ 0, an ellipse if 04eo1, a parabola if e ¼ 1, and a hyperbola if e41. In summary, application of these laws to a spacecraft operating in a central force field results in trajectories that are similar to those of celestial bodies. A coasting spacecraft that has sufficient energy will orbit Earth in a plane and in

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an elliptical fashion. With the additional thrusting it can be made to travel further out to loop around the Moon and return to Earth, or be captured in an orbit about the Moon. Even more thrusting will cause it to escape Earth’s gravitational pull and proceed on a parabolic or hyperbolic path to one of the other planets or into deep space. The resulting trajectories can be divided into conical segments, whose combinations are called patched conics. As mentioned earlier, limitations in onboard expendables require planning most space trajectories carefully in advance of the actual flight to achieve the most economical and efficient missions. Launch dates, orbital plane changes, midcourse velocity changes, and rendezvous points must be determined by working backward from a desired target along the way and at the end of the flight. This is usually done by using numerical integration with considerable trial and error adjustments. Once the mission plan is determined, navigational sightings are defined in terms of times, locations, and types of sensors to be used. At each point, the sightings are made, the information is routed to the onboard flight computer, and the inertial platform is aligned. If an adjustment is required in the state vector of the vehicle, attitude alignments of the vehicle are made and midcourse delta V’s are made using the thrusters on the vehicle. This is repeated as often as required to achieve mission goals. On the launch pad, the inertial measurement unit is held in a locked position relative to the spacecraft until the last practical moment, usually a few minutes before ignition. In the minutes after IMU release to actual vehicle liftoff, the IMU is controlled by a gyrocompass program to keep it aligned relative to Earth. At liftoff, the IMU is allowed to go inertial and remain so until the next navigational update in flight. During flight, there are several types of navigational updates. Ground radar tracking is the most common for spacecraft orbiting Earth. Optical sensors may be used to take sightings of Earth, the Sun, or the stars. These sensors include star trackers, horizon scanners, or Sun seekers for automatic navigation. For the manned NASA Apollo Lunar Landing Program, the astronauts made visual sightings using a telescope and sextant that were coupled electronically to the flight computer. All of these sensors are usually carefully mounted on a rigid navigational base that also supports the inertial measurement unit, so that angular resolution from an optical sensor to stable member axes can be made precisely. Navigational fixes can also be made using the Global Positioning System (GPS) satellites [see also the article on Global Positioning System (GPS) elsewhere in this Encyclopedia]. This approach can be used to determine the vehicle attitude as well as the location in space if multiple receivers are located precisely on the spacecraft and their relative positions are differentiated (15). Star Trackers. The star tracker is an automatic optical device used to determine the angle between the spacecraft and a luminous body typically outside the solar system. Planets do not make good targets because they lie in a fairly narrow band (the zodiac) and their motion is erratic compared to stars many lightyears away. A candidate group of stars is usually preselected and stored in the flight computer along with their general location and their brightness number. The spacecraft is oriented so that the star tracker points in the general direction

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of a candidate star and then searches until a match is made. Alternately, the tracker itself may be gimballed and allowed to move relative to the spacecraft axes. Two different sightings are enough to establish the spacecraft’s position in space or to align the IMU stable member, but a third or more additional fixes are used for confirmation and greater accuracy. Star trackers now in production have angular accuracy of 0.1 arc seconds or less (18). Star trackers are sometimes used to track other spacecraft, and they may be combined with cameras for photography. Horizon Scanners. The horizon scanner is used to determine the local vertical of a spacecraft that is orbiting Earth or other planetary bodies. The local vertical may be considered a vector from the center of mass of the vehicle to the center of Earth. Three or more sightings are taken of Earth’s horizon, as shown in Fig. 9, and the angles to the local vertical relative to the spacecraft body axes or inertial axes are computed geometrically by the flight computer. Once the local vertical is determined, star sightings can be made and the latitude and longitude of the vehicle determined. Since visible light is scattered by Earth’s atmosphere, the sensor is usually designed to detect infrared waves that more sharply define Earth’s horizon. Another advantage of infrared is that Earth radiates heat at infrared wavelengths even when the horizon is not in the direct rays of the Sun, and, consequently, sightings can be made on the dark side of Earth.

Local vertical

Local horizontal

Horizon tracking point

Horizon tracking point Horizon tracking point

Earth

Figure 9. Horizon scanner.

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Sun–Earth Sensors. Sun–Earth sensors, sometimes individually referred to as Sun seekers or Earth sensors, are sometimes used as attitude determination devices for spacecraft. They are relatively simple to build, and sightings are usually reliable because the Sun and Earth are large targets and hard to miss. But for the same reasons and because of the relatively rapid movement of Earth in its orbit, they are not usually as accurate for navigation as other devices previously discussed. A principal use of Sun seekers is to orient the vehicle relative to the Sun for thermal control. Radio Navigational Aids. For Spacecraft required to return safely to Earth and land (e.g., the NASA Space Shuttle Orbiter) or land on other bodies, one or more radio navigational aids may be required. The simplest of these is the radar altimeter, which allows the spacecraft to measure altitude autonomously above the surface of the landing zone. In the Space Shuttle Orbiter, the radar altimeter measures altitude from about 5000 ft down to touchdown. Tactical Air Navigation (TACAN) units are used on the Space Shuttle Orbiter to determine the slant range and magnetic bearing of the Orbiter during landing approach. This is not an autonomous capability, and several precisely located active ground stations are required. The maximum range of TACAN is about 400 nautical miles. The Orbiter acquires bearings at a range of about 300 nautical miles and 160,000 ft altitude after entry blackout. TACAN data can be used down to an altitude of about 1500 ft. For final approach, the Orbiter uses the Microwave Scanning Beam Landing System (MSBLS). This system requires active ground stations located immediately adjacent to the runway. The Orbiter acquires the MSBLS signal at a range of 8 to 12 nautical miles and at an altitude of approximately 18,000 feet. The Orbiter Commander usually takes over manually over the runway threshold at about 100 feet altitude. Information on the NASA Space Shuttle is taken from Ref. 12.

Control A number of functions must be performed during spacecraft flight that fall under the general heading of control: *

*

*

*

*

During navigational observations, the spacecraft must be aligned relative to an inertially fixed axis system and the attitude stabilized within a very narrow angular deadband. This is sometimes called the attitude hold mode. During periods of thrusting, the spacecraft must be pointed in the correct direction, and the thrust vector controlled. This is usually called the delta V mode. Spacecraft attitude hold relative to the Sun may be required, or perhaps the vehicle is slowly rolled (so-called ‘‘barbeque’’ mode) for thermal control. Attitude control may be required in conjunction with the deployment of certain mechanisms such as solar panels, radiators, docking mechanisms, and manipulator arms. If the spacecraft is to land on Earth or on one of the planets that has an atmosphere, control during atmospheric entry may be required and may

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necessitate blending of attitude control rockets and aerosurfaces. Landing control is likely to be required and may include control of aerosurfaces, speed brakes, parasails, drag parachutes, landing gear, wheel brakes, and steering on the ground. Effector command signals must be processed (scaled, mixed, prioritized, time-delayed) and routed.

Depending on the GN&C system design, attitude change commands may come from the flight computer, directly from the flight crew, or from the ground via radio links. Attitude stabilization commands may come from body-mounted gyros that are generally less accurate than those used for inertial guidance but may serve as emergency backups for the IMU. A newer technique for stabilizing spacecraft in Earth orbit uses differential Global Positioning System (GPS)-derived position data from multiple receivers located remotely from each other on the spacecraft. It has been found that attitude knowledge of the order of 0.051 is possible (15). Body-mounted accelerometers may be used for docking sensing or aerodynamic drag sensing during entry. Effectors. Devices that produce intentional reactive forces on the spacecraft are termed effectors and may include any of the following: *

*

*

*

*

*

Attitude control thrusters that use cold gases or reactive chemicals as propellants. These may also be used for small translations for such maneuvers as docking. Major engines that produce large changes in the velocity of the spacecraft (delta V). Both thrust level and direction (thrust vector control) may be controlled. Reaction wheels where the wheel rotational rate is accelerated or decelerated to achieve reactive torquing of the spacecraft and a corresponding change in attitude. Control moment gyros that are typically mechanical gyros torqued electrically so that the resulting precession produces a desired change in spacecraft attitude. Aerodynamic surfaces and drag devices. Tethers that produce electromagnetic thrusting.

Usually the effectors themselves are not considered part of the GN&C system, but their control electronics are. Analysis of the effects on vehicle motion is usually considered a GN&C responsibility. A simplified generalized block diagram of control function is shown in Fig. 10. Environmental Disturbances. There are several disturbances that can cause a variation in attitude and possibly tumbling or wobbling. These are usually more noticeable during quiescent periods when the spacecraft is allowed to drift: *

Gravity gradient effects, usually significant when orbiting or flying near a large body.

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Flight computer Star tracker Horizon scanner Radio navigation

Navigation

Guidance IMU

Body-mounted gyros Accelerometers

Control

Figure 10.

*

*

*

Effector driver electronics

Effectors

Control block diagram.

Aerodynamic drag and moments when flying near a body having an atmosphere. Solar radiation pressure, usually significant when the spacecraft configuration includes large planar surfaces, such as solar cell panels. Electromagnetic induction when flying through the magnetic field of a large body.

The effects of these disturbances are difficult to calculate accurately but may be estimated using the methods of Refs. 19 and 20.1 At the appropriate time, when a fixed attitude is required, these must be counteracted by the attitude control system. Often a small angular deadband is desired and is maintained by use of the reaction wheels, control moment gyros, or attitude control thrusters. In the last case, thruster propellant usage is always a consideration and sometimes leads to sophisticated electronic logic for duty cycle optimization. Reference Axis System. Since effectors and some sensors are mounted rigidly to the spacecraft structure, it is customary for control analysis to use an axis system originating at the center of mass of the spacecraft. This is depicted in Fig. 11. Symbols are defined in Table 1, followed by a sign convention in Table 2. Equations of Motion. Development of the equations of motion for all mission phases of a spacecraft is an arduous task and beyond the scope of this article. The usual approach to analyzing a particular phase is to make certain simplifying assumptions to reduce the number of mathematical terms significant in that phase. Later on, after achieving a basic understanding of the dynamics, correcting terms may be added for evaluation. 1 Another approach might be that used by the U.S. Navy to estimate aerodynamic stability derivatives in the transonic region for aircraft. Displaying remarkable initiative and refreshing candor, the Bureau of Aeronautics claims success using the method of omphaloskepsis (see BuAer Report AE-614, Fundamentals of Design of Piloted Aircraft Flight Control Systems, Volume II, Dynamics of the Airframe, February 1953, p. v-11.)

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Tn

Typical thruster set Tn

CN



i

M

q

Fy

N

Thrust vector control

v

e

y

a

c

CA



 r

 G

v

y

Tn

zi

G

 

IMU

r

E

( _)

XB Fx B L p 

B

Tn

TM

xi

i

Fz

zB

Figure 11. Control reference system.

For example, shown below are six-degree-of-freedom equations of motion for the coasting (in space) attitude hold mode. Simplifying assumptions are as follows: *

*

The spacecraft is considered a rigid body. Vehicle mass and inertia remain constant during the period of interest.

Table 1. Body Axes Coordinate System xB, yB, zB xi, yi, zi Fx, Fy, Fz L, M, N j, y, c p, q, r V a b dGy dGc de da dr Tn n ¼ 1, 2, 3y TME

Orthogonal body axes, right-hand rule Orthogonal inertial axes, right-hand rule Forces along the body reference axes Moments about the xB, yB, zB axes, respectively, right-hand rule Angular displacement about the x, y, z axes, respectively Angular velocity about the xB, yB, zB axes, respectively Velocity vector Angle between the x axis and the projection of the velocity vector in the x–y plane; angle of attack Angle between the x axis and the projection of the velocity vector in the x–z plane; angle of sideslip Main engine gimbal deflection (pitch) Main engine gimbal deflection (yaw) Elevon deflection Aileron deflection (produced by differential elevon deflection) Rudder deflection Attitude control thrusters; each has a unique number and may produce forces along all body axes Main engine thrust; may produce forces along all axes and moments about all axes

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Table 2. Sign Convention Cause (þ) (þ) (þ) (þ) (þ) (þ) (þ) (þ) (þ) (þ)

Effect

angle of attack (þ a) sideslip (þ b) rudder deflection (þ dr) elevon deflection (þ de) differential elevon deflection (þ da) gimbal deflection in pitch ( þdGy) gimbal deflection in yaw ( þ dGc) force moment attitude control thruster force (þ Tn)

( þ ) CN normal force coefficient ( þ ) CA axial force coefficient ( þ ) CY side force coefficient

*

*

(  ) z airload ( þ ) y airload ( þ ) y force, (  ) yaw (  N) (  ) z force, (  ) pitch (  M) (  ) roll (  L) (  ) z force, (  ) pitch (  M) ( þ ) y force, (  ) yaw (  N) ( þ ) linear acceleration ( þ ) angular acceleration ( þ or  ) moment depending on thruster location (  ) force along the zB axis (  ) force along the xB axis ( þ ) force along the yB axis

The spacecraft is symmetrical about the x–z plane, causing the products of inertia Jyz and Jxy to drop out. Terms involving the products of inertia, pJxz and rJxz, are small and may be ignored. x€ ¼ 1=m y€ ¼ 1=m z€ ¼ 1=m

hX hX hX

f€ ¼ 1=Ixx y€ ¼ 1=Iyy € ¼ 1=Izz c

Fx THRUSTERS þ Fy THRUSTERS þ Fz THRUSTERS þ

hX

hX hX

LTHRUSTERS þ

MTHRUSTERS þ NTHRUSTERS þ

X

i Fx ENVIRONMENT ;

X

i Fy ENVIRONMENT ;

X

i Fz ENVIRONMENT ;

X

i LENVIRONMENT ;

X

i MENVIRONMENT ;

X

i NENVIRONMENT :

It is helpful to write the equations in this form to get an understanding of the importance of the different terms. Numerical solution on digital computers is usually more convenient after conversion to matrix form. For further discussion on this subject, see Refs. 13, 21–23.

The Flight Computer There is no part of GN&C technology that has evolved during the past 50 years as dramatically as the onboard digital flight computer. It started in the early

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1950s as a fire control computer 4 cubic feet in volume with 250 vacuum tubes (24). Then came the addition of the onboard navigation function with transistortransistor logic (TTL). Then its use was expanded to include guidance where steering signals were injected into an analog flight control system. By the time of the second-generation Apollo spacecraft, the flight control function was added, integrated electronic circuits were implemented, and the first automatic fly-bywire flight control was realized. With the Space Shuttle came computer control of all GN&C functions for both atmospheric and spaceflight, computer-driven cockpit displays and computer processing of all manual flight control inputs, and the beginnings of distributed processing. Current flight computers are on the order of cubic inches in volume or smaller, use relatively modest amounts of electrical power, and compute at rates incomprehensible only a few years ago. Microminiaturization is being realized in current applications, and the application of nanotechnology seems likely within a few years. Today’s flight computer performs an impressive array of functions, depending on the mission of the spacecraft: *

*

*

*

*

*

*

*

*

*

*

*

*

Guidance sensor management and data processing, guidance algorithm computation, and steering signal generation. Navigation sensor management and data processing; course evaluation, correction, and forecasting; and trajectory calculation. Flight control sensor management and data processing, vehicle attitude control, command mixing and prioritization, effector selection, thrust vector control, attitude thruster control, aerosurface control, vehicle bending and longitudinal oscillation control. Aerobraking control, parachute and parasail deployment and steering, nose wheel steering, and wheel braking. Systems management for non-GN&C systems, such as electrical power generation, distribution and control; radio/radar/television communications; exploratory payload sensor management and data processing; environmental control systems; and hydraulic and pneumatic systems. Consumables accounting and management. Flight instrumentation control, data processing, and downlink control. Generation of cockpit displays and processing of manual commands for flight control. Accommodation of ground commands via radio for software updates, GN&C commands, and data downlinking. Redundancy management. Vehicle health management (an extension of onboard checkout). Multiplex data bus management. Distributed processing management.

Input/Output. Probably the most difficult problem in implementing flight computers is communication between the computer and the various devices that are commanded or generate data. The magnitude of this problem can be appreciated intuitively if one imagines the number of wire bundles and

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connectors required to link the devices to the computer. The problem is compounded if there are redundant computers linked to one another that share the same information. Also, put simply, the digital flight computer is an anomaly in the middle of an analog world, and the necessary conversion of signals from analog-to-digital (A/D) and digital-to-analog (D/A), plus attendant voltage scaling and device scale factor and bias accommodation, is a significant task. The computer communications problem has been solved successfully in various ways: *

*

*

*

Separating the input/output function and the computation function into two boxes called the input-output processor (IOP) and the central processor unit (CPU). In this approach, the IOP handles the A/D and D/A conversions, voltage scaling, temporary data storage, data bus management (if required), and other functions. Doing the A/D and D/A conversion, scale factor adjustment, and voltage scaling at the devices served. In some cases, small local processors are implemented in the devices to perform these relatively simple chores, so that messages to and from the main flight computer are reduced to significant flight data. (This is the beginning of distributed processing.) Implementing a multiplex data bus distribution system for communicating with the various devices. In this approach, various devices that have unique addresses are connected to a data bus managed by the IOP. An important advantage of this approach is weight reduction in wiring. Sharing the computation task among several processors on the spacecraft. It could be argued that this distributed processing approach is more of a fundamental change in the approach to hardware and software than just a solution to the I/O problem. Certainly, it significantly affects the total software design, implementation, and verification/validation approach. In addition, it usually reflects not just a GN&C decision, but total spacecraft system engineering methodology.

Real-Time Operation. The requirement of uppermost importance for flight computers doing GN&C computations is the ability to keep up with the dynamics of the vehicle and the sensors that provide input data. Whereas batch processing computers in typical ground settings simply run longer when necessary to complete a job, such a delay can be deadly in flight computers in the loop of highly dynamic flight scenarios. The flight computer must maintain system stability, provide computational precision, service sensors, accept interrupts, cope with failures, and even monitor itself while running on control software cycle times typically of 40 milliseconds. The computation bandwidth is almost impossible to estimate accurately early in the design phase of a program, and even the most generous growth margins are usually exceeded and call for compromises to be accommodated later on in the flight software. A key characteristic of the computer loading is the iterative nature of many calculations, particularly in navigation. There will be errors from a number of sources, including navigation sensor measurements and system mechanization.

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A technique for handling these errors is the Kalman filter (25,26), a stochastic analysis and estimation process used extensively in space, aircraft, and missile computers since the early 1960s. Beginning with a priori knowledge of system errors, the Kalman filter continually performs a statistical error analysis and predicts new values for system variables. The resulting product is continuing improvement in position determination. Flight Software. In the early days of general-purpose digital computers, and especially in digital flight computers, computation times were extremely slow by today’s standards, and random access memory (RAM) was bulky and expensive. The typical computer programmer was a mathematician who likely had a keen appreciation of the overall computation objectives and was driven to code austere programs in what was termed machine language in those days but called assembly language today. Major disadvantages in that software coding approach were that it was labor-intensive and only the original programmer understood the logic behind his organization of the computer program. If other programmers were brought in to modify the program, there was usually some unproductive period to learn the code already in existence. The major advantage was that the assembly code was understandable, at least to the original programmer, and was easily modified. The GN&C engineer often simply had only to speak in general terms to the programmer to have a change incorporated in a timely manner. Since those early days, three things have happened to cause a revolution in the way flight software is produced today: computation rates have increased by orders of magnitude, random access memory has become inexpensive, and higher order languages (HOL) have become predominant in the development of software programs. Now, software engineers and programmers are formally educated in information systems technology rather than mathematics. The use of higher order languages such as FORTRAN, C, C þ þ , Ada, and HAL/S makes the flight software engineer and programmer very productive in terms of assembly code generated. And once software engineers or programmers learn the rules and methods for a particular HOL, they can become productive immediately without knowing very much about the ‘‘big picture’’ of the flight program. This evolution in computer hardware and software production is not without drawbacks. The very universality of a typical HOL that makes it attractive to modern diverse users also produces very inefficient assembly code. Traditionalists cringe at the squandering of RAM, and even today there never seems to be enough RAM as a spacecraft program progresses. Then, there is the assembly code itself—often virtually indecipherable and difficult to change if relatively simple modifications (called patches) are necessary. The alternative is to make the changes in the source code in the HOL and recompile, a time-consuming process and so prone to introducing undesirable effects that lengthy reverification is usually required. An unfortunate result of this modern approach is that the GN&C engineer usually does not have a thorough knowledge of the flight software program and has difficulty developing an intuitive ‘‘feel’’ for the program. He must rely almost exclusively on the formal verification/validation of the program for confidence. Another characteristic of the modern GN&C flight software program is that it seems to ‘‘grow like Topsy’’ as the spacecraft design matures. This is an

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unfortunate product of an evangelistic-like effort in the early days of spaceflight, and still today, by enthusiasts to ‘‘sell’’ the digital flight computer by welcoming more and more functions into the software program as a cheap way of implementation. Software accommodation of a feature may appear initially to be a simple and seemingly inexpensive thing to do, but often the penalties in time and expense for the subsequent verification/validation are not anticipated. A vigorous GN&C system engineering program is probably the best solution for this quandary. The cost of flight computer software is likely to be the most significant item in the GN&C budget and may exceed the cost of the rest of the GN&C system. There are several important factors involved in containing this cost: *

*

*

*

*

*

*

It is critical to document software requirements in as much detail as possible or the computer may be undersized, or the software cost driven to astronomical levels by changes later on, or both. The HOL must be chosen carefully while balancing its suitability, maturity, and compilation time with the availability of software engineers and programmers familiar with it. The executive program (operating system) must be carefully planned for real-time operation with the attendant issues discussed earlier. Application software must be compatible with the devices commanded, and the software engineer-to-device engineer interaction is expensive and sometimes difficult to orchestrate. All software must be designed recognizing the time, manpower, and equipment costs associated with the verification process. Configuration management is an absolute must for the duration of the spacecraft project. The verification and validation approach must be selected. For verification, will flight computers or computer emulators be used? For validation, will a flight computer plus simple simulator be used, or will an ‘‘iron bird’’ or highfidelity avionics integration facility plus high-fidelity simulator be used? And will performance of the verification/validation be done (or repeated) by an independent organization (so-called independent verification and validation or IV&V)?

Current Trends. Flight computers continue to grow in capability—faster computational speeds and greater RAM—and are becoming smaller, less power hungry, and cheaper. Choices that were significant a few years ago are not issues anymore. For example, fixed-point versus floating-point arithmetic is now typically decided in favor of floating-point even though floating-point takes more computing time. Word length is now typically 32 bits, and 16-bit machines are passing from the scene. Lasers are being used for communication within the computer in place of copper wire. More and more devices in the GN&C system have embedded digital processors that take care of much of the computer overhead, thus off-loading the central flight computer. Other advancements are discussed under ‘‘Integrated GN&C’’ following.

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Integrated GN&C Modern spacecraft GN&C systems have their roots in the automatic pilots (autopilots) developed for aircraft in the middle of the last century to relieve pilots on long flights. Rapid advances were made in the 1930s and 1940s, inspired largely by World War II. Initial, relatively crude, vacuum-driven gyros plus vacuum control valves plus hydraulic actuators gave way to electrical sensors plus electronic (vacuum tube) signal processing plus electrical actuators. Throughout this period, the guidance and navigation functions were typically performed by the flight crew and manual commands could be issued to the aircraft directly or via the autopilot control system. Autopilots were the means used to control bombers with steering commands coming from the famous Norden bombsight. Conceptually, then, the guidance and navigation functions were an ‘‘overlay’’ to the autopilot control system. During the 1950’s to early 1960’s when onboard digital flight computers came on the scene to perform the G&N functions this overlay approach was followed. Flight computers provided steering commands to the vehicle via the analog control system in much the way aircraft crews had done in the past. The analog control system continued to provide three-axis stability and processed the computer-generated steering commands. This design approach was typical of aircraft and guided missiles of the period, and it was followed on the NASA Gemini spacecraft and the early NASA Block I Apollo spacecraft. It is interesting to note that toward the end of this period when manned spacecraft came on the scene mechanical control linkage gave way to electrical communications between the astronaut and effectors and the expression fly-by-wire originated. In the mid-1960s, as flight computers became more powerful and compact, the entire GN&C calculation function was assigned to them, and analog channels were retained as backup. Outputs were then made directly to the driver amplifiers of the analog control system to command the attitude control thrusters and large engines for changes in translational velocity and thrust vector control during thrusting. This integrated GN&C System became the standard design approach used in guided missiles, satellites, the NASA Apollo lunar mission spacecraft, the NASA Space Shuttle, and is followed today in the International Space Station. This evolution is shown in Fig. 12. The integrated approach has several advantages. Since the flight computer performs all the GN&C calculations, the computer program (‘‘flight software’’) development can be managed more easily. Flight software design changes can be controlled and implemented with less chance of error. System internal redundancy, if desired, can be implemented more efficiently. Overall GN&C electrical power usage is less. Total system volume and mass are reduced. This is not to say that the integrated GN&C System is necessarily less complex. Likely, it is not. Computer programs become larger, harder for nondeveloper users to understand, and almost impossible to test completely. Training of flight crew and ground support personnel is usually more tedious and lengthy. Preflight testing of the integrated system becomes more involved and expensive because of the fidelity required. Nonetheless, the integrated GN&C System provides the capability of performing more complex missions and is more flexible for design changes

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1940s autopilot

Analog autopilot control Autopilot electronics

Gyros

Computer G&N

Display

Pilot

Actuators

Mech. linkage

Aerosurfaces

1950s _ 60 s autopilot

Human backup Analog autopilot control

Autopilot electronics

Gyros

GN&C sensors

Digital flight computer

Pilot

Actuators

Human backup

Driver electronics

Figure 12.

Mech. linkage

Aerosurfaces

Mid-1960s digital fly-by-wire

Effectors Wire

Evolution of the autopilot.

needed for corrections, evolutionary development, and last minute mission changes. Digital Data Buses. A major advancement in the design of the NASA Space Shuttle was the use of multiplex serial digital data buses for communication with the flight computer. This approach helped solve the computer input/ output loading problem and also provided substantial weight saving. Some 28 data buses are used in the Shuttle design. The data bus is, physically, a twisted shielded pair of wires with transformer couplers for bus protection at each electronic ‘‘user’’ box. The computer uses command and data words that have unique addresses for each receiver so that although all boxes on the bus hear all communications, each responds only to words that have its address. Multiplexers/Demultiplexers convert and format serial digital commands into separate parallel discrete, digital, and analog commands for the various systems served (demultiplexing) and do the reverse for data collected and sent back to the computer (multiplexing). The USAF continued this development, and multiplexed data buses are now common on aircraft, spacecraft, and missiles (27). An approach similar to the copper data bus discussed before is the fiberoptic bus (28). Fiber has the advantages of potentially higher data rates and less susceptibility to electromagnetic interference or intentional jamming. Disadvantages include larger minimum bend radius for installation and the potential to become unserviceable due to age or environmental effects. Fiber optics is also considered part of a general class of communications called photonics. This is a term used for the general case in which the medium is

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light rather than electron flow. In GN&C, the expression fly-by-light is sometimes used. When direct line of sight exists between devices, the use of laser beams or infrared beams for data transfer is possible. Fly-by-Wireless?. A promising development that could be used in lieu of the data bus is digital spread-spectrum radio frequency communication between the flight computer and devices in the GN&C system. Spread-spectrum is a technique used to reduce or avoid interference by taking advantage of a statistical means to send a signal between two points using different frequencies at different times. The theory is that noise tends to occur at different frequencies at different times. Therefore, even though part of a transmission might be lost due to interference, enough of the message will come through to create noticeably better output compared to fixed-frequency systems. Further, using error correction techniques, the original message can be totally restored. The promise of this approach for aircraft and spacecraft is reduction in size, weight, and power of GN&C systems while offering immunity to natural interference or jamming from man-made equipment. INVOCON, based in Conroe, Texas, has had considerable success in developing instrumentation for the NASA Space Shuttle Orbiter and the International Space Station by linking sensors to a central controller/transponder via spreadspectrum radio. Using the same principle on a project sponsored by NASA Dryden Flight Research Center, INVOCON successfully replaced a data bus link from the flight computer to an elevon actuator on a NASA F-18 research airplane (29). This concept is shown in Fig. 13. Redundancy. The use of backup hardware in aircraft and spacecraft to allow operation in the event of failure dates back many years. In manned vehicles, the crew normally has had the capability to choose the backup, usually degraded in performance, but may have been incapable of making the decision in dynamic situations. The NASA Space Shuttle, which had a design requirement to remain fully operational with one failure and to remain safe after two failures, probably reached the maximum in complexity in the approach to automatic and manual redundancy management. Basic to the design approach is the provision of four redundant primary flight computers to handle two computer hardware failures automatically and a fifth identical backup flight computer loaded with dissimilar software to be chosen in a manual switchover in case of a generic software error in the primary set. Three or more redundant channels of sensors are shared by all computers, which then command four parallel redundant channels to the effector servoelectronics. The primary flight computers automatically compare

Pilot

Flight computer

RF transceiver

RF transceiver

Aerosurface actuator

Figure 13. Fly-by-wireless concept.

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and vote by ‘‘majority rule.’’ In addition, the aerosurface actuators, which have four channels at the secondary (pilot spool) stage, were designed to vote effectively by force fight, that is, three channels overcome the offending fourth channel by force summing three against one. A comprehensive discussion of these features can be found in Ref. 30. The requirement of redundancy and automatic redundancy management in spacecraft depends on the need for vehicle survivability weighed against the cost of redundancy implementation. For expendable vehicles, it may be cheaper simply to build more spacecraft than to make them internally redundant. In manned spacecraft, loss of life is so intolerable in this country that safety must be maintained, regardless of cost.

Simulation The sheer quantity of ‘‘black boxes’’ in a modern day GN&C system begs the question as to whether system end-to-end mathematical analysis is possible. Further, the time-varying, cascading signal flow from the multiple sensors to the flight computer to the various driver electronics for the motion effectors makes end-to-end analysis and solutions by conventional closed form methods virtually unachievable. Moreover, if the solution were tractable for a particular time and set of physical conditions and produced acceptable results from an operational point of view, the flight conditions could change in the next few moments, and the analysis would have to be repeated. The analytical approach commonly used in these situations is known as computer simulation, or simply simulation, where time is the primary independent variable. Simulations are used in the conceptual design phase of a program and grow in complexity and fidelity as the program matures. Mathematical models of the functions of the different pieces of hardware evolve in complexity from simple first-order equations to higher and higher levels of fidelity. Desktop computers are adequate in the early phases to define and analyze single sensors, signal amplifiers, effectors, and units as complex as inertial platforms. As the program progresses, multiple hardware units are combined, vehicle characteristics and operational environments become better known in more detail, and the simulations are forced to move to larger complexes of computers, data library devices, and resultant data readout devices. Additional complexity accompanies increased fidelity when the simulation is required to operate in a real-time mode. Real time means that a series of events simulated is calculated on the simulator in the same period of time as the real events would occur. Real-time simulation is used any time there is flight crew involvement or flight-like hardware is substituted for simulated hardware. For GN&C system-level simulations, the flight computer is usually one of the earliest units substituted for its modeled counterpart. If the flight vehicle is a manned spacecraft, a crew station (cockpit) is usually added. Active crew displays and controls, either flight-like or simulated, are added to allow realistic crew participation for evaluating the GN&C system. If visual cues are required for this crew participation, ‘‘out-the-window’’ scenes may be included. The NASA Johnson Space Center Engineering Simulator that has

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some two million lines of computer code is an example of these large man-in-theloop simulations. As the program matures, a simulator especially designed for flight crew and ground operations personnel training is usually constructed. This simulator– trainer typically has the highest fidelity of all simulations in the program. All onboard displays and controls are included, and system simulated failures may be introduced to enhance training. For manned spacecraft, out-the-window scenes are included for all phases of the intended vehicle mission, including prelaunch, launch, orbit transfers, midcourse deltaV’s, rendezvous and docking with another spacecraft, deorbit, and landing. Vital to the fidelity of simulations throughout a program are valid mathematical models for the components of the GN&C system, for effectors and other non-GN&C equipment, and for the environment in which the vehicle is supposed to be operating. Simple first-order equations are quite satisfactory early in the design phase, but as actual hardware is built, the fidelity desired increases, and the corresponding mathematical models become more complex. Similarly, equations of motion, environmental models for ambient physical conditions, aerodynamic data, ephemerides, and out-the-window scenes all become more intricate in the drive for fidelity as the program progresses and the design matures. This often results in large simulation complexes that are expensive to create and maintain, but they are indispensable for system development, anomaly investigations, and crew training.

Integration and Verification The term integration is an overworked word in the aerospace business, but it is quite descriptive in GN&C systems development. Hardware piece parts and modules are integrated (assembled) into line replaceable units (LRUs) or ‘‘black boxes’’2 that are replaceable as the first level of maintenance on spacecraft. The LRUs are then integrated (combined) to form the whole GN&C system. A similar process is followed in developing flight computer software. As this process is followed, engineering tests are conducted, and design errors are uncovered and corrected. This successive integration is part of the development process and is the most demanding part of GN&C engineering. Verification is the process of formal evaluation of the hardware and software and is done somewhat in parallel with the integration process. Four methods are generally used and are listed here in order of preference: test, inspection, demonstration, and analysis. These are defined as follows: Test—the stimulation of the hardware and software under prescribed conditions and the responses measured and evaluated against specifications. 2

The terms black boxes and line replaceable units (LRU’s) have their heritage in the airplane world. Before the space era, most airplane electronics boxes were, and still are today, painted black. Line replaceable units are the level of replacement for maintenance on the airport flight line or ramp. A term that has come into use on the International Space Station program is orbital replacement unit (ORU), and as the name implies, the ORU is the usual replacement unit for maintenance during orbital flight around Earth.

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Inspection—the visual examination of hardware, usually in static situations. Demonstration—operation in a test-like environment but where responses usually cannot be measured and performance must be evaluated subjectively. Analysis—mathematical and logical evaluation using mathematical models, drawings, flow charts, and photographs. Formal hardware verification testing begins with the acceptance test at the GN&C system vendor’s plant. This test is conducted before the buyer accepts delivery to ensure that the equipment is functioning in accordance with the requirements of the contract between the vendor and the buyer. Ideally, one or more LRUs are randomly selected from those already accepted for a subsequent qualification test. The qualification test is conducted at extremes in ambient conditions [pressure, temperature, vibration, input electrical power, electromagnetic interference (EMI), etc.] beyond the limits expected in normal operation. This test ensures that the hardware has comfortable operating margins. The next level of testing is the integrated system test. In this test, the entire GN&C system is assembled in a flight-like configuration, and stimuli at the sensors are processed through the system to the outputs of the effector driver electronics. This is also sometimes called a system end-to-end test. Debate continues as to whether this test is an engineering development test or a formal verification test, but in reality it is usually both. Certainly, any design errors or generic manufacturing errors that are uncovered must be corrected. A critical part of this exercise is the testing of interfaces with non-GN&C equipment. The importance of this aspect is discussed in Ref. 31. The final ground verification test is the mission verification test. The significant addition to the integrated system test configuration is a real-time simulation computer complex. This is used to ‘‘close the loop’’ (i.e., close the flight

Time Development Acceptance tests Procurement specs

Qualification tests Environment specs

Integrated tests

Software IV&V System end-to-end & interface specs

Mission verification tests

Software validated Mission requirements

Prelaunch tests

Figure 14.

Flight

Integration and verification sequence.

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677

Shuttle Avionics Integration Laboratory (courtesy of NASA).

control loop using simulated vehicle dynamics) and to model missing flight hardware and the vehicle flight environment so that the entire mission, or significant portions, may be ‘‘flown’’ as realistically as practical short of actual flight. Flight computer software is used in the flight computer and is validated in the process. Successful completion of the mission verification test allows certification for flight. This sequence is shown pictorially in Fig. 14. One of the most sophisticated facilities for the integrated system test and the mission verification test is the Shuttle Avionics Integration Laboratory at the NASA Johnson Space Center. An artist’s rendering of the facility is shown in Fig. 15. The avionics are installed in a high-fidelity, three-dimensional full-scale

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arrangement with flight-like vehicle wire harnesses. The outline of the Orbiter is shown in dashed lines for orientation. For manned spacecraft, one or more flight tests are usually required before the vehicle, including the GN&C system, is certified for normal operational use.

BIBLIOGRAPHY 1. Rauscher, M. Introduction to Aeronautical Dynamics. Wiley, New York, 1953. 2. Slater, J.M., and J.S. Ausman. Inertial and optical sensors. In C.T. Leondes (ed.), Guidance and Control of Aerospace Vehicles. McGraw-Hill, New York, 1965. 3. Sagnac, G. L’ether lumineux demontre par l’effet du vent relatif d’ether dans un interferometre en rotation uniforme. Comptes Rendus d’Academie des Science Francais, 95: 708–719 (1913). 4. Commission on Engineering and Technical Systems. Technology for Small Spacecraft. National Academy Press, Washington, DC, 1994. 5. Lefeve, H. The Fiber-Optic Gyroscope. Artech House, Boston, 1993. 6. Bryan, G.H. On the beats in the vibrations of a revolving cylinder or bell. Proc. Cambridge Philos. Soc. VII: 101 (1890). 7. Wright, D., and D. Bunke. The HRG as Applied to a satellite attitude reference system. In Guidance & Control 1994, Vol. 86, Advances in the Astronautical Sciences, R.D. Culp and D.R. Rausch (eds), Proc. Annual Am. Astronaut. Soc. Rocky Mountain Guidance Control Conf., February 2–6, 1994, Keystone, CO. 8. So¨derkvist, J. Micromachined gyroscopes. Sensors and Actuators A 43: 65–71 (1994). 9. Juneau, T., W.A. Clark, A.P. Pisano, and R.T. Howe. Micromachined rate gyroscopes. In Microengineering for Aerospace Systems. The Aerospace Press, El Segundo, CA, 1999. 10. NASA Manual ND-1021043, Apollo Command Module Block II Primary Guidance, Navigation, and Control System Manual, Rev. AA, NASA Contract 9-497, 10 March 1966. 11. Connelly, J., et al. MEMS-based GN&C sensors for micro/nano satellites, Guidance & Control 2000, Vol. 104, Advances in Astronautical Sciences, R.D. Culp and E. Dukes (ed.), Proc. Annu. Am. Astronaut. Soc. Rocky Mountain Guidance Control Conf., February 2–6, 2000, Breckenridge, CO. 12. NASA Training Manual, National Space Transportation System Reference, Vol. 1, Systems and Facilities, June 1988. 13. Thompson, W.T. Introduction to Space Dynamics. Dover New York, 1986. 14. Ebinuma, T., R. Bishop, and E.G. Lightsey, Spacecraft rendezvous using GPS relative navigation. In Spaceflight Mechanics 2001, Vol. 108, Part 1, Advances in the Astronautical Sciences, Proc. AAS/AIAA Space Flight Mech. Meet., February 11–15, 2001, Santa Barbara, CA. 15. DiPrinzio, D. and R.H. Tolson. Evaluation of GPS Position and Attitude Determination for Automated Rendezvous and Docking Missions, NASA Contractor Report 4614, Langley Research Center, July 1994. 16. Baker, R.M., and M.W. Makemson. An Introduction to Astrodynamics. Academic Press, New York, 1960. 17. Mueller, D.D. Introduction to Elementary Astronautics, unpublished class notes for a U.C.L.A. short course 1966. 18. Ball Aerospace and Technologies. Aspect Camera Star Tracker. Boulder, CO. 19. Thompson, W.T. Passive attitude control of satellite vehicles. In C.T. Leondes (ed.) Guidance and Control of Aerospace Vehicles. McGraw-Hill, New York, 1965.

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20. Abzug, M.J. Active satellite attitude control. In C.T. Leondes (ed.) Guidance and Control of Aerospace Vehicles. McGraw-Hill, New York, 1965. 21. Kaplan, M.H. Modern Spacecraft Dynamics and Control. Wiley, New York, 1976. 22. Noton, M. Spacecraft Navigation and Guidance. Springer-Verlag, London, 1998. (Note: This monograph has free implementation software written in C þ þ available on the Internet.) 23. Wie, B. Space Vehicle Dynamics and Control, AIAA Education Series, Reston, VA, 1998. 24. Smith, G.H. Overview of aerospace vehicle computer applications. In C.T. Leondes (ed.), Computers in the Guidance and Control of Aerospace Vehicles, AGARDograph No. 158, February 1972. 25. Kalman, R.E. A new approach to linear filtering and prediction problems. J. Basic Eng. March 1960. 26. Kalman, R.E., and R.S. Bucy. New results in linear filtering and prediction problems. J. Basic Eng. March 1961. 27. MIL-STD-1553B, Digital Time Division Command/Response Multiplex Data Bus, 1996 (replaces Rev. A dated 30 Apr 1975). 28. MIL-STD-1773, Fiber Optics Mechanization of an Aircraft Internal Time Division Command/Response Multiplex Data Bus, 02 October 1989. 29. NASA Tech Briefs, Vol. 24, No.2, Associated Business, New York, 2000, p. 8b. 30. Hanaway, J.F., and R.W. Moorehead. Space Shuttle Avionics System. NASA SP-504, 1989. 31. Euler, E.A., et al. The failures of the Mars Climate Orbiter and Mars Polar Lander: A perspective from the people involved. In Guidance & Control 2001, Vol. 107, Advances in Astronautical Sciences, R.D. Culp (ed.), Proc. Ann. Am. Astronaut. Soc. Rocky Mountain Guidance Control Conf., January 31—February 4, 2001, Breckenridge, CO.

JON H. BROWN Fort Worth, Texas

SPACELAB Spacelab is the European contribution to the Post Apollo Program. U.S. President Nixon made an offer in 1969 to the eight member states of the European Space Agency to participate in the planned recoverable launcher area in developing the Space Shuttle and its payloads. Spacelab is a reusable, multipurpose, modular laboratory and in the first 10 years of Space Shuttle operation, became its major payload. This payload remained attached to the Orbiter during the whole flight and maintained certain dependences. Spacelab was not a habitat, so that the Mission and Payload Specialists stayed in the Orbiter during take-off and landing and used it also as living quarters. Spacelab was designed for a 10-year lifetime and/or 50 missions. Inside the laboratory, called the module, a shirtsleeve atmosphere for the astronauts was required; experiments on the outside had to be conducted in the open space on a pallet. Both module and pallet were accommodated inside the Shuttle payload bay and remained attached to the Shuttle during the mission (Fig. 1). This fact limited the time of a Spacelab mission to the time of a Shuttle flight.

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Figure 1. Spacelab. This figure is available in full color at http://www.mrw.interscience. wiley.com/esst.

On 4 June 1974, the contract to design and build Spacelab was given by the European Space Agency (ESA) to an industrial consortium of companies from eight European countries under the leadership of the German space company ERNO Raumfahrttechnik in Bremen. The contract, won in a competition, was based on a modular design of the laboratory (Fig. 2) that applied state-of-the-art technology and experience in aircraft design to solutions of the turnaround requirements of frequent reuses. For most of the time from 1974 to 1983, the Spacelab development ran parallel to the Shuttle development on a no exchange of funds basis to use common hardware and subsystems as much as possible.

The Modular Concept The multipurpose objective of Spacelab led to a modular design to meet the needs of missions of very different natures for the module as well as for the pallet. To keep the cost down, the module and pallet elements were made a standard size, each 2.70 m. This allowed using standardized handling and transportation equipment and several combinations of module and pallet elements for the different missions. For pallet only modes, a special Igloo was

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Major spacelab flight elements Cover plate

Modular means Many different flight combinations Less design/manufacturing cost Less handling/operational complexity Fewer spares

AFT end cone Pallet Igloo (pallet-only missions)

Experiment segment Forward end cone Experiment racks Core segment

EVA airlock

Utility connections (to orbiter) Tunnel Tunnel adapter (to orbiter) AEG _ AERITALIA _ BTM _ DORNIER _ FOKKER/VFW _ HSD _ KAMPSAX_ MATRA _ SABCA _ SENER

Figure 2. The modular concept Major Spacelab Flight elements.

developed to house the equipment that controlled the functions of the pallet and its payloads.

The Design Concept The general shape and dimensions for this Shuttle payload were determined by the Shuttle payload bay. The Spacelab module has a cylindrical shape that is the diameter of the Shuttle payload bay. Although not needed in space, the cylinder contained a floor on which racks are mounted along the sidewalls. The required 50 reuses and therefore, the corresponding 50 refurbishments on Earth demanded a floor/ceiling configuration to allow proper handling when under gravitational forces. This fact led to the solution to mount the racks on floor elements which could be rolled out of the cylinder like freight containers out of an airplane (Fig. 3). By this concept, the payload and experiments could be removed from the laboratory on their floor elements and shipped directly to their origin, and the new payload could be rolled in without any delay. By using this solution, the originally required 2-week turnaround time on the ground between missions could be achieved. An aircraft structure design was selected for the pallet that allowed plain standardized surfaces for payload mounting and easy handling. A special instrument pointing system (IPS) was designed and built to be mounted on the

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Cargo airplane roll-on/roll-off feature AFT cone

Removable insulation bands Interface feedthrough MGSE alignment stand

Roll-out floor and standard rack assembly

11 July 1974

B _ 13

AEG _ AERITALIA _ BTM _ CIR _ DORNIER _ FOKKER/VFW _ HSD _ INTA _ KAMPSAX _ TERMA _ MATRA _ SABCA _ SEL_ SENER _ TCSF

Figure 3. Payload integration—system level.

pallet to allow high accuracy pointing for payloads, thus eliminating errors from the Shuttle attitude control system. An airlock on top of the module provided access of experiments to open space for direct exposure, and a window provided a view to the outside. Mission Specialists and Payload Specialists entered the laboratory through a tunnel from the lower aft flight deck of the Orbiter into Spacelab (Fig. 4). The 1.1-m diameter tunnel was not part of the European delivery program.

The Module Structure The basic structural element of the Spacelab module was a cylindrical segment 2.70 m long and 4.06 m in diameter. This cylindrical shell was made of AL- 2219 alloy sheets that had a chemically milled inside waffle pattern and were welded at the seams. End flanges of forged aluminum alloy stabilized the cylinder and were used to join the segments to each other and to the end cones. Two of these segments, two end cones, internal floors, and several racks were part of the initial module structure’s flight hardware deliveries. Cutouts in the shell for the window and the airlock required local strengthening to avoid distortion problems and to ensure pressure-tight mating surfaces (Fig. 5). The floor could take loads up to 300 kg/m at the center and up to 500 kg/m at the outsides, and the racks could hold a load of 580 kg/m per side. The module

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Jog

Crew airlock

Spacelab module

Orbiter cabin

Adapter

Tunnel

Figure 4. Offset concept for the transfer tunnel showing the jog required to permit centerline entrance to the Spacelab module.

and the pallet were supported in the Shuttle payload bay by trunnions and keel fittings which had to carry the weight of the loaded module on the ground and all loads imposed during launch and landing. A special surface finish had to be developed for these attachments to meet the very low friction coefficient to allow for twisting and bending of the Shuttle payload bay imposed by changing pressure, thermal loads, and landing impact. The module segment was attached by three fittings to the Orbiter payload bay, two trunnions and one keel fitting. The subsystem equipment which was needed to run the laboratory was to be placed in two double racks at the forward end of the module and on a subfloor beneath the main floor. All other single or double racks were available for the payload. There were up to seven overhead storage containers in each segment attached to overhead supports mounted on the module ceiling, which also served as attachments for the racks.

The Pallet Structure Like the module segments, the pallet segments were identical modular units 2.90 m long that could be mounted separately or together in pallet trains of two or three units or in combination with one or two module segments in the Shuttle payload bay. The basic U-shaped design allowed mounting experiments of up to 3000 kg on the inside or about 1100 kg/m equally distributed on standardized

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Drawing by T. Hall Flight International, London

Key Module 1 Insulation blanket 2 Close-outs (skin/racks) 3 Cabin air ducting (from subfloor) 4 RAAB 5 High data-rate recorder 6 Handrails 7 Water/freon heat exchanger 8 Utility tray 9 Gaseous nitrogen supply line 10 Gaseous nitrogen tank 11 Temperature transducer 12 Forward end-core

13 Module/Orbiter lower feed-through plates (two) 14 Insulation blanket supports 15 Freon pump 16 Water pump 17 Lithium hydroxide cartridge stowage 18 Freon lines 19 Control-centre rack 20 Debris traps 21 Workbench rack 22 Stowage container (lower for access) 23 Upper module/orbiter feed-through plate 24 Gaseous nitrogen fill-valve bracket 25 Gaseous nitrogen reducing valves (two-stage) 26 Position for double rack 53 Cold plates 54 Inner skin-panels 55 Outer skin-panels 56 Pallet/Orbiter primary pickup 57 Pallet/Orbiter stabilizer pickup 58 Connector support bracket 59 Pallet herd-points 60 Handrails 61 Support systems remote aquisition unit (RAU) 62 Experiment RAU (several) 63 Experiment power distribution box 64 Pallet/bridge supports 65 Experiment-supporting bridge 66 Electrical junction box 67 Integrally-machined aluminium-alloy ribs

27 Position for single rack 28 Keel-fitting 29 Subfloor 30 Aluminium alloy module shell 31 Electrical connectors for rack 32 Floor of aluminium-skinned honeycomb sandwitch (centre panel fixed, outer panels hinge up for access) 33 Overhead duct channels 34 Viewport 35 Nasa high-quality window 36 Fasteners for insulation blanket 37 Rack fire-suppression system 38 Double rack 39 Experiment airlock 40 Airlock controls

41 Overhead lights 42 Avionics cooling-air ducts 43 Aft end-cone 44 Radial support structure 45 Fire extinguisher (Halon) 46 Portable oxygen equipment 47 Foot restraint 48 Module/Orbiter pickups (four) 49 Module-segments joints, incorporating seals Pallets 50 Freon lines from module 51 Pallet interface 52 Cable ducts

Experiments 68 Synthetic aperature radar 69 Solar spectrum 70 X-ray astromony 71 Solar constant 72 Charged-particle beam 73 Advanced biostack 74 Isotopic stack 75 Microorganisms 76 Lyman Alpha 77 Waves 78 Low energy electron-flux

Figure 5. Spacelab, Europe’s space laboratory.

panels and on hard points that had an available volume of 33 m3 per pallet. Very thin aluminum faceplates were used on the panels to save weight. Four symmetrical support fittings were applied to suspend the pallet in the Orbiter payload bay (Fig. 6). For the pallet-only missions, a special equipment container, the so-called Igloo, was developed. The primary structure of the Igloo was a cylinder-shaped

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Figure 6. Pallet structure.

locally stiffened shell made of aluminum alloy forged rings. The lower end was closed, and the upper end was a mounting flange. The Igloo was attached directly to the forward end frame of the foremost pallet. The secondary internal structure was mounted on the primary structure and was finally closed by a cylindrical cover which was also a shell of aluminum alloy closed at the top and mated to the top flange of the primary structure. For the connection to the Orbiter, feedthrough of utility lines was provided by penetrations of the Igloo. The thermal control of the Igloo was both active and passive; most equipment was cold-plated to the FreonTM system. The internal atmosphere was air, and a drying agent was included to avoid condensation after closing the cover. Nitrogen was added via a fill valve to ensure sufficient internal pressure during the mission. Overpressure safety protection was provided redundantly by a relief valve and a burst disk.

The Instrument Pointing System One of the most interesting and ambitious subsystems was the instrument pointing system (IPS) for telescopes and sensors. The IPS provided three-axis attitude control and stabilization of payloads up to 2000 kg mounted on the pallet and exposed directly to open space. The pointing accuracy was 71 arcsecond for a payload that has a diameter up to two meters and a length up to 4 meters (Fig. 7). The design solution was an end-mounted approach in which the three gimbal systems would be mounted on the pallet providing support to a circular mounting frame to which the optical instrument would be attached. In zero gravity operation, the gimbal support structure would handle only the momentum of the instrument masses. During launch and landing, a payload gimbal separation mechanism would separate the IPS and the payload, a gimbal latch mechanism

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Optical Thermal sensor shield package

Roll drive unit Bumper device Payload attachment flange Payload clamp units

Elevation drive unit Power electronic unit

Replaceable PCA struts Gimbal support structure

Gimbal latch Harness Jettison Cross mechanism separator separation plane elevation drive unit (a)

(b)

Figure 7.

(a) Instrument pointing subsystem (b) Spacelab experiments testing.

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would lock the gimbal system, and a payload clamp assembly would provide a three-point clamping of the payload.

The Environmental Control Subsystem The environmental control subsystem (ECS) of Spacelab had to ensure the requirements of sea-level pressure and a shirtsleeve atmosphere for the crew and the experiments in the laboratory. It was agreed among the partners that the Orbiter would supply the oxygen and the Spacelab would provide its own nitrogen as well as the components to regulate and distribute the air supply throughout the module and remove excess moisture and carbon dioxide. A high-pressure nitrogen tank was mounted on the outside of the module. The CO2 created by the breathing of the crew was absorbed by LiOH cartridges mounted on the subfloor underneath the main floor of the Spacelab cabin as well as by the two air fans and the contamination filters. The heat generated by the module and the pallets was transferred to the Orbiter thermal control system. The minimum cabin temperature was supposed to be 19.51C within a range of 18–271C and the humidity between 30 and 70%. One circuit of air served the interior of the module cabin at a flow rate of 5–12 m/ min, and a second circuit served the cooling of the equipment mounted in the racks. By passing through a heat exchanger, the heat load of the air was transferred to a flow of water that was pumped through an Orbiter payload heat exchanger to transfer the total Spacelab heat load finally to the externally mounted Orbiter radiators. The total capacity of this system was 5.8 kW. A further component of the thermal control system was a FreonTM cooling circuit serving experiments mounted to cold plates on the pallet floor or lower sides. A total of eight such cold plates could be mounted on the pallet to accommodate experiments. Four thermal capacitors allowed storing of peak heat loads. As coolant, FreonTM 114 was used because water would freeze immediately when exposed to the outside environment of the pallets.

The Command and Data Management Subsystem The most complicated subsystem of Spacelab is the command and data management subsystem (CDMS). It is an extension of the Orbiter telecommunication system which finally transmits data generated by the laboratory and its experiments which have been acquired by the CDMS and multiplexed in low-rate housekeeping and high-rate scientific data streams. NASA used two real-time data transfer systems. The Space Tracking and Data Network (STDN) transferred relatively low data rates of up to 192,000 bits/s from the Orbiter to the ground (downlink) and 72,000 bits/s from the ground to the Orbiter (uplink). Higher data rates were transmitted with the help of the Tracking and Data Relay Satellite System (TDRSS). This link allowed downlink data rates up to 50 million bits/s. On the ground, the data were received by ground stations in White Sands, New Mexico, via S-band or Ku-band through the geostationary satellites of the TDRSS. From there, they were relayed to the Payload Operations Control Center

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at the NASA center in Clearlake near Houston and to the Spacelab Data Processing Facility at the Goddard Space Flight Center in Greenbelt, Maryland. The CDMS had to provide mass memory storage for the central computers and for the periods of time when the Shuttle was out of sight of the TDRSS storage of high-rate digital data. A high-rate multiplexer ensured the handling of up to 16 channels of 16-megabit-per-second data. Onboard computers, remote acquisition units that interfaced the functional laboratory equipment and the experimental equipment to the data bus, voice digitizer and intercom equipment, data display/keyboards, input/output and interconnecting stations, and a high-rate digital recorder were further elements of the CDMS. In principle, the CDMS was divided into two elements: the data processing assembly and the high-rate data assembly. In the data processing assembly, two remote acquisition units separately served the Spacelab subsystems and the experiments. Each of these parts had a computer, an input/ouput unit, and a 1-megabit-per-second digital data bus which was routed through the laboratory. Three keyboards and displays and the mass memory served the two parts. This assembly was to acquire data; to distribute timing and commands via the data bus; and to interface with the Orbiter multiplexer/demultiplexer, the pulse code modulation unit, and the master timing unit. The high-rate data assembly consisted of the high-rate multiplexer, the high-rate recorder, and the demultiplexer and high-rate data recorders on the ground. This assembly acquired data directly from experiments and time division multiplexed these data into a composite data stream of up to 48 megabits per second, which could be transmitted to the ground by the Ku-band communication system. Besides the high-rate data, some lowspeed data, as well as digitized voice and timing data, could be added into this composite data stream.

The Electrical Power Distribution Subsystem (EPDS) Spacelab received its power from the Shuttle Orbiter. The power was produced by three fuel cells that converted oxygen and hydrogen into electrical power by a chemical reaction. In orbit, the power of one of these three fuel cells could be dedicated to Spacelab. This meant providing 7 kW of dc at 28 volts. The maximum power for Spacelab could be raised to 12 kW for 15 minutes once within a 3-hour time frame. Inverters (400 Hz) in the laboratory could convert parts of the energy into three-phase ac power at 115 and 200 volts. Built-in control and regulation circuits provided protection for inverters and consumers against overvoltage and overcurrent. During takeoff and landing, the available power for Spacelab was reduced to 1 kW because most of the experiments and subsystems were dormant in these mission phases. The total available quantity of oxygen and hydrogen in the Shuttle Orbiter was the factor limiting the total energy available for a mission. This limited a normal mission to about 300 kWh available for experiments in the Spacelab module configuration and 550 kWh in a Spacelab pallet-only configuration. Additional fuel cells in the Shuttle Orbiters ‘‘Columbia’’ and ‘‘Endeavour’’ allowed extending the mission’s duration to 2 weeks as far as the power was concerned.

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Basically, of the 7 kW power available for Spacelab, 5 kW was required by the laboratory subsystems and mission-dependent equipment, leaving only 2 kW for the experiments; 2 kW in the case of the pallet-only mode, only 5 kW remained for experiments. The main power conditioning, distribution, and control of the power through a feeder from the Orbiter took place in a power control box. This box included a shunt regulator to limit the main bus voltage to 32 volts and melting fuses to prevent short circuits on the feeders. A subsystem power distribution box distributed the dc and ac power into dedicated subsystem feeders.

The Follow-On Production (FOP) The NASA/ESRO MOU of 1973 provided for NASA, to procure at least one further Spacelab no later than 2 years before the delivery of the first, provided that it met the agreed specifications and schedules and was reasonable in price. This agreement was realized by a procurement contract signed at the end of January 1980 and a preceding letter contract for the procurement of the essential longlead items.

The Mechanical Ground Support Equipment (MGSE) Parallel to the development of the Spacelab module and pallet and their auxiliary subsystems, sophisticated ground support equipment had to be developed that mirrored the dimensions and accuracy requirements of the flight hardware. Besides the integration process in its different phases, ground handling included transport on the ground and in the air. The modular design of Spacelab required a mechanical ground support equipment that allowed mating two modules and their two end cones on the ground to simulate the real flight configuration with respect to the accurate integration and verification of the O-ring seals. It also required mating a threepallet train. In addition to the integration equipment, the modules needed special transport equipment in which each module could be placed like a flat can, because its dimensions did not allow transport on the road in a normal upside position. Also, the eight double racks and four single racks, which were to hold avionics equipment and the experiments, had to be supported for integration, handling, and transport, including their floor sections onto which they were mounted. Besides the mechanical support, access equipment was needed such as workstands, scaffolding, access kits, floor covers, soft covers, desiccants, and transport tie-downs. For simulation and test on the ground, the gaseous nitrogen system, as well as the fluid loop of the active thermal control system and the environmental control life support system, had to be provided. Finally, checkout equipment for weight and balance, leak checks, and verification of the environmental control life support elements had to be provided.

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The Electrical Ground Support Equipment (EGSE) The parallel development of the Shuttle and Spacelab caused interface problems for the laboratory design because open issues or changes in the Shuttle development had a direct impact on the dependent payload, Spacelab. In particular, this problem existed for the electrical ground support equipment, which was to simulate the Shuttle on the ground during Spacelab integration, test, verification, and qualification. The electrical ground support equipment served for test and ground checkout during integration of the laboratory and during the integration of the payload into Spacelab. These tests and checkouts were to be conducted in a manual and in an automatic mode. Serving this purpose was Automatic Test Equipment (ATE), which was composed of a three-station console with keyboards and CTR displays, a CII MITRA 125 computer with related peripherals, recording and timing equipment, measurement components, stimulus generators, and interface equipment. Further elements of the electrical ground support equipment were the ground power unit to simulate the Orbiter power supply, an Orbiter interface adapter to simulate the Orbiter power and signal interface, an experiment segment/pallet simulator to simulate power loads for the Spacelab electrical power system and signals for the command and data management system, and an experiment subsystem simulator to simulate experiment power loads and data interfaces. Two complete sets of this electrical ground support system were developed and used first in Bremen for Spacelab development and later in the Kennedy Space Center (KSC) O and C Building to support mission preparation and payload integration.

The Flight Hardware Although not contained in the original flight hardware delivery program, ESA and NASA agreed to use the Spacelab engineering model pallet structures for orbital flight tests (OFT). These pallets were kept very simple, equipped only with an Orbiter FreonTM pump, cold plates, a power control box, and a flexible module for command purposes. With this additional agreement about the European hardware, the first Spacelab OFT-pallet arrived at NASA’s Kennedy Space Center in Florida on 4 December 1978. The pallet, packed in a special container, was shipped by sea in mid-November from Bremen to the Cape via Savannah, Georgia, and the Interstate Coastal Waterway. The second OFT-pallet arrived at KSC on 22 April 1979. They became part of the preoperational Orbiter payload, intended to collect data prior to the first operational flight of Spacelab. On 21 December 1981 the two segments of the long module were flown out of the Hanover airport because the runway was appropriately long for the USAF C-5 ‘‘Galaxy’’ nonstop flight directly to the Kennedy Space Center in Florida. The distance from Bremen to Hanover was covered on the road. The Igloo and three pallets followed on 28 July 1982. The follow-on production (FOP) flight hardware was delivered to KSC on 27 July 1984 (Fig. 8).

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Figure 8. (a) Spacelab Integration Hall at ERNO. Bremen Assembly of the Spacelab Flight Unit II, comprising Pallets plus Igloo. The Igloo (secondary structure in tilted position and without cover) is attached to the pallets, which are partly equipped with cold plates and subsystem equipment. (b) Spacelab hardware content/deliveries/configurations and use in pre- and operational Space Shuttle flights.

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Schedule and Cost From fall 1969 to summer 1972, the ESA and the European space industry investigated their possible contribution to the U.S. Post Apollo Program proposed by President Nixon. The direct participation in the Orbiter development and the development of a Space Tug by the Europeans was not accepted, and in 1973, ESA and NASA signed the agreement for Spacelab. After competitive study phases A and B, the development proposals were submitted to ESA by two industrial teams on 14 April 1974. The final contract was awarded to the ERNO team of Bremen on 5 June 1974, which was the Authorization To Proceed (ATP). With a great number of industrial and governmental partners from both sides of the Atlantic involved, the following milestones were achieved: 24 June 1974 Kick off 11 November 1974 Preliminary requirements review (PRR) 19 June 1975 Systems requirements review (SRR) 29 September 1975 Contract signature ESA/ERNO in Paris 02 July 1976 Preliminary design review (PDR) 03 March 1978 Critical design review (CDR) 04 December 1978 Delivery of the first OFT flight pallet to KSC 28 November 1980 Spacelab engineering model rollout 30 November 1981 Spacelab flight module (FU-1) roll out 11 December 1981 Spacelab flight module delivered to KSC 05 February 1982 Spacelab (FU-1) acceptance at KSC 08 July 1982 Spacelab flight pallets (FU 2) delivered 13 January 1983 Spacelab design certification review at NASA HQ 18 November 1983 Final flight readiness review at NASA HQ 28 November 1983 First Spacelab flight with STS 9 ‘‘Columbia’’ 27 July 1984 Follow-on production unit delivery More than 9 years after development began, Spacelab conducted its first flight. Although 3 years longer than originally assumed, this duration did not delay the Shuttle schedule. The launch date of STS-9 was well met by the completion of the development of the laboratory and by delivery of the first Spacelab payload (FSLP). The participation of nine European countries in the Spacelab program meant that the development work was executed by companies that had nine different currencies. ESA used, therefore, a neutral artificial currency for this project; the accounting unit (AU) was a kind of provisional EURO, which corresponded to the U.S. dollar with a slightly changing rate of exchange over time. The baseline cost of 193 MAU (million accounting units) in 1974 increased by changes in the scope of the contract, cost overruns, and a 103% cost escalation during the 9 years to 579 MAU. Not including the escalation, the overall program cost was below 140% of the agreed cost for this first European manned spaceflight program. By the end of the program, Germany had contributed 56.79%,

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Italy 11.34%, France 10.42%, the United Kingdom 6.67%, Belgium 5.29%, Spain 3.56%, the Netherlands 2.50%, Denmark 2.07%, Switzerland 1.07%, and Austria 0.29%.

The Missions Spacelab was designed for a 10-year operational lifetime and/or 50 reflights. The laboratory exceeded its design lifetime by more than 4 years, but it did not achieve 50 flights. The first pallet flight occurred in November 1981, and the last in July 2001, a time span of almost 20 years. The first module flight took place in November 1983, and the last in April 1998, a time span of almost 15 years. In total, 15 Shuttle flights carried a Spacelab module, 6 flights carried pallets plus an Igloo, and 10 flights carried pallets only (Tables 1 and 2). AN OFT pallet was used in the Orbiter ‘‘Columbia’’ to accommodate the Shuttle imaging radar SIR-A as early as the second flight. The third Shuttle flight STS-3 also contained a pallet in the payload bay. The first flight mission of Spacelab SL-01 was the Shuttle flight STS-9, and the Orbiter ‘‘Columbia’’ was launched on 28 November 1983 from Kennedy Space Center Pad B at 332:16:00:00 GMT. The Spacelab consisted of a long module and a pallet whose total mass was 15,265 kg (33,252 lb), both up and down. The flight was scheduled for 9 days. However, NASA and ESA decided at midtime to prolong it by 1 day because the usage of the resources on board allowed such an extension. So the mission was completed when the Orbiter ‘‘Columbia’’ landed at Edwards Airforce Base in California on 8 December 1983 at 343:23:47:00, accomplishing a mission duration of 11 days, 7 hours, 47 minutes. In that time, the Orbiter rounded the globe 166 times and covered a distance of about 7.5 million kilometers at an inclination of 571 about the equator (Figs. 9 and 10). The Spacelab mass was 12,780 kg (12,115 lb), of which 8145 kg (18,135 lb) were for the module and 3386 kg (7449 lb) for the pallet. The mass of the experiments and associated equipment was 3982 kg (8802 lb) and for the verification flight instrumentation 856 kg (2123 lb). Spacelab 1 was a joint mission of ESA and NASA. Each agency sponsored about half of the scientific payload. In the United States, the Marshall Space Flight Center was assigned responsibility for the NASA sponsored portion of the payload. In Europe, the responsibility for the ESA sponsored portion of the payload, referred to as the first spacelab payload (FSLP), was entrusted to a technical management team known as Spacelab Payload Integration and Coordination in Europe (SPICE). This team was set up by ESA in 1976 at the German Aerospace Research Establishment (DFVLR), Cologne-Porz. Spacelab 1 was a multidisciplinary mission of more than 70 experiments in five areas of scientific research: astronomy and solar physics, space plasma physics, atmospheric physics and Earth observations, life sciences, and material science. In total, 38 different facilities for these experiments existed. Of the 38 experiments, 16 required to conduct investigations were situated on the pallet and 20 in the module. Two of the 38 experiments had components both on the pallet and in the module.

694

Launch date & duration

28 Nov. 83 10 days 29 Apr. 85 7 days 29 Jul. 85 8 days 30 Oct. 85

7 days 2 Dec. 90 9 days 5 Jun. 91 9 days 22 Jan. 92 8 days 24 Mar. 92 9 days 25 Jun. 92 14 days 12 Sep. 92 8 days 8 Apr. 93 9 days 26 Apr. 93

STS carrier

STS 9 Columbia STS 51B Challenger STS 51F Challenger STS 61A

Challenger STS 35 Columbia STS 40 Columbia STS 42 Discovery STS 45 Atlantis STS 50 Columbia STS 47 Endeavour STS 56 Discovery STS 55

330 km 281 350 km 391 300 km 571 300 km 571 300 km 281 300 km 571 300 km 571 300 km 281

571 250 km 571 360 km 501 320 km 571

Orbit incl. & alt.

Table 1. Spacelab Module Missions

Mat. Science Solar Astron. Mat. Science

LM þ MPESS IG þ 3P þ IPS LM þ MPESS

SL-02 SL-D1

Mat. Science Life Science Atm. Physics Multidiscipl.

LM LM IG þ 2P LM þ EDO LM 1G þ 1P LM þ USS

SLS-01 IML-01 ATLAS-1 USML-01 SL-J ATLAS-2 SL-D2

Mat. Science Life Science Atm. Physics Solar Astron. Mat. Science

Life Science

IG þ 2P þ IPS

ASTRO-1

Life Science Astronomy

Multidiscipl.

Discipline

LM þ IP

Configuration

SL-01 FSLP SL-03

Mission

X

X

X

Major

X

X

X

X

Partial

European user participation

M. Schlegel U. Walter

D. Frimout

U. Merbold

R. Furrer E. Messerschmid W. Ockels

U. Merbold

European astronant

695

10 days 18 Oct. 93 14 days 8 Jul. 94 15 days 3 Nov. 94 11 days 2 Mar. 95 17 days 27 Jun. 95 10 days 20 Oct. 95 16 days 20 Jun. 96 17 days 4 Apr. 97 4 days 1 Jul. 97 16 days 2 Apr. 98

16 days

Columbia STS 58 Columbia STS 65 Columbia STS 66 Atlantis STS 67 Endeavour STS 71 Atlantis STS 73 Columbia STS 78 Columbia STS 83 Colombia STS 94 Colombia STS 90

Columbia

300 km

300 km 391 280 km 281 300 km 571 300 km 281 350 km 521 300 km 391 300 km 391 280 km 281 300 km 281 300 km 281 LM þ EDO LM þ EDO

LMS MSL-01

Life Science

LM þ EDO

USML-02

LM þ EDO

Life Science Mat. Science Mat. Science

LM

SL-M

NEUROLAB

Mat. Science

IG þ 2P þ EDO

ASTRO-2

Mat. Science

Astronomy

IG þ IP

ATLAS-3

LM þ EDO

Life Science Mat. Science Atm. Physics

LM þ EDO

IML-02

MSL-01R

Life Science

LM þ EDO

SLS-02

X

X

X

X

X

X

X

Favier

Clervoy

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Table 2. Spacelab Pallet Missions STS carrier

Launch data and duration

STS-2 Columbia STS-3 Columbia STS-41G

12 Nov. 81 2 days 22 Mar. 82 8 days 5 Oct. 84

Challenger STS-51A

8 days 8 Nov. 84

Discovery STS-39 Discovery STS-46 Atlantis STS-59 Endeavour STS-64 Discovery STS-68 Endeavour STS-75 Columbia

8 days 28 Apr. 91 8 days 31 Jul. 92 8 days 9 Apr. 94 11 days 9 Sept. 94 11 days 30 Sept. 94 11 days 22 Feb. 96 16 days

Spacelab pallet and purpose 1 Pallet/OSTA-01-Shuttle Imaging Radar SIR-A 1 Pallet/OSS-01-NASA Office of Space Science 1 Pallet/OSTA-03-Photographic and radar images of Earth 2 Pallets/Retrieval of Palapa and Westar communication satellites 1 Pallet/AFP-675-Air Force Program 675 1 Pallet/TSS-01-Tethered satellite 1 Pallet/SRL-01-Space Radar Laboratory 1 Pallet/LITE-Lidar In-Space Technology 1 Pallet/SRL-02-Space Radar Laboratory 1 Pallet/TSS-1R-Tethered satellite reflight

Some of the experiments on the pallet and in the module operated automatically; others were operated from the ground or by the crew remotely through the computer or by controls located on the instrument front panels. Other experiments in the module were operated directly by the crew. Of the 38 experiments, 13 were sponsored by NASA, and 25 by ESA. Of the 13 NASA experiments, 5 were located on the pallet and 8 in the module. Of the 25 European experiments, 11 were situated on the pallet and 14 in the module. The 38 experiments on board Spacelab 1 were selected from more than 400 proposals solicited by NASA and ESA in 1976. The STS-9 Shuttle/Spacelab crew consisted of six astronauts, five Americans and with Ulf Merbold, one German. The Commander was the experienced 52-year-old John Young on his sixth space flight. The crew was divided into two teams, a red team and a blue team, both to work in space in 12-hour duty cycle shifts: John W. Young Commander (CDR) Red Team Brewster H. Shaw Pilot (PLT) Blue Team Dr. Owen K. Garriott Mission Specialist (MS 1) Red Team Dr. Robert A. R. Parker Mission Specialist (MS 2) Blue Team Dr. Byron Lichtenberg Payload Specialist (PS 1) Red Team Dr. Ulf Merbold Payload Specialist (PS 2) Blue Team

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Rack 12 stowage Rack 8 MSDR Rack 4 primary metric camera

Pallet

Rack 11 prim. control for pallet exp. Rack 9 very wide field camera

Figure 9. SL-1 configuration.

After the first Spacelab module flight SL-01, a second Spacelab mission, SL-02, with a pallet-only mode, including the Igloo and an instrument pointing system (IPS), was launched on 29 July 1985 and flown for 8 days. This mission specialized in solar astronomy. Both missions were joint NASA/ESA missions and were considered the conclusion of the development program. After those missions, several other missions used Spacelab, both U.S. missions and international missions. With D-1 and D-2, Germany equipped two

Figure 10.

Side view.

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multidisciplinary Spacelab missions. Spacelab D-1 was launched on 30 October 1985 and was very successful. Spacelab D-2 was planned after the tragic ‘‘Challenger’’ mission in early 1986 and was delayed accordingly for more than 4 years until 26 April 1993. On 17 April 1998, the last Spacelab mission with a module took off into space with the Orbiter ‘‘Columbia’’ for a 16-day flight on a life science mission called NEUROLAB. This was 14 years and 5 months after its maiden flight in November 1983.

Lessons Learned The Spacelab program was NASA’s largest international cooperation program, because it involved with ESA on the European side, nine countries and its major payload was developed, parallel in time to the Space Shuttle, a unique situation with respect to interdependences and interfaces. And for this situation, the partners had agreed to a ‘‘no exchange of funds’’ funding and use of common hardware as much as possible. High hopes were attached to the Shuttle as the first recoverable launch system in the Post Apollo Program and its major payload Spacelab. For the European partner, however, Spacelab was the first manned space project ever. Under these circumstances, the program was a full success. On the European side, the selection of an industrial prime contractor for the first time for a project of this size was a solution, which proved to be very helpful, because most of the coordination effort between companies from nine European countries was executed within the consortium under the leadership of the prime contractor. The employment of U.S. industrial consultants in those areas that lacked experience in Europe, especially in the aspects of manned space flight, proved to be a very supportive measure. Representatives on all levels in the governmental organizations and in industry helped to avoid misunderstandings and served as on the spot mediators and real-time informants. A very tight network of clearly defined periodic meetings, working groups, and committees on all levels in government and industry kept all parties involved, provided the information to the need-to-know, and challenged the responsibilities in the decision-making process. The intentional application of state-of-the-art technology for the Spacelab development reduced the risk and was an important factor in keeping the schedule and cost plan that had so many partners of different experience involved. The frequent usage of the laboratory for very different missions with short turnaround times required well-known technologies and easy access to spare parts. The access to Electric/Electronic Equipment (EEE) parts during a period of more than 9 years of development and 15 years of operation was, however, a problem. During this period of more than 24 years, the fast advance of technology in the EEE parts industry caused problems of an undisturbed supply of parts for Spacelab. The availability of only one prototype of Spacelab, EM-1, in the program for cost saving reasons, resulted in the fact that there was no prototype in Europe to follow the operational program after EM-1 was delivered to the Marshall Space Flight Center in Huntsville, Alabama, on 28 November 1980.

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After the infrastructural system of Shuttle/Spacelab was completed, the partners, NASA and ESA, and the participating countries in Europe did not continue their joint effort for using of Spacelab as the public would have expected. There was no agreed upon organization for evaluating the results achieved in the laboratory and their importance for science, technology, and application. Attention was directed too soon to the next generation of a manned space facility, the Space Station. The limited mission time of the Shuttle-attached Spacelab was one of the reasons for this development, although it took more than 10 years before the decision for the next generation of manned orbital infrastructure was made. Fortunately, much of the good Spacelab experience could be saved and flow into the even greater challenge of a permanent, manned station in low Earth orbit launched with the Space Shuttle. Again, a module will be supplied by Europe incorporating Spacelab experience and spirit.

BIBLIOGRAPHY 1. Lord, D.R. Spacelab. An International Success Story. NASA Scientific and Technical Information Division, Washington, DC, 1987. 2. Messerschmid, E., R. Bertrand, and F. Pohlemann. Raumstationen, Systeme und Nutzung. Springer-Verlag, Berlin, Heidelberg, New York, 1997. 3. Buedeler, W., and S. Karamanolis. Spacelab, Europas Labor im Weltraum. Wilhelm Goldmann Verlag, Muenchen, 1976. 4. Von ERNO bis Asrtrium, Wir zeigen die Vergangenheit und sprechen von der Zukunft Schriftenreihe der Raumfahrthistorischen Archivs Bremen e. V. Heft 1. Stedinger Verlag, Lemwerder, 2001. 5. Sebesta, L. Spacelab in Context. European Space Agency, Paris, France, 1997. 6. Longdon, N. Spacelab Data Book. European Space Agency, Paris, France, 1983.

HANS E.W. HOFFMANN ORBCOMM LLC Dulles, Virginia

SPUTNIK 1: THE FIRST ARTIFICIAL EARTH SATELLITE Humanity’s leap into space was one of the greatest scientific achievements of the twentieth century. However, one of the ironies of history is that this great scientific and engineering achievement was largely facilitated by the Cold War between the two superpowers, the United States and the Soviet Union. In the late 1950s, the United States held a significant advantage over the Soviet Union in number of nuclear warheads and capability to eliminate the most important strategic facilities on its enemy’s territory. Although the Soviet Union already had nuclear weapons, it did not have second-strike capability because it

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did not have any delivery vehicles capable of delivering warheads to U.S. territory. The senior political leadership of the Soviet Union assigned its missile science and industry a task of ‘‘special governmental importance’’–development of a ballistic missile that could render the United States vulnerable. The early experimental intercontinental ballistic missiles developed by Soviet scientists and engineers were used to place the first artificial Earth satellites into orbit. Thus, the history behind the development and launching of the first satellite into orbit is primarily the story of a missile that became a launch vehicle for a satellite rather than a missile carrying a weapon of mass destruction. As long ago as February 1953, a government decree assigned NII-88, the lead institute for missile technology, the task of performing ‘‘theoretical and experimental research related to developing a two-stage ballistic missile and studying the future applications for a missile with a range of 7000–8000 km.’’ This task was led by Sergei Korolev, who at that time was the Chief Designer of OKB-1 (at that time part of NII-88). Korolev, for his part, was head of the Chief Designer’s Council, which included the following Chief Designers: Valentin Glushko for rocket motors, Nikolai Pilyugin for trajectory control systems, Vladimir Barmin for ground launch systems, Mikhail Ryazanskii for radio systems, and Viktor Kuznetsov for gyro control devices. The preliminary design for the missile called for developing a 170-metric-ton, two-stage missile with a 3000-kg detachable nose cone. The nose cone was designed to hold a nuclear warhead weighing up to 1000 kg and having an 80-kiloton maximum yield. However, in 1953, the Soviet Union successfully tested a thermonuclear warhead—the ‘‘hydrogen bomb,’’ with a yield a few dozen times that of a nuclear warhead. In late 1953, Korolev was ordered to change the design and increase the throw weight to 3000 kg. This required an increase in the total payload mass to 5500 kg. Major revisions in the missile design were required to maintain the desired range specifications. The basic changes implied by the increased throw weight requirement were as follows: 1. change launch configuration that used the upper load-bearing members of the side-mounted engines as supporting points for the four launch trusses and removed as the missile begins to climb; 2. switch to a high-thrust, four-chamber propulsion system; 3. use movable steering engines in place of exhaust vanes; 4. increase fuel capacity; 5. increase of 38 metric tons in the thrust of each propulsion system; 6. addition of backup gyro control devices; 7. addition of systems to monitor drainage of tanks and synchronization of fuel consumption by the side-mounted engines; 8. development of a fundamentally new launch system that reduced the load on the missile structure, thereby substantially reducing the weight of the missile.

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The Government decree ordering development of the R-7 two-stage ballistic missile was approved on 20 May 1954. A subsequent decree provided a schedule and list of deliverables. Flight testing was scheduled for February 1957. Development and fabrication of all systems and the first few missiles took place during 1955–1956. The missile design represented a fundamental improvement in structure, frame configuration, dimensions and mass, propulsion-system thrust, control actuators, a new dynamic launch system configuration, and new inertial and radio navigation techniques. The missile itself consisted of four side-mounted, first-stage engine assemblies mounted around the central engine assembly of the second stage. The internal design of the side-mounted engine assemblies and central engine assembly were similar to the single-stage designs then in use. The engines used kerosene and liquid oxygen as propellants. All five engine assemblies operated from the ground up and were started almost simultaneously. The side-mounted engines were turned off upon stage separation, and the second stage remained active. Each engine assembly was based on a standard four-chamber engine with sea-level thrust greater than or equal to 80 metric tons. This package arrangement required a synchronized tank drainage system and an engine-thrust control system. This missile marked the first use of special steering chambers that moved by electrohydraulic actuators in response to commands issued by the control system. The development process for this missile included extensive experimental testing of individual systems, in addition to experimental testing of the overall design. Final testing of the control system was performed on special M5RD test missiles based on the R-5 missile. Flight testing of 10 M5RD missiles using the new control-system equipment enabled verification of the apparent-speed-control system, the tank drainage system, and a new telemetry system, including the sensor instrumentation (in particular, a vibrational sensor system). The R-5R test missiles were used for flight testing telemetry-based missile velocity measurements, using a centimeter-wavelength pulsed radio system, radio-wave attenuation in engine jets, and correctness of the underlying design principles for the radio direction finder. Flight testing three R-5R missiles provided a large amount of experimental material that enabled substantial modifications to the instrumentation fabricated for initial flight/design testing of the R-7 missile. Almost all of the sensor equipment for the onboard instrument system was developed from scratch. In all, seven sets of recording sensors were installed on board the missile. In addition, special ground-based recording equipment was used to monitor the missile launch and the behavior of launch systems. More than 600 in-flight parameters were measured using a total of 2800 kg of onboard instrumentation. Unlike all previously developed single-stage missiles, the R-7 missile and launch fixtures formed a single dynamic system. The launch ground system included more than 30 individual systems and assemblies. The interfaces between ground systems and missile and the interfaces between individual ground systems were tested using special missile mockups interfaced to the ground systems at the launch facility. However, before this, appropriate system testing of the launch system and missile was performed at

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the Leningrad Machinery Plant. The missile-lift procedure was simulated using high-capacity lifting cranes and a missile filled with water in place of fuel. It was essential to ensure that the gantries retracted simultaneously and that the missile frame would mechanically mate with the launch system. The complete ground system at the launch site was tested using a simulated R-7 missile, which was repeatedly placed on the launch stand and repeatedly fueled. The Scientific Research Institute for Firing Tests (NII-229) built several high-capacity rocket test stands for testing the propulsion systems in combination with the missile frame. From August 1956 to March 1957, NII-229 performed five firing tests of the side-mounted engine units, three tests of the central engine unit, and two tests of the complete package of five engine units. The first R-7 missile for flight testing arrived at the Tyuratam support facility (Scientific Test Site NIP-5, the future Baikonur) in March 1957. The horizontal testing performed at the NIP-5 support facility included electrical and pneumatic testing of each engine unit, postshipment verification of engine-unit alignment, assembly of the engine units into a single package, and integration testing of all electrical and pneumatic systems (‘‘horizontal system testing’’). The first meeting of the State Flight Test Committee occurred on 10 April 1957. The Committee was chaired by M.V. Ryabikov, Chairman of the Military Industrial Complex, and had the following members: Chief Marshal of Artillery M.N. Nedelin (Deputy Chairman); Chief Designer S.P. Korolev (Engineering Manager); Chief Designers V.P. Glushko, N.A. Pilyugin, V.P. Barmin, M.S. Ryazanskii, V.I. Kuznetsov, S.M. Vladimirskii (Deputy Chairman, State Committee for Radio Electronics), A.I. Nesterenko (Manager, Scientific Test Site 5), G.N. Pashkov (State Planning Committee [Gosplan]), I.T. Peresypkin (USSR Ministry of Communications), and G.R. Udarov (Deputy Chairman, State Committee for Military Equipment). The missile was first launched on 15 May 1957. The flight appeared to proceed normally for the first 60 seconds, at which time a fire broke out in the tail compartment. Reduction of the telemetry data revealed that one of the sidemounted engines fell off 98 seconds into the flight, when the missile became unstable. The root cause of the accident turned out to be a leak in a fuel line. Nevertheless, this launch did confirm that the control-system parameters for the first-stage segment were correct and gave us confidence in the launch dynamics. The second scheduled launch attempt on 11 June 1957 was unsuccessful because the disc on the main oxygen valve for side-mounted unit C froze and an error had been made during installation of the nitrogen blowdown valve on the oxidant line for the central engine unit. The missile was returned to the Support Facility. The third launch was on 12 July 1957. Thirty-three seconds into the flight, the missile became unstable. The root cause of the accident turned out to be a short circuit between the control-signal circuits and the housing in the angularvelocity integrator for the roll channel. The fourth launch on 21 August 1957 was successful, and the missile, for the first time, hit a target on Kamchatka Peninsula. A TASS statement announcing the launch of a long-range, multistage intercontinental ballistic missile was carried by the USSR mass media on 27 August. The fifth launch of the R-7 missile on 7 September 1957 confirmed the results of the previous launch. However, although the missile components and systems operated normally during the active portion of the flight, the warhead

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reentry vehicle (RV) broke up upon reentry into the lower atmosphere. A significant amount of time was required for research and development work related to, and fabrication of, new warhead RVs capable of withstanding the high temperatures and large gas-dynamic loads that occurred during the reentry portion of the flight. With the consent of the Government and the State Commission, it was decided to use two of the original twelve R-7 missiles built for development testing as launch vehicles for the first artificial Earth satellites. These launches provided a practical opportunity to accumulate additional experimental data on all missile systems except for the forward compartment containing the nuclear warhead. Thus, these early satellite launches were an integral part and outgrowth of development testing for early intercontinental ballistic missiles. A fortunate convergence of historical fates in the early 1950s led to the renewal of creative contacts between Sergei Korolev and Mikhail Tikhonravov, his former colleague in development of early perwar unguided missiles during the 1930s. In the late 1940s, Tikhonravov led a group of enthusiasts from the highly classified Scientific Research Institute 4 [NII-4] who were studying designs for space launch vehicles, as well as a variety of issues related to subsequent manned spaceflight. Tikhonravov and his group were the first professional people in the Ministry of Defense system who dared to say that the R-7 intercontinental ballistic missile developed under Korolev’s direction could, with slight modifications, be used as a satellite launch vehicle. With slight changes to the flight plan, the missile could place a satellite weighing up to 1500 kg into orbit instead of delivering a 5.5-metric-ton nuclear warhead at a range of 8000 km. The Ministry of Defense generals did not support Tikhonravov’s initiative, but Korolev quickly grasped the possibilities for practical implementation of his long-held dream of human spaceflight, and was able to arrange for Tikhonravov’s transfer from NII-4 to OKB-1 (Korolev’s design bureau). In May 1954, Korolev had presented a proposal to the Minister of Armaments (Dmitrii Ustinov), the Council of Ministers, and the USSR Academy of Sciences to develop the world’s first artificial Earth satellite and launch it with the R-7 missile for research purposes. The basic idea contained in Korolev and Tikhonravov’s report was that ‘‘the artificial Earth satellite is an inevitable phase in the development of space hardware, following which interplanetary missions will become possible.’’ In August 1954, the USSR Council of Ministers approved a proposal to study the scientific and engineering issues involved in a spaceflight. Korolev had support from the following senior government officials, as well as the Academy of Sciences: Deputy Chairman of the Council of Ministers V.M. Malyshev, D.F. Ustinov, and Ministers B.L. Vannikov, M.V. Khrunichev, and K.N. Rudnev. Despite the military’s fears that development of a spacecraft would be a distraction from the primary tasks involved in developing the R-7 missile, on 30 January 1956, the Council of Ministers approved the development of an unstabilized spacecraft (‘‘Object D’’) weighing 1100–1400 kg and carrying 200–300 kg of scientific research instrumentation. The Government gave the USSR Academy of Sciences responsibility for developing the scientific research equipment. The required task orders were issued as directives to the Ministry of Defense Technology (the lead ministry) and all other ministries and organizations involved in missile development and production.

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OKB-1 and its subcontractor organizations completed their design work in 1956 and then moved on to fabricate Object D, which was to become the first artificial Earth satellite. Object D was to be used for scientific research in the following areas: density and ionic composition of the high-altitude atmosphere, solar particles, magnetic fields, cosmic rays, temperature conditions within the spacecraft itself, braking of the spacecraft in the upper atmosphere, and the accuracy of spacecraft position and orbit determination. The scientific instrumentation and spacecraft onboard systems were to have been powered by solar panels and storage batteries, and the spacecraft would have had an automatic temperature control system. This spacecraft was also to have been the first recipient of an onboard control system with a special programmable timer. In-flight control of the scientific research program was to have been performed via a radio command link from the ground. A special radio telemetry system was to have transmitted the research results to the ground. An extensive network of ground stations, forming a unified command and telemetry system controlled from a single center at NII-4, was to have been constructed. By late 1956, it was clear that the schedules for fabricating the scientific instrumentation had slipped, and it became uncertain whether the spacecraft could even be launched during the U.N.-sponsored International Geophysical Year. Because press reports indicated that the United States was also preparing to launch a spacecraft early in the International Geophysical Year, Korolev proposed postponing the launch of Object D and launching a very simple spacecraft, the PS, carrying no scientific instrumentation instead. On 15 February 1957, the Government accepted the proposal by Korolev and the Academy of Sciences to launch an extremely simple unstabilized Earth satellite (Object PS) to verify that the PS could be observed in orbit and that signals transmitted by the PS could be received, as well as to ensure worldwide priority in the space race. The Government would permit the spacecraft launch to occur only after one or two successful launches of the R-7 missile. The extremely simple satellite (the PS) was designed, fabricated, and prepared for launch using the missile in only 8 months. The launch was scheduled for 6 October. However, intelligence reports from overseas indicated that the Americans were preparing their own satellite for launch in early October, so Korolev sped up the preparations, and the PS was launched on 4 October. The first artificial Earth satellite was launched into space on the fifth launch of the first intercontinental missile. Two of the preceding four launches had been unsuccessful because the problems encountered by the developers of the first intercontinental missile were new. The goal of this launch was to place an extremely simple satellite into orbit and also to obtain additional experimental data on the dynamics of the missile launch and propulsion systems, the control and guidance systems, the stage separation system, the onboard sensors, the operation of equipment on the ground, and the command and telemetry system on the ground. During the round-the-clock effort to prepare the rocket for launch, several problems were identified and eliminated right on the launch pad. In response to one problem report during rocket fueling, one of the fuel tanks for the side-mounted engines was drained and refilled to test the ‘‘tank full’’ alarm system.

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The guidance system was adjusted to place the missile into an orbit with the following parameters: perigee: 223 km; apogee: 1450 km; period: 101.5 min. These orbital parameters could be achieved using half of the guaranteed fuel supply, provided that the guidance systems and all engine systems operated properly. Trajectory measurements indicated that the rocket performed normally during the active portion of the flight. However, the second-stage engine ran out of fuel before it was scheduled to be turned off by the guidance system. The world’s first artificial Earth satellite was launched into space on 4 October 1957 at 22:28 Moscow time. It ended up in an orbit with the following parameters: perigee; 228 Km: apogee; 947 Km: inclination; 65.11: and period: 96.17 min. Reduction of the telemetry data enabled collecting a large amount of experimental data concerning operation of the missile, various individual missile systems, and the ground launch systems. The spacecraft weighed 83.6 kg; the body of the spacecraft consisted of a sphere 580 mm in diameter with four collapsible-whip antennas (2.4 m and 2.9 m long). The body of the first artificial Earth satellite consisted of two aluminum alloy hemispheres filled with dry nitrogen (pressure 0.13 MPa); the joint between the two hemispheres was sealed using a rectangular-cross-section, vacuum-grade O-ring. The pressurized enclosure held an electrochemical power source and two radio transmitters that continuously transmitted on 20.005 and 40.002 MHz (wavelength 17 and 7.5 m, respectively). The transmitter signals consisted of alternating telegraph ‘‘marks’’ and ‘‘spaces,’’ each 0.3 s long. The ‘‘mark’’ on each frequency was transmitted simultaneously with the ‘‘space’’ on the other frequency. The radio transmitter system had a total mass of 3.5 kg, and each transmitter had an output of about 1 W. The telemetry data (temperature and interior pressure) were transmitted to the ground by modulating the frequency of the ‘‘mark’’ and ‘‘space’’ signals. Each transmitter had two collapsible-whip antennas (approximately 701 apart). Each pair of antennas had a nearly spherical antenna pattern. The temperature control system had a radiator with a fan-driven, sealedloop, forced-gas heat exchange system designed to maintain a stable interior temperature in the face of variable external thermal fluxes. The temperature control system used a bimetallic thermal relay as the sensor element. Whenever the temperature increased above 361C, the fan came on, and nitrogen circulated through the system to transfer heat away from the hemisphere that was acting as a radiating surface (emission coefficient 0.35–0.4, solar absorption coefficient 0.23–0.4). The fan was turned off whenever the temperature fell below 201C. The intended purpose of the automated onboard electrical system was to turn on the electrical power to the instruments, once the spacecraft reached orbit (i.e., upon separation from the launch vehicle). During the launch phase, the spacecraft was placed under a nose cone for protection against aerodynamic and thermal effects; the nose cone was jettisoned when the second-stage engine shut down. The spacecraft carried three silver-zinc batteries weighing 51 kg. The batteries could support operation of the instruments for 3 weeks. The first

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satellite lasted 92 days, and completed B1400 orbits around Earth. On 4 January 1958, it reentered Earth’s atmosphere and burned up. The orbital parameters of the first Soviet artificial Earth satellite were such that it was visible from all continents across a wide range of latitudes. Observations of the spacecraft’s motion, reduction of the observations, and prediction of the future motion of the spacecraft based on these results served as an early practical exercise in using ground-based spacecraft control systems and equipment for measuring spacecraft parameters. The first satellite was observed using radio equipment, as well as optical instruments at astronomical observatories. The news that the first artificial Earth satellite had been launched aroused extremely broad interest among radio amateurs and amateur astronomers around the world. In the Soviet Union, 66 optical observing stations and 26 clubs with extensive collections of radio observing gear regularly observed the spacecraft. In addition, thousands of radio amateurs attempted radio observations of the spacecraft. State scientific radio monitoring stations also observed with radar and radio range finder systems. This was the first rich set of statistical data concerning the transmission of meter-wavelength radio waves through the ionosphere and also represented the first opportunity to receive radio signals at two different frequencies from regions of the ionosphere that had heretofore been inaccessible, that is, above the ionization peak or even above the entire ionosphere. Extremely valuable data were also obtained on radio-wave absorption in previously unstudied layers of the ionosphere, as well as new data on the structure of these regions and the ion concentration at various altitudes and times of day. These systematic measurements showed that the altitude of the main peak in the ionosphere and the peak electron concentration vary from day to night, north to south, and east to west. Radio propagation measurements at the frequencies emitted by the spacecraft at various altitudes provided a new avenue for ionospheric research. These observations led to the discovery that the decrease in electron concentration in the upper ionosphere (above the main peak) with altitude is five to six times slower than the increase with altitude below the peak. For example, when the observations were made (in October), the electron concentration increased by approximately a factor of 10 from 100–300 km altitude, whereas it decreased only by a factor of 2 from 300–500 km in altitude. The initial indications that micrometeoroids were not hazardous to spacecraft was an extremely important result. The radio methods used included radio ranging and Doppler observations of the radio signal from the spacecraft. These early experiments indicated that the Doppler effect could be used successfully to determine spacecraft orbital parameters. It became obvious that the accuracy of orbit determination could be quite high if the transmitter frequency were increased and if automated frequency measurement equipment were used. Early, high-sensitivity photographic techniques were developed for observing this spacecraft. Image-converter tubes turned out to be especially promising in this regard. The news that the Soviet Union had launched the world’s first artificial Earth satellite turned out to be quite an unexpected sensation for the entire human community. The flight of the first satellite around Earth caused a stunning resonance around the world. Virtually the entire world press carried front-page banner headlines reporting the news, thereby underlining that Soviet

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science had taken the lead. The American government was shocked, along with American scientists confident of their superiority. The senior leadership of the Soviet Union was surprised by this enthusiastic reaction from their people and people around the world. Therefore, to consolidate the political success that had been so unexpectedly achieved in the Cold War, the General Secretary of the CPSU Central Committee, Nikita Khrushchev, proposed that Korolev launch a new satellite in honor of the 43rd anniversary of the October Revolution. The first, extremely simple satellite was still operating in orbit, and there was no sense in launching another, similar satellite. A second spacecraft was readied in less than a month and launched on 3 November 1957. The dog Laika, who later became famous, was the first experimental animal to orbit Earth. The second spacecraft did not separate from the second stage of the rocket, and, for the first time, data on the behavior of a dog in space was transmitted to the ground via a multichannel telemetry link. Object D, which was to have been the first spacecraft, was not launched until 15 May 1958. It was the third Soviet spacecraft and the first that could truly be called a space laboratory based on the amount of scientific instrumentation on board. The successful launch of the first artificial Earth satellite marked the beginning of humanity’s journey into space. Many extremely urgent scientific problems required direct experiments at altitudes of hundreds or thousands of kilometers above Earth’s surface. Although the significance of artificial Earth satellites had long been understood, launching them had remained an insoluble problem. The main difficulty had been developing a rocket that could give a spacecraft a velocity of the order of 8000 m/s. Only after the Soviet Union had developed an intercontinental ballistic missile did it become possible for the first time in history to launch an artificial Earth satellite. The superior design of this missile enabled placing a spacecraft with the required weight of scientific instrumentation into orbit. Further improvements of the R-7 missile—in particular, the addition of a third and then a fourth stage led to manned space programs, communications satellites, and the first automated interplanetary spacecraft for studying the Moon, Venus, and Mars. Modernized versions of the R-7 missile are currently used for manned and cargo spacecraft to support the International Space Station. BORIS CHERTOK ENERGIA Space Association Russia

SUN THE SUN AS A STAR Although many cultures deified the Sun, for example, Ra, the Egyptian sun god,  RYA, the sun god in Ancient India, humans became aware of its real and SU

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significance to terrestrial beings only in the last 400 years. Before the Copernican heliocentric model of the solar system was accepted, the Sun and the ‘‘wandering stars’’ (which we now know are planets), the Moon and the fixed stars were objects whose apparent daily motions needed to be explained. The Ptolemaic system hypothesized intricate moving equants and epicycles to describe the motion. By placing the Sun logically at the center of the solar system and observing moving spots on the Sun, Galileo established its place in our current world view. Contemporary humans have learned that the Sun supplies a nearly constant flux of visible radiant energy to Earth and is also magnetically connected to it so that powerful releases of energy at a distance of 1 AU (1.5  108 km or B93 million miles) from Earth can produce surprisingly large local effects, such as the extensive power failure in October 1989. Basic Properties of the Sun. The Sun is only one of several billion stars in our galaxy, but as a special example, it is important to understand how it compares to other stars. Following the development of stellar spectroscopy, techniques became available to classify stars by their locations on the Hertzsprung– Russell (HR) diagram, which is a plot of stellar luminosity as a function of spectral type. Figure 1 from Chaisson and McMillan (1) shows an HR diagram for several varieties of stars. The ordinate is the object’s optical luminosity or the total energy emitted per second by the object, and the abscissa is the star’s spectral type, which is a measure of its surface temperature. To place a star on the HR diagram, the star’s distance must be known, so its absolute luminosity can be determined from the measured luminous flux using the inverse square law. As in Fig. 1, the luminosity is usually expressed in terms of the Sun’s luminosity. The star’s spectral type is determined by measuring the star’s ‘‘color’’ and characterizing its spectrum. The sun is a ‘‘G’’ class star. Stars that have higher surface temperatures, for example, are ‘‘O’’ class, and ‘‘M’’ class stars have lower surface temperatures. All normal stars span a temperature range from B40,000 K to B4000 K. Actually a star’s spectral class is determined by comparing its spectrum with the spectra of a set of standard stars. Some physical properties of the Sun and the way the properties are determined are shown in Table 1 (from Reference 2). Of most interest here is the group of stars on the main sequence, the socalled ‘‘normal stars,’’ that are in the early stages of their evolution. (The Sun is indicated by the symbol .) Formally, the Sun’s spectral class is G 2 V, and its surface temperature is B6000 K. It is considered an average main sequence star in all respects: luminosity, surface temperature, and size. The M class stars are typically 0.1 times smaller than the Sun, and the O class stars are 100 times larger than the sun. However, due to the sun’s proximity to us, detailed studies of it and the radiations it emits can be done with great precision, and this knowledge can be used to understand other stars better. From direct measurement of isotope ratios, we know that the Sun’s age is B4.5 billion years, and from the theory of stellar evolution, that it may live at least another 5 billion years. When its current internal energy source is depleted, it will expand and become a red giant and eventually a dwarf star similar to those on the HR diagram in Fig. 1. As far as its origin is concerned, the Sun and planets may have formed as a result of a supernova explosion near the early solar nebula. This aspect of the Sun’s early life and its future are fascinating topics

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Figure 1. Most stars have properties within the shaded region known as the main sequence. The points plotted here are for stars lying within about 5 pc of the Sun. The diagonal lines correspond to constant stellar radius, so that stellar size can be represented on the same diagram as luminosity and temperature. (Recall that  stands for the Sun.) From Reference 1.

(see, for example, Reference 3), but our goal here is to describe the present Sun in some detail. The current understanding of the Sun draws on advances in both groundbased observing techniques and sophisticated space instrumentation, as well as in theoretical computer models using the observations as input. This review first briefly describes some important ground-based observations and identifies several space missions whose main purpose has been to advance our knowledge of our nearest star. Next, the structure of the Sun and its atmosphere is described in more detail, beginning at the center of the Sun and emphasizing our current understanding in appropriate detail. A major goal of this work is to give an up-todate summary of our knowledge of solar activity. The new science of helioseismology is discussed later, and space weather, as it affects Earth, is also covered briefly later. For each topic, a brief history will be given with a summary of current knowledge followed by identification of important unsolved problems. References for further study will be given for each topic discussed. A general reference for the topics discussed can be found in chapter 16 of Astronomy Today by Chaisson and McMillan (1).

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Table 1. Properties of the Suna Datum Mean distance

How found Radar reflection from planets

1.471  108 km

Minimum distance to Earth Mass

Acceleration of Earth

Mean angular diameter

Direct measure

Diameter of photosphere

Angular size and distance

Mean density

Mass/volume

Gravitational acceleration at photosphere (surface gravity)

GM/R2

Luminosity

Spectral class Effective temperature

1 AU 149,597,892 km 1.521  108 km

Maximum distance to Earth

Solar constant

Value

Measure with instrument such as bolometer Solar constant times area of spherical surface 1 AU in radius Spectrum

333,400 Earth masses 1.99  1033 g 310 5900 .3 109.3 times Earth diameter 1.39  1011 cm 1.41 g/cm3 27.9 times Earth surface gravity 27,300 cm/s2 1.9 cal/min/cm2 21.368  106 ergs/s/cm2 3.8  1033 ergs/s

G2V

Derived from luminosity and radius of Sun

5800 K

Received visible flux Apparent magnitude and distance

 26.7 þ 4.8

Rotational period at equator

Sunspots and Doppler shift in limb spectra

24d16h

Inclination of equator to ecliptic

Motions of sunspots

71100 .5

Visual magnitude Apparent Absolute

a

From Reference 2.

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Ground-Based Observations of the Sun.

Heracleides (388–315 B.C.) and Hipparchus (190–120 B.C.)1 can be considered the first who observed the Sun with a scientific motivation. The Ptolemaic Earth-centered system, as promoted by Aristotle, held sway until Copernicus (1475–1543) proposed the heliocentric model for the solar system, even though Aristarchus (B320B250 B.C.) had earlier considered such a model. However, serious study of the Sun began after Lippersheim invented the telescope in 1608. Galileo (1610) in Padua and Fabricius (1610) in Wittenberg used it in their independent discoveries of sunspots. Solar physics investigations for the next 192 years consisted of recording the number of sunspots on the solar disc until 1802 when Wollaston discovered dark lines in the solar spectrum. Using an improved spectrograph, Fraunhofer discovered 547 additional dark lines, features that were named after him. The origin of these strange features on the Sun, as explained in 1859 by Kirchhoff and Bunsen, is the selective absorption of light from the hot solar surface that is directed toward Earth, by the cooler atoms in the Sun’s atmosphere. By 1929, the chemical composition of the Sun was known as a result of the pioneering work of Hale and Russell and many others. That the Sun rotates in a period of about a month was first inferred from the observation of sunspots. A more complete summary of the knowledge of the Sun’s properties, obtained before the era of space age astronomy, can be found in several readable texts such as Menzel (4) and Kiepenheuer (5). Early Solar Physics by Meadows (6) contains an interesting review of ideas about the Sun in the period 1850–1900 and has several original papers. Hufbauer (7) traces the development of knowledge about the Sun from Galileo’s time. Space-based observations of the Sun are complemented by very significant ground-based observations. From the latter group, we discuss the following: *

*

*

The discovery of radio emissions from the Sun in 1942 during the development of radar, which led to the study of solar radio bursts by using radioheliographs. The discovery of neutrinos from the Sun whose flux was significantly below theoretical expectations; this discrepancy is driving a revision of our understanding of stellar interiors and also fundamental particle physics. [See notes added in proof.] The development of helioseismology, beginning with the discovery of 5minute oscillations of the Sun in 1960, which led to the development of the Global Oscillation Network Group (GONG) project, the Michelson Doppler Imager/Solar Oscillation Investigation (MDI/SOI), and, the Extreme-Ultraviolet Imaging Telescope (EIT) instruments on the Solar and Heliospheric Observer (SOHO) satellite.

Observations of the Sun from Space. Table 2 is a chronological list of the principal space missions, which have, or are expected to, provide significant advances in our knowledge of the Sun. Here we give a brief summary of the main facets of the Sun that a given mission is intended to study. In the following 1

See Asimov’s Biographical Encyclopedia of Sciences and Technology, Doubleday, Garden City, 1972, for biographical sketches of ancient to modern scientists.

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Table 2. List of Space Missions Relevant to Solar Physicsa Mission name Sounding rockets Vanguards Solrads Orbiting Solar Observatories (OSO) Mariner 2 Skylab Helios 1 Helios 2 ISEE3/ICE P78-1 Solar Maximum Mission (SMM) Hinotori Space Shuttle GRANAT Ulysses Gamma-1 Compton Gamma-Ray Observatory (CGRO) YOHKOH Solar and Heliospheric Observer (SOHO) Advanced Composition Explorer (ACE) Trace

Launch date 1946 1959 20 June 1960 1962 27 August 1962 1973 1974 1976 1978 1978 14 February 1980 21 February 1981 1982 1 December 1989 6 October 1990 1991 5 April 1991 30 October 1991 2 December1995 25 August 1997 2 April 1998

Termination date Continuing 1975 1975 14 December 1962 1974 1984 1981 1982 1985 3 December 1989 Continuing 1974 Continuing 4 June 2000 Continuing Continuing Continuing Continuing

a

The National Oceanic and Atmospheric Administration (NOAA) maintains regular observations of solar emissions, X-rays, and particles, with its GOES satellites. [See notes added in proof.]

sections, we describe in more detail some of the observations or discoveries made on these missions. The first observations of the Sun from space were made by Naval Research Laboratory (NRL) scientists in 1946 using ultraviolet (UV) and X-ray spectrometers carried on V-2 rockets captured from the German military in 1945. Further observations were made of solar UV and X-ray emissions using Viking and Aerobee rockets. After 1980, Black Brant rockets were used as well as the Aries and the second stage of the Minuteman I ICBM. Sounding rockets of various types are still used today to obtain—mainly UV, extreme UV (EUV), and X-ray spectra of the Sun using precise imaging telescopes. Sometimes, sounding rocket instruments are used for cross-calibration of instruments carried on satellites such as SkyLab or the Solar Maximum Mission. (A discussion of some of the early results is given in Reference 8.) ˚ or 121.6 nm) The first satellite observations of solar Lyman alpha (1216 A ˚ and soft X rays (1–8 A or 10–80 nm) were made by NRL from SolRad I during a period of 5 months after launch on 10 June 1960. Additional SolRad satellites produced continuous data on UV and soft X rays through 1975. Starting in 1962, NASA launched its Orbiting Solar Observatory (OSO) Series, which made many important observations of the quiet and active Sun through 1975. The first white light coronagraph and high-resolution gamma-ray spectrometer were flown on OSO 7, launched in 1971 and operated for 15 months.

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During its 4-month journey to Venus in 1962, Mariner 2 made comprehensive and definitive observations of the solar wind. Following OSO 7, the SkyLab (the last phase of the manned Apollo program) was launched in 1973 and continued into 1974. Imaging soft X-ray and EUV telescopes made pioneering observations related to the solar wind (SW) and coronal holes. (See also Reference 9 for a brief review of early results from the SolRads, OSOs, and SkyLab.) From 1974 to 1976, the European Helios 1 and 2 solar-orbiting satellites made plasma and high-energy electron and ion observations at distances from the Sun as close as 0.3 AU. From 1978 to 1982, the International Sun Earth Explorer 3 (ISEE3) was at the Sun–Earth Lagrangian point (L1) and made numerous valuable solar flare observations. As the International Comet Explorer (ICE), it continued flare observations beyond 1982. In 1978, the U.S. Air Force launched the P78-1 satellite that carried a white light coronagraph (Solwind) and observed many coronal mass ejections (CMEs). In the early 1980s, the P78-1 mission unceremoniously ended when it became a target in a ‘‘Star Wars’’ test. On 14 February 1980, the Solar Maximum Mission (SMM) satellite was launched. It carried six instruments for a coordinated study of solar flares, at the maximum of sunspot cycle 21, covering several spectral regions: the UV, soft and hard X rays with the hard X-ray burst spectrometer (HXRBS), and gamma rays, with the gamma-ray spectrometer (GRS). The instrument payload included a cavity radiometer for measuring the total solar luminosity or ‘‘solar constant.’’ In April 1984, SMM was taken onboard the space shuttle Challenger to replace gyroscopes and to refurbish some instruments. SMM was returned to orbit and made important flare observations during the rising phase of sunspot cycle 22. Its mission was ended on 2 December 1989 by orbital decay caused by expansion of earth’s atmosphere due to heating from solar UV and X-ray emissions, the very radiations it was designed to study! Early results from all SMM instruments are described in a report of an SMM workshop (10). The Japanese Hinotori solar-observing satellite, launched on 21 February 1981, made many observations of solar flares, some of which overlapped SMM/ GRS observations. (A review of high-energy solar flare observations during this period is given in Reference 11.) In 1982, the manned Space Shuttle program began. Several small specialpurpose solar observations were made from the SPARCS satellites, which were developed earlier for sounding rocket flights. SPARCS was deployed during a 7 to 10-day mission by shuttle astronauts, coorbited with the shuttle, and then was retrieved at the end of mission. It was returned to Earth, and a new solar instrument was installed, ready for another shuttle launch. The major Shuttle mission of importance to solar physics was the launch of Spacelab 2 on 29 July 1985. Spacelab 2 carried three solar telescopes and several other instruments with two solar physicists, Loren Acton and John-David Bartoe, to operate the solar telescopes. (See Reference 12 for a complete history of the Spacelab program.) The Russian Granat satellite was launched in December 1989 carrying the French/Russian SIGMA, a new type of imaging gamma-ray telescope and a

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scintillator gamma-ray spectrometer, PHEBUS, that responded omnidirectionally. Therefore, PHEBUS could detect many gamma-ray flares, but not with the same efficiency as the SMM/GRS. In October 1990, the ESA Ulysses was launched. Its mission is to study the high-latitude heliosphere, that is, above and below the plane of the ecliptic. It also carried instruments to study the solar wind composition and charge state and an X-ray spectrometer to study solar flares. A review of the results from the first orbit of Ulysses around the Sun was presented by Reference 13. The Russian spacecraft GAMMA-1, operational during 1991, made valuable observations of two high-energy flares. On 5 April 1991, the Compton Gamma-Ray Observatory (CGRO) was launched. Its primary mission was to study cosmic sources of gamma rays with four instruments, which covered the photon energy range from a few keV to 30 GeV. However, as a secondary goal, all CGRO instruments can make solar flare observations and have made some significant discoveries. On 3 June 2000, CGRO underwent a controlled reentry over the Pacific Ocean to reduce the chance of debris impacting populated areas. The Japanese YOHKOH solar-observing satellite, launched on 31 August 1991, carried the soft and hard X ray imaging telescopes (SXT) and (HXT). The dramatic imaging capabilities of the SXT and HXT have revolutionized the study of solar activity. On 2 December 1995 the Solar and Heliospheric Observer (SOHO) was launched with 12 instruments for studying several phenomena of the quiet and active sun. Helioseismology, solar wind (SW) and coronal mass ejection (CME) observations from SOHO are discussed in individual sections later. The Advanced Composition Explorer (ACE), launched on 25 August 1997, is designed to study solar flare particle emissions, cosmic rays, the SW, and other heliospheric energetic particle phenomena. It is positioned between the Sun and Earth at the Sun/Earth neutral gravity point, L1. (For solar flare particle emissions, see the section on Space Weather, in this article.) The Transition Region and Coronal Explorer (TRACE) was launched on 2 April 1998 to image (to 1 arcsecond spatial resolution) the photosphere, the transitional region, and the corona at three EUV wavelengths and several UV wavelengths. The solar plasma can be studied at selected temperatures from 6000 K to 107 K.

STRUCTURE OF THE SUN The structure of the Sun is illustrated in Fig. 2 as a series of concentric spherical shells, or zones, which are understood to have different physical characteristics, as deduced from observations and the theoretical development of the standard solar model (e.g., the SSM described by Bahcall and Ulrich in (14a) and later by Bahcall and Pinsonneault (14b). The inner zone or core of radius B2  105 km, the region of energy generation by nuclear burning, is followed by the radiative zone of outer radius B5  105 km and then the convective zone of outer radius B7  105 km. The thin outer zone of the solar disk, which is called the photosphere, has an accurately determined thickness of B5  102 km and a

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Figure 2. A schematic of the Sun showing the relative location of various regions and their approximate thicknesses. Not drawn to scale. From Reference 1.

temperature of B5800 K. These zones lie within the Sun’s visible disk as measured at a wavelength of 500 nm, although we see only the photosphere. Figure 3, from Reference 15, shows the internal physical properties of the Sun predicted by the SSM. A new means of studying the interior of the Sun uses helioseismology techniques on the SOHO satellite that complement the GONG observations. Above the photosphere lies the solar atmosphere that has three distinct regions: the ‘‘cool’’ chromosphere (B4500 K) B1.5  103 km thick, sometimes called the color sphere and seen in some total eclipses followed by the transitional region B8500 km thick in which the temperature rises rapidly to B8000 K, and then the low-density, hot corona at a temperature of B1  106 km, also seen in a total eclipse. The expansion of the corona into the interplanetary medium (IPM) becomes the solar wind, which extends throughout the solar system. The Core and the Neutrino Problem. It was realized in the mid-1800s that if the source of the Sun’s energy was due to its gravitational contraction, the so called ‘‘Helmholz–Kelvin hypothesis,’’ then the Sun could not have existed longer than B3  107 years. However, by the 1930s, age dating of terrestrial rocks containing radioactive minerals indicated that they had existed for B1.6  109 years., so clearly gravitational energy is inadequate to account for the existence of the Sun and, in fact, all stars on the HR main sequence. That transmutation of elements might be a source of stellar energy was considered as early as 1920 by Eddington and others (see Reference 16). Numerous authors investigated the energy release from possible nuclear reactions, but a paper by Bethe in 1939 (17) provided details for the fusion, or nuclear burning, of

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12

10

1.0

1000

7

100

10

5 million K 0.8 L

0.6 L

10 g/cm

1 million K

3

6

10

Density 0.4

3

0.4 L

1 g/cm

5

10

Luminosity 0.2

0.2 L

8

0.1 g/cm

Pressure

10

10 3

0.6

Temperature

1

Density, g/cm

9

Luminosity, L/ L

Pressure, bars

0.8

10

8

10

10 million K

11

10

Convective zone

1.0 L

10

10

Radiative zone

Core

Temperature, °K

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3 4

10

0.1

0 3

7

10

10

0

0.2

0.4

0.6

0.8

0.01

1.0

Fractional radius, r /R

Figure 3. Internal compression. The variation of pressure, luminosity, temperature, and density with fractional radial distance from the Sun’s center (left) to its visible surface (right). At the Sun’s center, the temperature is 15.6 million Kelvin, and the density is 151 grams per cubic centimeter; the central pressure is 233 billion times that of the Earth’s atmosphere at sea level (one bar). Thermonuclear energy is produced in a core region that extends to about one-quarter (0.25) of the solar radius; the core contains almost half of the Sun’s mass. The convective zone begins at 0.71 of the solar radius where the temperature has dropped to about 2 million Kelvin and the density has fallen to about 0.2 grams per cubic centimeter; the convective zone comprises about 2% of the Sun’s mass. In the photosphere, the temperature is 5780 Kelvin, and the pressure and density have dropped off the scales of the graph. [Prepared from the standard solar model data computed by John Bahcall and Marc H. Pinsonneault, Rev. Mod. Phys. 64: 885–926 (1992). From Reference 15.]

hydrogen to helium via the CNO cycle and the p–p chain. These two pathways for producing the Sun’s energy are shown in Table 3, although there are three alternative p–p chains. The basic CNO cycle, in which the heavy elements act only as catalysts, was suggested in 1938 independently by Bethe and von Weizsaecker. These reactions convert four hydrogen atoms into one helium nucleus that results in a mass loss of B4.8  10  29 kg, or an energy release of 4.3  10  12 J, or 26.8 MeV, for every helium nucleus produced by ‘‘burning’’ B6.7  10  27 kg of hydrogen. From this basic fact one can estimate that the solar luminosity of B4  10  26 J/s would require consuming B6.2  10  11 kg of hydrogen per second. Considering the Sun’s mass of B2  10  30 kg, which is currently mostly hydrogen, nuclear burning could easily power the sun for billions of years. In a sense, the transmutation of hydrogen into helium is very inefficient because only about 0.7% of the available mass can be converted into energy to supply the solar luminosity; however, no more efficient energy source

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Table 3. The pp Chain and the Carbon–Nitrogen Cyclea Reaction number Proton–Proton Chain I-1 I-2 I-3 I-4

Reaction p þ p ! 2 H þ e þ þ ne p þ e þ p ! 2 H þ ne 2 H þ p ! 3 He þ g 3 He þ 3 He ! 4 He þ 2p

Neutrino energy, MeV 0–0.420 1.44

II-5 II-6 II-7

3

He þ 4 He ! 7 Be þ g e þ 7 Be ! 7 Li þ ne 7 Li þ p ! 4 He þ 4 He

0.816, 0.383

III-8 III-9

p þ 7 Be ! 8 B þ g 8 B ! 8 Be þ e þ þ ne 8 Be ! 4 He þ 4 He þ g

0–14

Carbon–nitrogen cycle

p þ 12 C ! 13 N þ g 13 N ! 13 C þ e þ þ ne p þ 13 C ! 14 N þ g p þ 14 N ! 15 O þ g 15 O ! 15 N þ e þ þ ne p þ 15 N ! 12 C þ 4 He

0–1.20

0–1.73

a

From Reference 23.

has been suggested. A direct experimental proof that the nuclear burning is currently in progress seems to have been presented now (see later and note added in proof). Because the weak interactions in the p–p chain produce neutrinos, it was clear (see e.g., Reference 18 and 23) that detecting them with a flux consistent with theoretical predictions could confirm that nuclear fusion reactions were the source of the Sun’s energy. All of the reaction cycles shown in Table 3 produce neutrinos, but these were not all considered in the early papers about nuclear fusion as the source of stellar energy (see, e.g., Reference 17). The net result of hydrogen burning is 41H-4He þ 2e þ þ 2ne. Because the theory of beta decay was untested (the neutrino was not observed until 1956 at the Oak Ridge nuclear reactor), Pontecorvo (in 1946) and Alvarez (in 1949) proposed detecting neutrinos by the reaction 37Cl(ne, e  )37Ar. However, the possibility that (the expected) solar neutrinos could be detected eventually led Davis to develop a large neutrino detector consisting of B100,000 gallons of cleaning fluid (CCl4).2 A history of the development of this neutrino detector, other experimental possibilities, and related theoretical work is described well by Bahcall and Davis (19). To understand how this detector relates to the neutrinos produced by nuclear fusion, according to the Standard Solar Model (SSM), Bahcall and Shaviv (20) had deduced the neutrino flux at Earth versus the energy of the neutrinos 2

Note that the neutrinos detected by chlorine are ‘‘electron neutrinos,’’ ne, but there are at least two other types of neutrinos (see below).

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Flux, cm

_2

s

_1

1012 1011

Water

Chlorine

Gallium pp

1010 109 108

13

N

13

Solar neutrino spectrum

7

Be

O

107 17

F

106 105

6

7

Be

104 103 102 101

pep 0.1

B

hep

1

10

Neutrino energy, q, MeV

Figure 4. Solar neutrino spectrum. This figure shows the energy spectrum of neutrinos predicted by the standard solar model. The neutrino fluxes from continuum sources (like pp and 8B) are given in the units of number per cm2 per second per MeV at one astronomical unit. The line fluxes (pep and 7Be) are given in number per cm2 per second. The spectra from the pp chain are drawn with solid lines; the CNO spectra are drawn with dotted lines. The upper abscissa scale shows the energy range, from threshold, to which three types of neutrino detectors are sensitive. From Reference 21.

from the different reactions shown in Table 3 before the first results from the chlorine detector had been reported. As can be seen in Fig. 4, from Bahcall (21), the chlorine detector can respond to all neutrinos whose energies are above 0.814 MeV, which can come from 8B, from electron capture by 7Be, or from the rare proton–electron–proton (pep) reaction. The chlorine detector has been operated, from 1967 to 1993, in the Homestake gold mine in South Dakota. The first results from the chlorine detector (22) showed that the CNO cycle produces less than 9% of the Sun’s energy. Later results are shown in Fig. 5, a plot of the measured flux of neutrinos from late 1986 to late 1992 (23). The average value of the measured neutrino flux was 2.3270.22 SNU, where the solar neutrino unit (SNU) corresponds to one neutrino capture per second per 1036 atoms of 37Cl3. This flux corresponds to a production rate of 0.43770.042 atoms of 37Ar per day, the units on Fig. 5. The most recent theoretical neutrino flux value, based on the current solar model, is þ 1:4 9:31:2 SNU. So there is a clear discrepancy, raising the question, Is the experiment or the SSM and basic neutrino physics wrong? This question and results from three new neutrino experiments, in addition to the Davis results, are discussed in Bahcall (21). Because all of the experiments disagree with the theory, three solar neutrino problems are identified. The discrepancy between the chlorine result and theory is the ‘‘first’’ solar neutrino problem. Now, as shown in Fig. 4, neutrinos whose energy is higher than 7.5 MeV 3

To consider detection of neutrinos by target nuclei other than 37CL, 1 SNU is more generally defined as 10  36 interactions per target atom per second ((21), p. 201).

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109

1.5 90

91 94

99 100

92

97 95

1.0

106

119

98

112 102

116 120122 115 5 1 114 118

105 107 103 96

0.5

104

125 124 126

111 110

108

113 121

37Ar

production rate, atoms/day

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0 1986

1987

1988

1989

1990

1991

1992

Figure 5. A plot of the 37Ar production rates versus time. This plot summarizes the data obtained after the Homestake experiment resumed operation in September 1986. Run numbers are shown above each plotted point. Not plotted are run 93 for SN1987A and run 117 for the solar flare of 4 June 1991. From Reference 23.

from 8B decay can be detected by neutrino–electron scattering. Such a reaction can be detected by the Kamiokande II (K II) 4500-ton pure water detector, that records the (Cherenkov) light in a large array of photomultipliers from electrons knocked out of H2O molecules by neutrinos. After 1000 days of operation, this experiment reported in 1991 a neutrino flux of 0.4470.06 SNU, about 2.5 times þ 0:17 below the predicted value of 1:00:14 SNU. So both experiments confirm a deficit in the expected solar neutrino flux. However, the chlorine and Kamiokande results are incompatible with one another, and Bahcall identifies this as the ‘‘second’’ solar neutrino problem (see the left and central bar graphs in Fig. 6).

+6

+1.2

137 _7

9.3_ 1.4 +0.14 1.0_ 0.17

79 ± 12 69 ± 13 0.44 ± 0.06 2.55 ± 0.25

Ga (E > 0.2 MeV) H2O (E > 7.5 MeV) _ p p, pep Experiment

Cl (E > 0.8 MeV) Theory

7Be 8B

CNO

Figure 6. Comparison of measured rates and standard model predictions for four solar neutrino experiments. From Reference 21.

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The SAGE and GALLEX experiments use gallium as the detector and can detect neutrinos whose energies are 40.2 MeV. The average for the gallium exþ6 periments is 74 SNU, but a rate of 1377 SNU is expected (bar graph on the right in Fig. 6). This conflict is the ‘‘third’’ neutrino problem. (‘‘As far as the theoretical predictions are concerned, several teams have developed computer codes for predicting the current neutrino flux and, when using the same input data, obtain results that agree to within a few percent’’) (see Reference 24). Bahcall (21) also concludes ‘‘that at least three of the four operating solar neutrino experiments give misleading results or else physics beyond the standard electroweak model is required to change the neutrino energy spectrum (or flavor content) after neutrino production.’’ Earlier, Bahcall and Bethe (25) argued that the SSM is valid and that any nonstandard solar model consistent with the K II data predicts a Homestake detection of 4 SNU, which is too high. So if Homestake is correct, then neutrinos are missing. Now, the neutrinos from the p–p chain are electron neutrinos, whereas at least two other flavors of neutrinos, muon and tau neutrinos, are known. If the electron neutrinos are mixed with some other type of neutrino that has a small mass, then they could change their form in transit from the solar core to the detector. This phenomenon, known as ‘‘neutrino oscillations,’’ could be the answer. This and several other possibilities are discussed in Bahcall (21). Finally, non-standard models of the solar interior (mentioned before) can possibly be tested with helioseismological results from SOHO and GONG. They can detect ‘‘p-oscillation’’ modes of the solar interior and possibly ‘‘g’’ modes as well. The latter have amplitudes in the solar interior that are larger than the ‘‘p’’ modes and relate more closely to the neutrino calculations. [See note added in proof.] Radiative and Convective Zones. The energy produced in the solar core by nuclear reactions is initially in the kinetic energy of the charged reaction products and some in the form of gamma rays of BMeV energies. Tracking the energy release in the core to its eventual release into space, to give the solar luminosity, is now done by sophisticated computer programs that use the observed properties of the Sun as constraints. According to the SSM mentioned before, the energy released in the core is transported by radiation through the radiative zone, hence its name, by repeated scattering of high-energy photons as they degrade in energy until they reach the bottom of the convective zone B106 years after the original gamma ray left the core! The radiative opacity of the solar material is an important parameter input to these calculations (see Reference 26 for a basic definition). The energy at this point is transported to the solar surface by the actual movement of hot gas that rises to the surface, a physical process completely different from energy transport by radiation. This process is convection, the rise of hot gas to the photosphere through a hierarchy of cells of different size so that the cooler gas above sinks and there is a net flow of thermal energy to the solar surface. The variations in temperature, density, and pressure as energy moves outward from the core through these zones are shown in Fig. 3, according to recent SSM calculations (14b). (See also Reference 27 for a clear treatment of heat transfer in the convective zone.) Photosphere. At optical wavelengths, the photosphere defines the Sun as we see it, and much of the activity of the Sun can be seen by features that change in the photosphere, such as sunspots. As just mentioned, all of the energetic

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photons from the core have been absorbed by the time they reach the outer boundary of the radiative zone (see Fig. 2), and the energy they carry is finally transported convectively upward to the photosphere where the gas density is too low to sustain convection. The photospheric surface then radiates essentially as a blackbody because it has been heated to an average temperature of B5800 K. A white light image of the full solar disk actually shows a limb darkening. This is understood as a result of the rapid fall in temperature with height, so that at the limb, the cooler gas at a higher altitude radiates at lower intensity than at disc center. There, the visible radiation comes from a greater photospheric depth where the gas is hotter and its radiant intensity is higher. According to this picture, the lower level of the photosphere has a temperature of B6200 K, the upper level B5400 K, and the average disk temperature is B5800 K. A study of the photospheric absorption spectra, combined with theoretical models, has led to a determination of the composition of the Sun. Our current knowledge on this important topic also involves a study of spectra of solar prominences and sunspots and direct measurements in the solar wind, topics to be covered later. Table 4 lists the abundances of elements up to calcium (Z ¼ 20 (28)), although most elements up to thorium (Z ¼ 90) have a measurable photospheric value. The value A in the table is the power of 10 of the element’s abundance based on reference to a hydrogen abundance of 1012. The solar composition given

Table 4. Recommended Values for Solar Abundancesa E1

AE1

Comments

1 2 3 4 5

H He Li Be B

12.0 10.9 1.0 1.1 2.3

Reference element Prominence, flare part., solar wind Photosphere, spot Photosphere Photosphere

6 7 8 9 10

C N O F Ne

8.7 7.9 8.8 4.6 7.7

Photosphere Photosphere, corona Photosphere, corona Spot Corona

11 12 13 14 15

Na Mg Al Si P

6.3 7.6 6.4 7.6 5.5

Photosphere, Photosphere, Photosphere, Photosphere, Photosphere,

16 17 18 19 20

S Cl Ar K Ca

7.2 5.5 6.0 5.2 6.3

Photosphere, corona Photosphere, spot Corona Photosphere Photosphere, corona

Z

a

From Reference 28.

corona corona corona corona corona

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by ‘‘mass fractions’’ used in the SSM calculations is typically X ¼ 0.71, Y ¼ 0.27, and Z ¼ 0.02 for the hydrogen, helium, and heavier element abundances, respectively (14). An up-to-date treatment of the topics covered here can be found in Solar Astrophysics (27). Granulation. The upper portion of convection cells can actually be observed, as was done at Mount Wilson which showed Doppler shifts of Fraunhofer lines from individual granules. Excellent observations were made by instruments on the Skylab in 1973. These observations show that the visible solar surface has a granular appearance that continually changes; bright granules move upward, as shown by a small blue Doppler shift of certain spectral lines corresponding to a velocity of B1 km/s. The darker portions of this granulation, when viewing the same spectral lines, show a redshift, indicating cooler material that is falling back into the convection zone. The granules, of typical size B1000 km, are in continual motion and have a lifetime of several minutes. Thus, it is believed that this granulation corresponds to the upper layer of the convection zone. But on a larger scale size of B30,000 km, a flow pattern is also seen that is similar to the granulation, where similar upward and downward motion of gases occurs at the centers and edges of the cells, respectively. It is believed that these larger features, called supergranulation, are the imprint on the photosphere of larger convective cells deeper in the convective zone (27,29). Active Regions. Historically, ‘‘activity’’ on the sun was recognized by visually observing sunspots, which changed in number apparently in an apparent cyclic manner, and the bright patches, called faculae, seen close to sunspots when observed near the solar limb. Today, areas on the photosphere, where spots, faculae, other features, called plages, and dark filaments appear together, are called active regions. Transient phenomena such as solar flares and prominences are manifested in these regions, hence the appellation—active. Figure 7 shows an excellent high-resolution image of faculae and a sunspot near the solar limb at a wavelength of 575.5 nm taken at Pic du Midi, France. The activity of the Sun, which will be discussed further later, is closely related to, and probably due to, solar magnetism that is evident most strongly in sunspots. The magnetic field on the Sun was detected by Hale in 1908, using the Zeeman effect, which was discovered in the laboratory in 1896 by Pieter Zeeman. Hale discovered the sunspot magnetic field in 1912, again using the Zeeman effect. The sunspot field can be as strong as several thousand gauss. The central darkest area of a sunspot is called the umbra, and the surrounding annular ring the penumbra. (See Reference 30 for excellent photographs of all features of active regions.) Sunspot Number and Magnetic Polarity Cycles. Many terrestrial inhabitants are aware of the relationship between the spots on the Sun and impressive auroral displays and inconveniences such as high-frequency communication disruptions and power outages. Fewer are aware of the roughly periodic, cyclic nature of the number of spots. The sunspot cycle was discovered in 1843 by a German amateur astronomer, Heinrich Schwab, who, after only two decades of observation, stated, ‘‘The total number of sunspots has a period of about 10 years.’’ After B1850, the Sun was observed out on a regular basis, so the study of the cycle reported by Schwab could be studied in detail. Actually, sunspots often appear in groups that can sometimes last for a 27-day solar rotational

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Figure 7. High-resolution image of faculae and sunspots near the limb (upper center), obtained at the Pic du Midi Observatory, France, at a wavelength of 5750 with filter of ˚ passband. By permission of R. Muller. 100 A

period, the time for a solar rotation relative to moving Earth. A convenient way to understand the characteristics of the sunspot cycle is illustrated in Fig. 8. In the lower part of the figure, the sunspot areas are shown as a function of time. This is simply a time plot of the total area of all of the sunspots on the disc at a given time. The cyclic time behavior of the area is very similar to a plot of sunspot number versus time, in which the time between the peaks is B11 years, although the interval could be a few years shorter or longer. A plot of the relative sunspot number, which is actually a running mean, typically peaks a year or two before the area plot. As pointed out by Zirin (30), sunspot area is a more significant indicator of the degree of solar activity than number. At the top of Fig. 8, the latitude on the sun where a spot appears is plotted versus time. This type of plot, known as Maunder’s butterfly diagram, shows that as a cycle proceeds, spots appear at lower and lower latitudes; the first spots of a cycle appear near 301 N and S latitudes, and the last ones of the cycle appear near the solar equator. This latitude behavior, named after its discoverer, is called Sporer’s law. In this plot,

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Sunspot areas

Sunspot latitudes

1880

1890

1900

1910

1920

1930

1940

1950

1960

1970

40 30 20 10 0 _ 10 _ 20 _ 30 _ 40 6000

300

400

500

600

800 900 1000 1100 1200 1300 1400 1500 1600 CARRINGTON rotation number

700

4000 2000 0 1880

1890

1900

1910

1920

1930

1940

1950

1960

1970

Figure 8. Butterfly diagram (compiled by J.A. Eddy) showing the drift of sunspot position toward the equator. The length of each cycle is about 12 years, causing the old and new cycles to overlap (HAO). From Reference 30.

the length of a cycle is about 12 years and has a small overlap of the end of a cycle and the beginning of the next cycle. What about the behavior of the magnetic properties of the sunspots? If one considers a pair of sunspots in a given hemisphere during a particular sunspot cycle, the leading spot moving west will have one magnetic polarity (N or S), and the following spot has the opposite polarity, but this pattern is reversed in the other hemisphere. In the next solar cycle, the polarities are also reversed, which means that two sunspot cycles, or 22 years, must elapse before the orientation of the magnetic polarity is the same again. This characteristic of sunspots, discovered in 1912, is known as the Hale–Nicholson law. How are we to understand this fascinating behavior of sunspots? The most obvious features of the magnetic characteristics of active regions are the strong local magnetic fields, the Hale–Nicholson polarity laws, Sporer’s law, and the 22-year magnetic cycle. Forty years ago, Horace Babcock, who developed the first solar magnetograph at the end of the 1940s with his father Harold, proposed a ‘‘solar dynamo’’ model that explains qualitatively the four features just mentioned (31). The essential ingredients consist of a weak B1gauss solar surface field due to an axisymmetric dipole and a rotating Sun that has a convective zone that rotates faster at the equator than at higher latitudes. According to Babcock, it is assumed that the total magnetic flux for the equivalent dipole field, which extends over all space outside the Sun, is B8  1021 maxwells (or B8  1015 webers) and lies in a thin convective layer about 0.1 R thick just below the photosphere. However, the current view is that the dynamo lies in a transitional region between the convective and radiative zone called the tachocline (see, e.g., References 32 and 33). The differential rotation of the Sun with latitude has recently been determined directly by the Michelson Doppler Imager (MDI) on the SOHO satellite.

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24 0° latitude

Period of rotation, days

26

30° latitude

28

30 60° latitude 32

0.4

Core

Radiative zone

Convective zone

0.8

0.6

Surface

1.0

Fraction of sun's radius

Figure 9. Internal rotational rate of the Sun at latitudes of 0, 20, and 601 has been inferred using data from the Michelson Doppler Imager on the SOHO space craft. Down to the base of the convection zone, the polar regions spin more slowly than the equatorial ones. Below the convection zone, uniform rotation appears to be the norm, although scientists have not yet determined rotational rates within the Sun’s core. From Reference 34.

Figure 9 (from Reference 34) shows that the rotational period, as a function of the fractional radius of the Sun, in the radiative zone that extends to B0.7 R from the Sun’s center is the same at all latitudes, but in the convective zone, the period varies with latitude. The belief that the convective zone is in differential rotation is strikingly confirmed by MDI measurements. Babcock suggests that if the dipole field is initially normal, about 3 years into a new sunspot cycle, as a result of the frozen-in subsurface field and convective zone differential rotation, ‘‘a spiral wrapping of five turns in the North and South Hemispheres (occurs).’’ The field lines become stretched and twisted leading to a greatly enhanced field up to a factor of 45, depending on latitude. ‘‘Twistingy by the faster shallow layers in low latitudes forms ‘ropes’ with local concentrations that are brought to the surface by magnetic buoyancy to produce bipolar magnetic regions (BMRs) with associated sunspots and related activity.’’ The essence of Babcock’s original model is shown in Fig. 10, where the opposite polarities of the BMRs in the Northern and Southern Hemispheres are clearly shown. (Excellent summaries of Babcock’s model, which still gives a clear qualitative picture of the magnetic cycle, may be found in References 27 and 29.) Chromosphere and Transitional Zone. The fortuitous equality of the angular diameters of the Sun and Moon permits astronomers to see the relatively thin (B1500 km thick) and cool (B4500 K) lower atmosphere of the Sun for a few seconds before and after the total phase of an eclipse. The fact that this layer, the chromosphere, appears reddish in color is the reason that it is sometimes called

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N

f p

E

W

f p

S

Figure 10. Bipolar magnetic regions (BMRs) are formed where buoyant flux loops of the submerged toroidal field are brought to the surface after several solar rotations have twisted the original global field. The BMRs continue to expand, and flux loops rise higher into the corona. From Reference 31.

the ‘‘color-sphere.’’ The color is due to the Balmer Ha red emission line at a ˚ (656.3 nm). Figure 11, based on model calculations of wavelength of 6563 A Vernazza et al. (35), shows a schematic diagram of the temperature versus height above the outer edge of the photosphere.4 Thus, the temperature first falls, as expected, for the first 500 km, reaches the ‘‘temperature minimum,’’ and then slowly climbs to B7000 K at an altitude of 2000 km, after which the temperature rises abruptly and reaches B500,000 K. The region where the thermal gradient abruptly increases is called the ‘‘transition zone.’’ Beyond the transitional region is the corona where the temperature continues to rise above a million K. Actually the beginning of the chromosphere is defined as just above the ‘‘temperature 4

The ‘‘edge’’ is defined to be at an ‘‘optical depth’’ equal to 1 for a wavelength of 500 nm. Here, a photon will have a B37 % chance of escaping to space.

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Quiet-Sun EUV brightness components 30 Mg || k line k1

k2

k3

Ca || K line 20

K2

K3

K1

H
(core)

H
(wing)

3 cm 1 cm 3 mm 600 µm

3

T, 10 K

1 mm 300 µm

10

150 µm

8

50 µm

C (109.8 nm)

L
(center)

Si (152.4 nm) L
(peak) 6

Fe, Si (157.5 nm) H (70 nm) Si (168.1 nm) H (90.7 nm)

L
(1 Å) L
(5 Å)

4 2500

2000

1500

1000

500

0

h, km _5

10

_4

10

10

m, g cm

_3

10

_2

10

_1

1

_2

Figure 11. The average quiet-Sun temperature distribution in the chromosphere and transition region derived from the EUV continuum, the La line, and other observations. The approximate depths where the various continua and lines originate are indicated. From Reference 35.

minimum’’ where the temperature rises relatively slowly until the rapid rise. Note that the abcissa is reversed and height increases to the left. Figure 11 also shows the approximate region in the chromosphere where different emission features arise. Remember that when looking at the solar disc, the strong Fraunhofer absorption features from the photospheric continuum that pass through the chromosphere are dominant in the solar spectrum. Conversely, during a total solar eclipse, viewing the chromosphere at the limb reveals a beautiful spectrum, the ‘‘flash spectrum,’’ devoid of Fraunhofer dark lines. This now becomes an emission spectrum, which enables studying the composition of the chromosphere. Actually the full interpretation of the flash spectrum is very complex and a treatment in depth is given in Zirin (30). At higher altitudes above

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the limb, unusual ‘‘coronium’’ lines are seen which were eventually identified as lines from highly ionized Fe and Ca. (These lines are discussed in the next section on the corona.) When observed in the red Ha line of hydrogen, for a few seconds during a solar eclipse, one can see prominences and spike-like jets called spicules. These features are always present above the solar surface but cannot be seen against the bright disc. The spicules can extend up to 10,000 km above the solar limb, have a thickness of about 900 km, and typically move upward at velocities of B25 km/s with lifetimes of B5 min. They are concentrated at the edges of the supergranulation cells. Of course, the immediate question concerning the chromosphere is how it (and the corona) can reach the higher temperatures shown in Fig. 11. The corona can also be observed in the absence of a solar eclipse using coronagraphs on Earth or in space. Corona. During a total solar eclipse, depending on the phase of the solar cycle, the light that extends well beyond the disc and the narrow chromosphere takes on different forms and is typically as bright as the full Moon (B10  6 of the sun’s brightness near the chromosphere falling to B10  8 at B2 R. This ‘‘crown’’ of light is called the solar corona. The most dramatic characteristics of the corona are its high temperature and its extremely low density, surprising for such a luminous object. The changing forms of the corona are associated with the sunspot cycle and the changing magnetic structure. Near the minimum solar sunspot number, the coronal intensity near the poles is depressed relative to that when the solar activity is high. The corona, observed during an eclipse, emits a continuous spectrum similar to the photospheric continuum but has no Fraunhofer lines and instead bright emission lines, first observed in 1869. It is believed that the continuous spectrum is due to the scattering of the disc continuum from free electrons (Thomson scattering). The lines, it was thought, are from a new element, ‘‘coronium.’’ Then in 1941, Grotrian and Edlen identified the lines as ‘‘forbidden’’ transitions from highly ionized atoms such as Fe X, Fe XIV, and Ca XV, which require million degree temperatures. So the increasing temperature of the solar atmosphere above the temperature minimum reaches a few million K—more than two orders of magnitude hotter than the photosphere. This surprising fact has yet to be explained. Several mechanisms have been suggested, but from 1948 until the late 1970s, sound waves produced by the material motions in the convective zone, it is believed, produce supersonic shock waves that have sufficient energy to heat the corona. However, space-based observations did not detect any shock motions at the lower edge of the corona; presumably, if the shocks are produced outside of the photosphere, they dissipate all of their energy in the chromosphere, explaining the temperature rise seen in Fig. 11. Alternative mechanisms all seem to involve ways of transforming magnetic field energy to heat in a low-density gas, and there is at least circumstantial evidence for an analogous mechanism. This comes from the Einstein Observatory observation of intense X-ray (therefore, hot) coronas around dwarf K and M stars, which may have strong magnetic fields, because large star spots have been detected on their surfaces. Therefore, it would be reasonable to expect that the Sun’s corona could be heated in a similar manner. (A very readable description of attempts to explain coronal heating is found in Reference 29.)

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The part of the corona described above that extends to B2 R is called the K (for ‘‘kontinuerlich’’) corona. Beyond 2 R, faint Fraunhofer lines (the F corona) from the photosphere are visible. This is the near-Sun enhanced ‘‘zodiacal light’’ produced by dust in the interplanetary medium. The corona is now studied in EUV and X-ray wavelengths from space using sounding rockets and satellites. The SKYLAB mission in 1973 obtained soft X-ray images of the inner corona that showed bright loops outlining magnetic field lines that connect opposite magnetic poles of a bipolar magnetic region. Solar Wind. We understand today that the solar wind (SW) is the dynamic expansion of the Sun’s hot, ionized corona, a plasma that moves radially away from the Sun at a speed of B400 km/s and extends throughout the heliosphere beyond Earth. The study of certain tails of comets led Ludwig Biermann in 1950 (36) to propose ‘‘solare korpuskular-strahlung’’ as the explanation for the deflection of their tails5 in the radial direction away from the Sun. This proposal credits Biermann with discovering the SW because we now know, from direct measurements in space, that it is composed predominantly of electrons, protons, alpha particles, and other ions from the Sun. The helium abundance of B4–5% is apparently enhanced in the corona compared to the photosphere. The SW has been studied intensively since the first space probes and planetary missions were launched. Its properties are continuously monitored by instruments on several spacecraft beyond Earth’s influence. The SW plasma properties vary with solar activity; the low speed and high speed winds have speeds of 327 and 702 km/s, respectively. For the average SW (speed B4687 116 km/s), the average flux of ions, directed radially away from the Sun, is B4  108 particles/cm2 s at 1 AV. The temperatures of the principal SW constituents, protons, electrons, and alpha particles, determined in the frame of the bulk plasma motion, are 120,000, 140,000, and 580,000 K, respectively. The alpha-particle abundance is highly variable in association with transient energetic particle events. An important feature of the SW, directly associated with the Sun’s largescale magnetic field, is the variability of its magnetic field. In particular, the radial component of the interplanetary magnetic field, swept out by the SW, switches polarity several times as the solar rotation sweeps the field past Earth in B27 days (the synodic period of the Sun’s rotation, i.e., relative to Earth.) indicating sectors of opposite polarity. The field lines of opposing directions, it is understood, are separated by a thin current sheet, lying approximately in the plane of the ecliptic, which is warped, so as the Sun rotates, the alternating polarities are seen. The sectors of opposite magnetic polarities are referred to as the ‘‘sector structure’’ of the interplanetary magnetic field. The pioneering work in developing a theoretical model for the SW was done by Eugene Parker (37). See also Parker (38). He assumed that the SW was driven by the thermal pressure of the hot coronal gas. The plasma in the corona is sufficiently hot that it is not gravitationally bound. This mechanism appears sufficient to account for the low-speed SW, but not for the high-speed SW. On the other hand, the origin of the high-speed SW seems to be directly associated with coronal holes. However, the origin of the SW is still not fully understood, and this 5

The cometary tails referred to here are ‘‘ion’’ tails as opposed to ‘‘dust’’ tails.

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remains one of the major areas of investigation in solar physics along with heating of the corona. (A detailed treatment of the solar wind can be found in Reference 27.) Coronal Holes. Apparently, coronal holes were first observed by X-ray telescopes carried on early sounding rocket flights. In 1973, a telescope on SKYLAB photographed the full solar disc in soft X rays (o1 keV). The image obtained (see Reference 39 for an impressive example) showed bright emission over and above most of the disc that extended into the corona. However, near the central meridian, a dark region extended from pole to pole. The impressive feature observed was a large coronal hole, a part of the corona, which is slightly cooler (B106 K) than the bright hotter B2–3  106 K regions. The brighter regions that generally cover the largest area of the corona are due to hot, X-ray emitting gas confined to closed magnetic loops whose foot points are anchored in the photosphere at bipolar magnetic sites or at more complex magnetic sites. The magnetic field lines of coronal-hole regions are open, that is, only one end of a line is tied to the photosphere at a unipolar magnetic site, and the field at several solar radii presumably becomes effectively radial, tied to the solar-wind flow. Using X-ray photographs of coronal holes and measurements of the magnetic fields at the photosphere, the high coronal magnetic field geometry is computer modeled, assuming that there are no electric currents above the photosphere, that is, that the coronal field is potential. When high-speed SW flows are extrapolated back to the Sun along field lines, it is found that the magnetic polarity of the stream in space corresponds to that of the coronal hole at the Sun. So the source of the high-speed SW is a coronal hole! As mentioned before, the Parker model of the SW assumes that it is driven by thermal pressure, but that is not sufficient to accelerate the fast SW. One possibility is the presence of Alfven waves that propagate along magnetic field lines and can, if sufficiently intense, carry the plasma along to form the fast SW. Another point of historical interest is that some geomagnetic disturbances (e.g., cosmic-ray variations) due to enhanced particle fluxes at Earth recurred in a 27-day period. ‘‘M regions’’ on the sun were hypothesized as the source of the corpuscular radiation that causes the disturbance (see Reference 40). These regions are now identified as low-density coronal-hole regions near the sun. (A detailed discussion of this topic and its connection to the SW can be found in Reference 41). During the first orbit of the Ulysses mission, 1992–1998, Keppler (13) in a recent review reports that there was minimal solar activity so that large polar coronal holes were present and the Solar Wind Isotopic Composition Spectrometer (SWICS) observed fast SW at speeds of 750–800 km/s at the higher latitudes. When Ulysses was at lower latitudes, above the so-called ‘‘streamer belt,’’ the slow SW at speeds of 300–450 km/s was observed, consistent with the view that it arises from closed magnetic regions. Apparently, polar coronal holes extend down to a latitude of 451 near the Sun (within o1 R of the photosphere). The Ulysses observations also indicate a continuous mass loss, during a solar minimum, of B6  108 kg/s that is miniscule compared to the solar mass of B2  1030 kg. Above 451 latitude, the fast SW makes the heliosphere extend approximately in the polar directions, but because the Sun’s magnetic dipole axis is B101 from its rotational axis, the extension must wobble in a B24-day sidereal

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period. It is interesting that the boundary between the slow and fast SW is sharp and occurrs within oB21 of latitude. The SWICS instrument has also shown that the chemical compositions of the regions are different (42). The Ulysses instruments provide considerable data on the properties of the Interplanetary Medium (IPM), particle acceleration in so-called ‘‘co-rotating interaction regions,’’ and the ‘‘anomalous cosmic rays.’’ Some of these topics are reviewed in Keppler (13). The Ulysses mission will continue until at least 2001, so we can expect many more valuable results related to solar activity and phenomena in the IPM. (See the section on Space Weather, in this article.) Solar Activity. Solar activity has traditionally referred to dynamic and transient changes in photospheric active regions, where faculae, filaments, plages, and sunspots are seen. Even solar flares, historically the most dramatic symbol of solar activity, result from subtle changes in loop magnetic fields coupled to the photosphere. Within the past 30 years, vast eruptions have been observed in which 20 billion tons of matter have been expelled from the sun in a single event, a coronal mass ejection (CME), that sometimes has disastrous effects on Earth. All of these phenomena result from what might more appropriately be called solar magnetic activity. Great advances have been made in understanding the causes of solar activity, but major new missions will be necessary to unlock the secrets of this aspect of the Sun. We will first discuss some important properties of flares in the next section and then, more briefly, CMEs near the Sun. Solar Flares and Prominences. In 1859, an intense brightening was observed near a large sunspot group by English observers Carrington and Hodgson; it was the first recorded report of a solar flare, or what is sometimes also called a solar eruption (see Reference 43 for an early history of astronomy). We know now that this type of transient event, which usually can last for minutes or hours, is caused most certainly by the release of energy that resides in the strong magnetic fields that are present in a plage/sunspot region. Before discussing the possible cause of solar flares, some of their general properties must be presented. First, a solar flare is the most energetic transient phenomenon that takes place in the solar atmosphere; the largest releases as much as 1032 ergs overall, although CMEs may release a comparable amount of energy. This is more than 2 billion times more energy than is released in a 1-megaton nuclear explosion. In addition, it is now known that a large flare emits a major fraction of its energy as electromagnetic radiation (photons), ions, and electrons. The photon energies range from about 10 microelectronvolts (meV) for meter-wave radio emissions, to electron volts, for optical emissions, and to a billion eV (GeV) for gamma rays. Major flares can produce both ions and electrons that have relativistic kinetic energies, as well as relativistic neutrons, that can reach Earth. The frequency with which flares occur follows the sunspot cycle and, as mentioned earlier, the more flares occur, the larger the total area of sunspots. Because the intensity of flare emissions appears roughly proportional to the area of the flare in the optical region of the electromagnetic spectrum, solar astronomers have developed a classification scheme to report on the ‘‘importance’’ of a flare. In particular, flares are routinely observed in the light of the red Balmer line of hydrogen, the Ha line whose wavelength is 656.3 nm. Now, the microwave radio flux and the soft X-ray intensity are also routinely measured as well. Table 5 (from Reference 30) shows the classification scheme used to report

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Table 5. Flare Classification by Areaa Area, square deg

Area in 10  6 A

r2.0 2.1–5.1 5.2–12.4 12.5–24.7 424.7

r200 200–500 500–1200 1200–2400 42400

Class

Typical flux at 5000 MHz, sfu

Typical SXR class

S 1 2 3 4

5 30 300 3000 30000

C2 M3 X1 X5 X9

a

From Reference 30.

the properties of an optical flare developed by the International Astronomical Union (IAU) in 1966. The first and second columns in the table refer to the observed area of the enhanced Ha emission region in terms of square degrees on the solar surface6 in the heliocentric coordinate system (column 1) or in millionths of the area of the solar hemisphere (column 2), and the third defines the optical class of the flare. Sometimes a letter, f, n, or b, follows the letter class indicating that the estimated visual brightness is faint, normal or bright, respectively. The fourth column gives the microwave flux at a frequency of 5 billion cycles/s (5 GHz) in solar flux units (1 sfu ¼ 10  22 watts/meter2/Hertz), and the fifth column gives the ˚ ) band, as measured by the NOAA GOES soft X-ray flux (SXR) in the (1–8 A spacecraft. The SXR class labels Bn, Cn, Mn, and Xn mean that the X-ray flux is 4n  10  7, 4n  10  6, 4n  10  5 or 4n  10  4 watts/m2, respectively. The number after the letter is a multiplier. As an example, if a major flare is designated as 4bX9, its Ha emission covers an area of 42400 millionths A, is bright, and has a soft X-ray intensity of 9  10  4 watts/m2 at Earth. In the flare literature, there are numerous different types of flares depending on their shape or what technique is used for observing, but in Ha, the designations compact and two-ribbon seem sufficient (44). Generally, flares occur in magnetic loops or flux tubes whose foot points are rooted in, or near, sunspots. Compact flares are small and apparently have a single loop, whereas two-ribbon flares are much larger and have two bright ribbons moving away from a single dark ribbon or filament. When viewed on the disc, a prominence appears as a dark filament. The larger flares can last for hours, and the brightest are seen as white light flares. The Ha flare is often called a chromospheric flare, or ‘‘cool’’ flare, as contrasted to the ‘‘hot’’ flare that occurs in the corona at a temperature of B107 K. (See an extensive discussion in Reference 45). It is instructive to see a flare in a wavelength other than in the standard Ha band; and both sounding rocket and satellite images are readily available now. On 11 September 1989, a sounding rocket carrying a high-resolution X-ray telescope, NIXT (Normal Incidence X-ray Telescope, the forerunner of TRACE), observed the solar corona in a narrow wavelength band centered at 6.35 nm, which includes X-ray spectral lines from highly ionized Fe XVI and Mg X atoms. 6

One square degree on the Sun corresponds to an area of 1.48  108 km2. One solar hemisphere has an area of 3.04  1012 km2.

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Figure 12. A high-resolution X-ray image of the solar corona that has active regions and a flare (indicated by the arrow). The angular resolution is 0.7500 , and the temperature of the brightest areas is about 3  106 K. This image was made by L. Golub et al. in 1989 from a rocket using the 0.25-m X-ray telescope NIXT (Normal Incidence X-ray Telescope) in a narrow wavelength range around 6.35 nm; this range includes the emission lines of Fe XVI at 6.37/6.29 nm and of Mg X at 6.33/6.32 nm. In contrast to an image-forming X-ray telescope of the Wolter type that has grazing incidence, here the image is produced at normal incidence as by an orinary optical mirror. Reflection of the X rays is made possible by vapor deposition of alternating thin layers of cobalt and carbon, so that constructive interference results for the wavelength 6.35 nm. NIXT was developed by L. Golub and coworkers at the Smithsonian Astrophysical Observatory. Cambridge, together with the IBM Thomas J. Watson Research Center, Yorktown Heights, New York (photo courtesy of SAO and IBM Corp.). (See Reference 46 for description of rocket flight.) This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

The coronal temperatures required to produce these lines is in the range of a few million degrees (B107 K). The image obtained is shown in Fig. 12 (from Reference 46) and shows the full solar disc where bright emission is across the several active photospheric regions. In addition, the most intense feature, northwest7 of the Sun’s center is a small two-ribbon subflare, SbC5, so the soft X-ray intensity is 5  10  6 W/m2 in the 10- to 80-nm band. This flare was observed by NIXT 2 minutes after the GOES monitoring satellite observed the peak of the X-ray emission and, judging from its brightness compared to the coronal emission across the several active regions, its temperature is probably 107 K. What is particularly interesting and significant about this observation is that the tangled coronal structure, controlled by the local magnetic fields, is undoubtedly due to similar magnetic complexity in the active photospheric regions.

7

Generally, in images of the solar disc, east is on the left, and west is on the right.

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The basic question of what causes a flare is still unanswered, but optical and magnetic observations of active regions and the morphology of two-ribbon flares have led to the general view that reconnection, or annihilation, of opposing magnetic fields occurs in an ‘‘arcade’’ of magnetic loops that are anchored in the photosphere. It is also now believed that the energy release occurs in the corona and there is adequate energy stored in sunspot magnetic fields so that only a small fraction of the existing fields need be annihilated to supply the energy for a major flare. (See the next section for a specific example—the Masuda flare.) There is a vast literature on solar flares that was developed before the space age began, and key observations of their electromagnetic emissions were made by both ground-based optical and radio telescopes. Much of the early and current knowledge about flares is expertly treated by Zirin in Astrophysics Of The Sun (30) and earlier by Svestka in Solar Flares (47). Recent discussions may be found also in Tandberg-Hanssen and Emslie (48) and Somov (49). High-Energy Flare Emissions. A singular event occurred in 1942, when several cosmic-ray detectors on Earth responded to an increased flux of particles, presumably produced in association with an intense flare that occurred 2 days earlier. If the increased ground-level cosmic-ray flux was due to protons at the top of the atmosphere, then the protons must have relativistic kinetic energies. This ‘‘solar flare effect’’ in ground-level high-energy cosmic-ray detectors recurred about every 4 years until after WWII, when cosmic-ray groups around the world began carrying instruments to high altitudes on balloons and the so-called solar cosmic rays could be studied at lower energies using other instruments. It is this aspect of solar flares that is most important to focus on because high-energy flare emissions produce the greatest effects on Earth and on spacecraft. Recognition of this fact led in 1963 to a major conference on ‘‘The Physics of Solar Flares’’ at the Goddard Space Flight Center (see Reference 50). This article concentrates mainly on recent results obtained from SMM, YOHKOH, CGRO, SOHO, and TRACE after some brief history. (See Reference 11 for a review of results from 1942–1995.) The dramatic increases in ground-level cosmic-ray intensity are often referred to as ground-level events (GLEs), enhancements of the count rate of a cosmic-ray intensity monitor. The first GLE observations were made using ionization chambers and Geiger–Mueller counters that detect predominantly the charged secondary components produced in the atmosphere by a spectrum of energetic extraterrestrial particles. After the end of World War II, neutron monitors were developed that respond to the secondary nucleonic component produced by the lower energy particles in the spectrum incident at the top of the atmosphere. See Simpson (51) for a detailed discussion of this subject. Using this new technique, on 19 November 1949, a GLE was observed after a large solar flare. The intensity recorded during the maximum of the event was six times the normal value, a relative increase, which was considerably larger than those of earlier bursts. Major activity in the early part of 1950 was low; then on 26 February 1956, a large GLE occurred. This event became a powerful stimulus to research on high-energy charged particles and photons associated with solar flares. The neutron monitor increase showed that protons that had energies 415–30 GeV were produced in close temporal association with the flare optical and radio

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emissions. The protons that caused this GLE, it was believed were accelerated by the ‘‘Fermi mechanism’’ in the turbulent magnetic field produced by the chromospheric solar flare. The first observations of energetic photons from a solar flare were made on 20 March 1958 from a balloon-borne experiment. These detector increases coincided exactly with the maximum intensity of the optical flare (which began 5 minutes earlier) and with an 800 MHz microwave radio burst! Analysis of the individual detector rates indicated that the burst was caused by bremsstrahlung from B1035 electrons that had energies of B0.5 MeV. After the USSR launched Sputnik, (4 October 1957) and the US satellite program was initiated, Morrison (52) published a stimulating paper predicting the fluxes expected for gamma-ray emission from several celestial objects and other phenomena, including solar flares. Also in the post World War II period, the systematic study of radio emission associated with solar flare ‘‘eruptions’’ began. By the mid-1950s, radio spectrographs were observing solar radio bursts. Eventually, interferometers were added, that could give the spatial locations of the sources, and a concrete picture of the cause of the radio phenomena developed. This led in the 1960s to a two-phase model that explained that the radio bursts are due to accelerated B100-keV electrons followed in large flares by Fermi acceleration of protons and electrons to very high energies; this explained the GLE events and the low-energy, flare-associated protons detected by satellite, balloon, and riometer (relative ionospheric opacity meter) detectors. As will be discussed, this model existed until the 1980s before new observations complicated the picture. In the decade 1970–1980, the successful series of OGO and OSO satellites continued solar observations. Experiments on SKYLAB as well as rocket and balloon experiments were dedicated to study the Sun. Instruments on High Energy Astronomical Observatories HEAO-1 and HEAO-3, dedicated to cosmic observations, could also record high-energy solar emissions. In the early 1970s, the University California San Diego (UCSD) group’s X-ray spectrometers on the OSO 7 satellite obtained, copious X-ray observations in the range of 2–300 keV. An auspicious development in this period was the appearance of the complex active region (McMath 11976) on 11 July 1972 that heralded the occurrence of the August 1972 series of major flares that had concurrent intense high-energy emissions and associated terrestrial effects. On 2, 4, and 7 August 1972, major flares from a single active region produced the most intense and long-lasting radiations across the full electromagnetic spectrum observed until then that had associated, intense, long-duration charged particle fluxes in space. Fortuitously, during a 10-minute period in the rising phase of the 4 August flare, the OSO 7 gamma-ray spectrometer recorded, for the first time (53), an intense gamma-ray line and continuum spectrum, that dramatically confirmed the 1958 prediction by Morrison (52) that nuclear reactions in the solar atmosphere during flares could produce a 2.223 MeV line detectable on Earth. The spectrum obtained also gave evidence of nuclear lines at 0.511 MeV, 4.4 MeV, and 6.1 MeV and continuum photons to 10 MeV. The ground-based optical observations of the two-ribbon flares on 4 and 7 August are discussed in detail by Zirin (30).

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These observations implied that acceleration of protons above 30 MeV was closely associated with the production of the relativistic electrons that produced the bremsstrahlung continuum. This concept was not taken seriously though, probably because of the 3-minute time resolution of the observations. Therefore, the two-phase model, mentioned before, was still considered valid. Also, based on the published OSO 7 data, Svestka in his 1976 monograph on solar flares (47) proposed for the first time that electrons and ions were accelerated in the ‘‘same (one-step) acceleration process, or that the second-step acceleration (giving rise to 430 MeV protons) immediately follows the process of pre-acceleration.’’ Due to the launch of the SMM satellite on 14 February 1980 and the Japanese Hinotori satellite on 21 February 1981, the continuing International Sun– Earth Explorer/International Comet Explorer (ISSE 3/ICE) and Interplanetary Monitoring Probe (IMP) operations, and the coordinated ground-based solar observations of the Solar Maximum Year (SMY), an avalanche of new information on solar flares became available. Satellite observations by the hard X-ray and gamma-ray spectrometers on these new missions provided dramatic new diagnostic tools for determining the properties of particle acceleration associated with solar flares. Within the first few months of flare observations in 1980 by the gamma-ray spectrometer (GRS) on SMM, it became clear that acceleration of B50-MeV protons and relativistic electrons was simultaneous to within the instrument time resolution of 1–16 s (54), confirming the earlier suggestion of Svestka (47). Note that this observation is contrary to the two-phase acceleration model mentioned earlier. During an intense limb flare on 21 June 1980, the SMM GRS detected, for the first time, a burst of energetic neutrons at Earth, after a 1-minute-long burst of gamma-ray lines and electron bremsstrahlung that extended to more than 100 MeV in photon energy. This confirmed the 1951 prediction (55) that relativistic protons, accelerated during a solar flare, could produce a flux of high-energy neutrons observable at Earth. Then, during the impressive east-limb flare on 3 June 1982 that produced intense nuclear line emission, it was shown (56) that the photon spectrum above 10 MeV included contributions from meson-decay gamma rays and electron bremsstralung that extended to more than 100 MeV. Neutron monitors in Europe also recorded, for the first time, secondary neutrons produced in the atmosphere by primary solar neutrons (57) (see Fig. 13). The protons that resulted from the decay of solar neutrons in space were observed by ISEE 3/ICE and IMP detectors. Several flares observed by the SMM GRS show rich gamma-ray line spectra, from which it is possible, in principle, to determine the composition of the ambient solar atmosphere where the nuclear reactions occur. Analysis of the spectrum obtained for the 27 April 1981 flare, a long duration gamma-ray flare, indicates that the ambient medium composition differs from that of both the photosphere and the corona and requires an enhanced neon abundance. For this flare, it is suggested that significant gamma-ray line production could take place in the corona. In impulsive flares, most of the line emission is expected to peak in the lower chromosphere, and there is essentially no contribution from the corona. Figure 14 shows the spectrum for this flare calculated by Ramaty (58) that gives the best fit to the observed line spectrum. In general, the SMM GRS

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104 SMM _ Gamma-ray spectrometer 3 June 1982

103

103

102

102

100

101 MCW (4.1_ 6.4) MeV

100

102

102

Counts per second

Counts per second

X-ray (56 _ 199 keV) 101

HE-matrix (>25 MeV) 101 Sunset at satellite ~ 1204UT 100

100 Jungfraujoch neutron monitor

6

6

4

4

2

2

0

0

_2 11.40

_2 11.50

% deviation from mean, 8798 counts / min

% deviation from mean, 8798 counts / min

101

12.00

Universal time

Figure 13. Temporal history for several data channels from the SMM GRS and for the Jungfraujoch neutron monitor count rate for the 3 June 1982 flare. Peak count rates in the GRS (X-ray) and MCW (4.1–6.4) MeV energy bands are uncertain because of pulse pile-up, excessive dead time, and photomultiplier gain shifts and should be used with care. The highest MCW count rates have been estimated using measured live time values, averaged over 16.384 s, and a derived gain shift correction. Error bars are based on count statistics only. From Reference 57.

observations were confirmed by the gamma-ray spectrometer (GRS) on the Hinotori satellite. By the end of solar activity cycles 20 and 21 in 1986, all of the highenergy radiations predicted (52,59) had been observed; they revealed new insights into the solar flare particle acceleration enigma. Table 6 summarizes the principal characteristics of the emissions from high-energy solar flares. These may be considered observational constraints, which any theory of particle acceleration must satisfy. The temporal evolution of a flare may be

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Vol. 2 10 27 April 1981 best _ fit abundances

e+
_


n 20Ne

56Fe

1 cm _ 2

24Mg

Photons, keV

_1

28Si 20Ne

12C

16O

24Mg

0.1

15O,15N

16O,15N

0.01

0.001

0

2

6 4 Energy, MeV

8

10

Figure 14. Calculated solar flare gamma-ray spectrum corresponding to abundances that best fit the observed spectrum of the 27 April 1981 limb flare. From Ramaty (58).

separated into three phases: onset, impulsive, and extended phases. A particular flare may be of short duration (a few seconds or minutes) or of long duration (hours); each has emission characteristics shown in the three columns of the table. Table 7 shows the properties of the accelerated electrons and ions of the larger flares. After the demise of the SMM satellite, on 1 December 1989 the French/ Russian satellite, GRANAT, was launched and observed high-energy flare neutral emissions from two flares on 15 May, 24 May, and 11 June 1990. The Russian spacecraft GAMMA-1 provided the only gamma-ray data available on the intense flare at 20:28 UT on 26 March 1991. Since April 1991, additional solar flare observations were made by the four instruments on the Compton Gamma-Ray Observatory (CGRO). All CGRO instruments directly viewed three X-class flares on 9, 11 and 15 June 1991. Because the instruments on the CGRO, particularly, were much larger than those on the SMM and Hinotori, extended, hours long emissions of gamma rays and neutrons at low flux levels could be identified following some of these flares. The most dramatic example of extended emission was obtained by the CGRO EGRET, following the 11 June 1991 flare when high-energy gamma rays (4100 MeV) from meson decay and ultrarelativistic bremsstrahlung were observed continuing more than 8 hours after the peak of the flare (60). The PHEBUS spectrometers on GRANAT also recorded gamma-ray line and continuum emissions throughout the intense part of the flare when the EGRET spark chamber and the COMPTEL were disabled because of electronic dead time caused by intense X-ray flux in anticoincidence shields. The EGRET Total Absorption Spectrometer Calorimeter (TASC), however, could also function independently of the spark

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Table 6. Observational Constraints for Acceleration Theory of Temporal Evolutiona [I.] Onset phase

[II.] Impulsive burst

[III.] Extended emission

Duration is typically minutes

Sudden enhancement of bremsstrahlung to several hundred MeV

Impulsive events with simple decay

Brightening occurs in UV, soft X rays, or Ha

Simultaneous enhancement of g-ray line emission

Impulsive event and continued production of high-energy emissions

Type III burst at coronal height B(0.4  1)R

Occasional bursts of 410 MeV photons

Succession of impulsive events and g-ray line emission dominant over bremsstrahlung

Type I noise storms

Emission of decay g rays

Succession of impulsive events bremsstrahlung dominant over g-ray line emission

Increasing intensity of hard X-ray bremsstrahlung

Arrival of high-energy neutrons at Earth B10 MeVoEn42 GeV

Low-level g-ray line emission develops

Production of escaping particles (SEP)?

Bursts of > 10 MeV photons sometimes occur New radio sources appear as new energetic emissions appear

Type III/V radio burst

Type II and IV radio bursts

a

From Reference 11.

chamber, so high-energy emissions were recorded throughout the flare. After the intense phase of this flare, analysis of the COMPTEL data indicated a flux of neutrons in the energy range of 40–60 MeV, and the EGRET TASC also showed possible evidence of even higher energy neutrons. Figure 15, from Dunphy et al. (61), shows the energy loss spectra in the EGRET TASC for several time intervals during this long event. Phases I-1 and II correspond to the impulsive and extended phases listed in Table 6. Two important questions posed by the long-term high-energy gamma-ray emission concern the acceleration mechanism/geometry and whether the particles are accelerated in a short impulsive phase that lasts a few minutes, are trapped in magnetic loops, and subsequently precipitate into the lower corona. Or are the particles continually accelerated by processes other than that which produced the particles causing the initial burst of high-energy radiation? The first possibility was considered by Mandzhavidze and Ramaty (62), who

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Table 7. Some Extreme Particle Propertiesa Parameter Number

Electrons

Ions (protons)

1041 ð> 20 keVÞ 1036 ð> 100 keVÞ 51034 ð> 300 keVÞ

Rise time, s

10  2

Duration, s Total energy, ergs

10 ! 1034 ð> 20 keVÞ 1029 ð> 100 keVÞ 1028 ð> 300 keVÞ

31035 ð> 30 MeVÞ

1032 ð> 300 MeVÞ >1 60 !

1030 ð> 30 MeVÞ 31028 ð> 300 MeVÞ Power, ergs s  1 1032 ð> 20 keVÞ 21028 ð> 300 MeV) a

From Reference 11.

concluded that the trapping model can explain the observations. The second possibility has been considered in interpreting the GAMMA 1 and COMPTEL data for the 15 June 1991 flare. Electric fields produced in a reconnecting current sheet (behind a rising CME) accelerate the ions and electrons and explain, particularly, the delayed microwave emission, the 1- to 8-MeV and 100-MeV, and 1-GeV gamma rays. An earlier analysis of essentially the same data concluded that the delayed (gradual phase) emission resulted from prolonged production due to stochastic acceleration. Further interpretations of flares from AR6659 and the question of storage or continuous acceleration of charged particles have recently been reviewed (63). The high-energy flares discussed before were observed by using nonimaging instruments, so the actual location of gamma-ray and neutron source regions is unknown. However, imaging of relatively small regions on the Sun is possible at photon energies below 100 keV. Due to the launch of the Japanese solar observing satellite YOHKOH and its soft X-ray telescope (SXT) and hard X-ray telescope (HXT), flare observations have been revolutionized. On 13 January 1992, these two telescopes took images of a compact M2.0-class X-ray flare on the west limb. Figure 16 shows a schematic of the flare geometry as deduced from the SXT and HXT images given in Masuda et al. (64). The SXR loop represents the SXT image at B2 keV, and the three features marked HXR correspond to the major emissions recorded by the HXT above B25 keV. The flare in X rays below B33 keV was of 1-minute duration or less. The image in hard X rays shows the loop top source B700 km above the top of the SXR loop, and it is believed that the former is the site of acceleration of the electrons that stream down the two sides of the SXR loop and produce the HXR foot-point sources. The hypothetical magnetic field geometry above the loop top is based on observations of two-ribbon flares

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10000.00 1000.00

Phase I-1

100.00 10.00 1.00 0.10 0.01 10000.00 1000.00

Phase I-2

100.00 10.00 1.00

Counts/MeV-s

0.10 0.01 10000.00 1000.00

Interphase

100.00 10.00 1.00 0.10 0.01 10000.00 1000.00

Phase II

100.00 10.00 Neutrons

1.00

π-Decay gammas 0.10 0.01 1

10 Energy, MeV

100

Figure 15. Plots of energy-loss spectra (background subtracted) observed by the TASC for several time intervals during the 11 June flare. The spectra were fit with a multicomponent model spectrum described in the text. From Dunphy et al. (61).

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Reconnection point

Reconnection flow (v ~ 3,000 km s−1)

Shock HXR loop-top impulsive source

Energetic electrons

SXR loop

Evaporation

HXR double footpoint sources

Figure 16. The magnetic-field geometry for reconnection derived from the present observation. The important features are elongated antiparallel magnetic fields above an arcade of closed loops; a current sheet (or neutral sheet) formation between them and the reconnection point impinges on the underlying closed loop and forms a shock, resulting in a high-temperature (TeB2  108 K) region just above the closed loop. It is also likely that electrons are accelerated in the shock and stream down along the reconnected field toward the double foot-point sources. From Masuda et al. (64).

(65) and is consistent with the view that energy release in a coronal current sheet due to magnetic reconnection produces downward (and outward) moving shocks, which can both heat material in the loop top to temperatures as high as 2  108 K and accelerate nonthermal electrons. Figure 17 shows an SXT image of a major X9 ‘‘behind the limb’’ flare that has major emission above the limb in the corona. This observation has stimulated further theoretical and interpretive research to understand this type of flare, which may be fundamental for all flares. Recently, Forbes (66) studied possible trigger mechanisms for both solar and stellar flares and concluded that it is likely that both ‘‘involve the sudden release of magnetic energy’’ via reconnection but there may be different trigger mechanisms involved simply because the magnetic field structures are probably not the same. In flares, the ‘‘trigger’’ causes a loss of stability, or equilibrium, in a

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Figure 17. SXT images of the X9 event of 2 November 1992 as the sum of the raw white light and soft X-ray images at about 03:18 UT. The contours show the location of the limb. The flare occurred about 101 behind the limb and was the largest of the Yohkoh eta. By Hugh Hudson. This figure is available in full color at http://www.mrw.interscience. wiley.com/esst.

coronal magnetic field configuration and could be due to new magnetic flux that emerges from the photosphere or twisting of field lines tied at foot points. Either would produce stresses in the field that increase to a limit above which equilibrium is lost. Coronal Mass Ejections. During satellite missions in the early 1970s, particularly during OSO 7 and the SKYLAB observations, bright transient features that moved rapidly away from the Sun were seen in the corona. These events, which are called coronal mass ejections (CMEs), were observed using white light coronagraphs, which occult the bright disc emission, making an artificial eclipse. They are made visible by scattering, toward the observer, of the photospheric continuum by free electrons (Thomson scattering) from the leading edge of the feature. This technique permits observing the corona from about 2–30 solar radii (R) from the center of the Sun. CMEs have speeds, measured by the projected position of the leading edge versus time, which range from o50 to B2000 km/s, and estimates of the total mass involved are B1012 to 1014 kg. CMEs are actual expulsions of matter and magnetic fields from regions lower in the solar atmosphere. Such features are sometimes called coronal transients (see (27)). The most recent studies of CMEs have found that they are more frequently associated with eruptive prominences than solar flares. The sudden disappearance (or disparition brusque) of filaments, it is believed, is associated with most

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Figure 18. This EIT full Sun image, taken on 14 September 1999 in the He II emission ˚ shows the upper chromosphere/lower transitional region at a temperature of line at 304 A about 60,000 K. The bright features are called active regions. A huge erupting prominence escaping the Sun can be seen in the upper right part of the image. Prominences are ‘‘cool’’ 60,000 K plasma embedded in the much hotter surrounding corona, which is typically at temperatures above 1 million K. If an eruption like this is directed toward Earth, it can cause a significant amount of geomagnetic activity in Earth’s environment and a following spectacular aurora. Instrument: EIT; Taken; 14 Sept 1999, 07:19UT (courtesy of SOHO/ EIT.SOHO is a project of international cooperation between ESA and NASA). This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

CMEs (30). In Figure 18 is shown a recent SOHO image, made by its ExtremeUltraviolet Imaging Telescope (EIT), of an eruptive prominence on the west solar limb. Presumably the CME that would be associated with such an eruption would lift off from the Sun because the matter is no longer tied to the Sun by the magnetic field of the preexisting prominence. The cause of such eruptions is not known at the present time (30). Studies have also shown that flares are not always associated with CMEs and sometimes the flare initiation lags the projected time of the CME onset. The CMEs that have speeds in excess of B600 km/s are potential drivers of shock waves moving through the interplanetary magnetic field, and it is believed that acceleration of energetic particles to GeV energies can occur at the shock front. One proposal is that all of the long duration, or energetic gradual, particle events seen at Earth that produce GLEs are caused by particles accelerated by CME shocks. This is actually a very controversial subject because all of the major GLEs at Earth are also associated with major solar flares. As discussed earlier, it is known that GeV ions and electrons are accelerated in a solar flare. Major advances in studying CMEs are currently in progress using the Large Angle

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Spectroscopic Coronagraph (LASCO) on the SOHO spacecraft. Further discussion of CMEs may be found in Kahler (67). Coronae of Solar Type Stars. It might be surprising to find that activity on some solar-type distant stars, B42 light years away compared to the Sun’s distance of 8 light-minutes, should be observable from near Earth. The photon flux in any wavelength band would be reduced by a factor of B1014 from such a star, compared to the Sun. So-called flare stars have been observed by astronomers (professional and amateur) at visible wavelengths since stellar observations have become commonplace. An important example of flare stars is the binary group known as RS Canum Venaticorum (68), whose discovery is closely related to the discovery of starspots (68). The appellation, ‘‘flare stars,’’ was given to this type of star because transient emission was similar to the optical intensity variation (in Ha) seen in solar flares. Flare stars have also been seen at radio wavelengths since the late 1950s and also at ultraviolet and X-ray wavelengths (see later and also References 69 and 70). In 1974, a rocket observation was made of Capella, an RS CVn variable, by Catura and Acton at soft X-ray wavelengths. This object was identified as a member of ‘‘a new class of galactic X-ray sources.’’ It is understood now that the X-ray emission is most likely associated with the degree of magnetic activity on the star, which again is related to the size of starspots on the star. Since the Einstein Satellite’s study of X-ray stars (71) and more recently, using ROSAT data, astronomers have studied the X-ray luminosity of stars ranging in size from giant stars to main sequence and subgiants. Because it is believed that the magnetic activity of the Sun, represented by the number (or total area) of sunspots on the Sun (see Fig. 8), is due to the action of the solar dynamo that results from the differential rotation of the sun, it is believed that the intensity of stellar X-ray emission could be related to the observed rotational speed of the star. This correlation of the X-ray luminosity of a variety of stars with known rotational speeds is shown in Fig. 19 from Pallavicini et al. (71) and has been confirmed in a recent study by Haisch and Schmitt (72). This result also can be understood as follows: recall in our discussion of the origin of sunspots, that two factors were essential: differential rotation of the Sun versus latitude and the presence of a convective zone that provides the conditions for the solar dynamo and leads to the creation of sunspots, emerging magnetic flux, and solar activity in general. Now, the RS CVn binary variables are the most intense emitters of X rays, which it is thought, is related to the short orbital period of the pair about their mutual center of mass. Because the separation of the pair is relatively small, o0.3 AU, there are strong tidal forces, and the individual star’s rotational periods become synchronized with the rapid orbital motion of the pair. It is believed that this is the basic cause of the greater magnetic activity and hence greater X-ray intensity of this system, as shown in Fig. 19. Haisch and Schmitt (72) also discuss the recent 1995 observation of a huge X-ray flare on the binary T Tauri star, V773 Tau. The peak X-ray (0.7–10 keV) luminosity was 1033 ergs/s, and the total energy release in the same energy band was 1037 ergs; these values are 104 and 105 higher than the same values for solar flares. T Tauri binary star systems also rotate rapidly, not from tidal action, but because they are young stars that contract from the protostellar cloud stage. It has also been inferred that large sunspots exist in this system, so these observations give

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32

Log Lx, erg s−1

31

RS CVn's

30 29

28 IV + V III + II Empty circles: Sp GO-M5 Filled circles: Sp F7-F8

27 Sun 26 0

1

2

Log V sin i, km s−1

Figure 19. X-ray luminosities plotted versus projected rotational velocities for stars of spectral type F7-F8 (filled symbols) and G0 to M5 (open symbols). The position of RS CVn stars in indicated. From Reference 71.

strong-support to the belief that rotation is the prime cause of their magnetic activity, that is, flaring and this idea carried to the Sun supports the theoretical model of sunspot creation as a result of differential rotation twisting a poloidal magnetic field See also Haisch 1991 (73). Another important aspect of this solar/stellar connection is the necessity of convective zones for producing a dynamo. The association of coronal activity in the Sun with its magnetic activity, probably produced at the base of the convective zone, is shown in Fig. 20. Helioseismology. In 1960, Leighton, using the 150-foot tower telescope at Mount Wilson solar observatory to study the Doppler-shift motions of Fraunhofer lines over granulations, discovered that the radial velocities of certain lines oscillated in a period of 5 minutes. These were later explained, independently by R. Ulrich and R. Stein, and another solar physicist, J. Leibacher, by the superposition of nonradial sound waves trapped in the convective zone, which acted as a resonant cavity. Many other modes of oscillation, called acoustic modes, have subsequently been discovered. Because the Sun is a sphere of hot gas, it has oscillatory amplitudes described by spherical harmonics, which are designated by eigenvalues l (angular degree) and m (azimuthal). The 5-minute oscillations now have been identified with l-values from 0 to 41000. The modes that have small l-values (long wavelengths) can penetrate to the center of the Sun. The waves are also designated as p-modes, or pressure modes, and g-modes, or gravity modes; the latter are due to buoyancy. The g-modes have larger amplitudes and are better for investigating the Sun’s structure near the core; however, there is not yet definite evidence of their existence. Clearly, because disturbances are

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Figure 20. The violently hot solar corona is visible in X radiation. These images from the Yohkoh satellete show how the corona varies with the 11-year solar activity cycle. The bright corona at the left shows high activity, which may be associated with magnetic storms on Earth and other injurious effects. The dark corona on the right shows the present situation (ca. 1995) of low magnetic activity. The plot at the bottom shows the time variation. Activity started to increase again in 1997–1998. The solar X-ray images are from the Yohkoh mission of ISAS, Japan. The X-ray telescope was prepared by the Lockheed Palo Alto Research Laboratory, the National Astronomical Observatory of Japan, and the University of Tokyo with the support of NASA and ISAS. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

propagated at the velocity of sound, which depends on temperature and density, the study of solar oscillations can give information about these parameters in the solar interior. Currently, the two major projects to study solar oscillations are GONG and the several instruments on the SOHO spacecraft, which were described earlier. The Global Oscillation Network Group (GONG) began operation on 5 October 1995 and consists of six solar observatories, fitted with special telescopes, which are spaced around Earth so at least one will have the Sun in its field of view. SOHO was launched on 2 December 1995, reached the inner Lagrangian point L1 on 14 February 1996, and began constant observations of the Sun. On 25 June 1998, contact with SOHO was lost, but was established again several months later.

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The results of new observations from these projects have not been fully synthesized, but one important result gives the sound speed and density profiles throughout the Sun. At a COSPAR meeting in Nagoya in 1998, Shibahashi (74) listed several vital issues in solar physics that helioseismology research will impact, and we mention a few: *

*

*

*

*

Attempts to solve the solar neutrino problem will benefit by improving the SSM. [See note added in proof.] Measuring the rotational rate of the core can aid in determining the evolution of the Sun because of conservation of angular momentum, which suggests an increasing angular velocity with age. Conversely, the surface layers are expected to lose angular momentum to the SW, so the Sun’s internal angular momentum distribution should be determined. Related to this is the differential rotational rate of the solar surface and convective zone as a function of latitude. The mechanism for depletion of surface lithium abundance is unknown as is the physical cause of the existence of the chromosphere and corona. The cause of solar activity and the energy release mechanisms and triggers for flares and CMEs is unknown. How can the change in solar luminosity with solar activity, which is connected to energy transport in the convective zone, be understood? Finally, helioseismologists expect to develop a model of the solar interior, independent of the evolutionary SSM.

For further reading on this fascinating subject, see References 27,29,30,74. Of most value are the web home pages: *

*

GONG: http://helios.tuc.noao.edu/helioseismology.html SOHO: http://sohowww.nascom.nasa.gov/

Space Weather. The effects of solar activity are dramatic in space and even damage or disable satellites. Sometimes the effects on Earth can affect everyday life by disrupting communications or causing electrical power outages. Major effects are caused by intense fluxes of ionizing radiation (primarily energetic protons) and intense X-ray bursts. The most damaging are the large solar energetic proton (SEP) events. The deleterious effects of solar activity have therefore caused the development of a new business—space weather forecasting. The organization charged with this task is the National Oceanographic and Atmospheric Administration (NOAA) in Boulder, Colorado, whose Space Environment Center (SEC) issues daily reports and sends alerts of impending space storms to all who may be affected. Table 8 lists the known ‘‘Solar Proton Events Affecting the Earth Environment’’ which is available on the SEC web page. Following is the SEC web address, where the reader can browse to find topics of particular interest: http:// www.sel.noaa.gov/.

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Notes Added in Proof 1. Firm Evidence for Neutrino Oscillations. This possible solution to the solar neutrino problem was discussed in the original manuscript. Now a new neutrino detector, which began observing solar neutrinos in late 1999, provides compelling evidence that the neutrinos oscillate or change their ‘‘flavor’’ in transit from the solar core to the earth. (See Ahmed et al., Phys. Rev. Lett. 89, 011301 (2002) and Physics Today, July 2000). The detector, The Sudbury Neutrino Observatory (SNO), consists of a kiloton of heavy water in a transparent spherical shell of 12 m diameter and viewed by 9456 photomultiplier tubes. This apparatus is surrounded by a large ultrapure light water shield and resides in a nickel mine at a depth of 6010 m of water equivalent near Sudbury in Ontario, Canada. The SNO can detect the continuous spectrum of 8B neutrinos above 2.2 MeV (see text—Fig. 4) through the reactions 1: ue þ d ! p þ p þ e 2: ux þ d ! p þ n þ ux and also by elastic scattering via 3: ux þ e ! ux þ e Here x stands for all neutrino flavors. x ¼ e, m, t. Reactions (2) and (3) are sensitive to all neutrino flavors while 1 is only sensitive to the electron-type neutrinos which are directly produced in the fusion reactions in the core of the sun (see text—Table 3). The essential results are flux of 8B electron neutrinos, (1.7670.11)  106 cm  2 s  1 total flux of all neutrinos, (5.0970.062)  106 cm  2 s  1 The SSM prediction for the 8B flux is (5.0570.91)  106 cm  2 s  1, in excellent agreement with the measured total flux of all neutrinos. Thus about twothirds of the Sun’s 8B neutrinos have changed their flavor, i.e. presumably transmutated into muon (m) or tau (t) neutrinos. 2. RHESSI – A New Satellite to Study Solar Flares. On 5 February 2002, the Reuven Ramaty High Energy Solar Spectroscopic Imager (RHESSI) was launched by Pegasus XL spacecraft into a low-altitude orbit at an inclination of B381. The RHESSI mission consists of a single spin-stabilized spacecraft. The only instrument on board is an imaging spectrometer with the ability to obtain high-fidelity color movies of solar flares in X-rays and gamma rays. It uses two new complementary technologies: fine grids to modulate the solar radiation and germanium detectors to measure the energy of each photon very precisely. RHESSI’s imaging capability is achieved with fine tungsten and/or molybdenum grids that modulate the solar X-ray flux as the spacecraft rotates at

750

92 58 21 42 12 13 13 18 29

28 3,500 2,000 53 450 27 110 68 15 18 17 17 54 9,200 44

1988 Jan 03/0835 Mar 25/2330 Jun 30/1140 Aug 26/0045 Oct 12/0930 Nov 09/0635 Nov 14/0235 Dec 17/0855 Dec 18/0150

1989 Jan 05/0130 Mar 13/0645 Mar 18/0920 Mar 24/0110 Apr 12/0125 May 05/1000 May 06/1045 May 23/1350 May 24/0905 Jun 18/1910 Jun 30/0710 Jul 01/0720 Jul 25/1225 Aug 13/0710 Sep 04/0510

Jan 02/2325 Mar 25/2225 Jun 30/1055 Aug 26/0000 Oct 12/0920 Nov 08/2225 Nov 14/0130 Dec 17/0610 Dec 17/2000

Jan 04/2305 Mar 08/1735 Mar 17/1855 Mar 23/2040 Apr 11/1435 May 05/0905 May 06/0235 May 23/1135 May 24/0730 Jun 18/1650 Jun 30/0655 Jul 01/0655 Jul 25/0900 Aug 12/1600 Sep 04/0120

Proton flux, pfu at 410 MeV

Maximum

Start, day/UT

Particle event

78 10 4,600 390 14 2,700

17 44

10,000

1993 Mar 04/1735 Mar 13/0155 1994 Feb 21/0900

Mar 04/1505 Mar 12/2010

Feb 20/0300

350 22 3,000 1,400 110 2,300 30 14 240 12 40 94

Proton flux, pfu at 410 MeV

1992 Feb 07/1115 Mar 16/0840 May 09/2100 Jun 26/0610 Aug 06/1210 Oct 31/0710

May 13/0910 Jun 01/0445 Jun 11/1420 Jun 15/1950 Jul 02/1010 Jul 08/1645 Jul 11/0450 Jul 12/0205 Aug 27/1830 Oct 01/1810 Oct 28/1440 Oct 30/0810

Maximum

Feb 07/0645 Mar 16/0840 May 09/1005 Jun 25/2045 Aug 06/1145 Oct 30/1920

May 13/0300 May 31/1225 Jun 04/0820 Jun 14/2340 Jun 30/0755 Jul 07/0455 Jul 11/0240 Jul 11/2255 Aug 26/1740 Oct 01/1740 Oct 28/1300 Oct 30/0745

Start, day/UT

Particle event

Table 8. Solar Proton Events Affecting Earth Environment January 1988–18 February 2000

Associated CME

751

57 4,500 22 40,000 43 71 380 7,300

950 16 18 13 12 150 410 180 45 79 21 230

240 13 43,000 20 52

Sep 13/0825 Sep 30/0210 Oct 06/0825 Oct 20/1600 Nov 09/0610 Nov 15/0910 Nov 28/1105 Dec 01/1340

1990 Mar 19/2315 Mar 29/1005 Apr 08/1330 Apr 11/2130 Apr 17/0655 Apr 28/1735 May 22/0750 May 25/0115 May 29/0100 Jun 12/1700 Jul 26/2315 Aug 01/2015

1991 Jan 31/1620 Feb 25/1305 Mar 24/0350 Mar 30/0330 Apr 04/1000

Sep 12/1935 Sep 29/1205 Oct 06/0050 Oct 19/1305 Nov 09/0240 Nov 15/0735 Nov 27/2000 Nov 30/1345

Mar 19/0705 Mar 29/0915 Apr 07/2240 Apr 11/2120 Apr 17/0500 Apr 28/1005 May 21/2355 May 24/2125 May 28/0715 Jun 12/1140 Jul 26/1720 Aug 01/0005

Jan 31/1130 Feb 25/1210 Mar 23/0820 Mar 29/2120 Apr 03/0815

1,700 150 210 670 44 1,200 11 310

14 32 14 48 64

13

1998 Apr 21/1205 May 02/1650 May 06/0945 Aug 26/1055 Sep 25/0130 Oct 01/0025 Nov 08/0300 Nov 14/1240 1999 Jan 23/1135 Apr 25/0055 May 05/1955 Jun 02/1010 Jun 04/1055 2000 Feb 18/1215

Jan 23/1105 Apr 24/1804 May 05/1820 Jun 02/0245 Jun 04/0925

Feb 18/1130

72 490

Apr 20/1400 May 02/1420 May 06/0845 Aug 24/2355 Sep 25/0010 Sep 30/1520 Nov 08/0245 Nov 14/0810

1997 Nov 04/1120 Nov 07/0255

63

1995 Oct 20/1210 1996

Oct 20/0825

Nov 04/0830 Nov 06/1305

35

Oct 20/0340

Oct 20/0030

W/ 18 0954

NA Halo/24 1331 Halo/ 03 0606 Halo/ 01o1937 NW/ 04 0726

W/20 1007 Halo/02 1406 W/06 0829 NA NA NA ? NA

W/04 0610 W/06 41300

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 15 rpm. Up to 20 detailed images can be obtained per second. This is sufficient to track electrons as they travel from their acceleration sites, believed to be in the solar corona, and slow down on their way to the lower solar atmosphere. High-resolution spectroscopy is achieved with nine cooled germanium cystals that detect the X-ray and gamma-ray photons transmitted through the grids over the broad energy range of 3 keV to B17 MeV. Their fine energy resolution of about 1 keV is more than sufficient to reveal the detailed features of the X-ray and gamma-ray spectra, clues to the nature of the electron and ion acceleration processes. The spinning spacecraft pointing at or near the Sun center provides a simple and reliable way to achieve the rotation required for the HESSI imaging technique. The RHESSI mission has already observed several flares, so considerable new knowledge about solar flares should be obtained in its nominal 2–3 year lifetime. For the current status of the mission and information about the observed flare events, see the RHESSI homepages at http://hessi.ssl.berkeley.edu and http://hesperia.gsfc.nasa.gov/hessi/

ACKNOWLEDGMENTS

The author would like to acknowledge Terry G. Forbes, who reviewed an early draft of this article and an anonymous reviewer for valuable suggestions. Discussions on relevant high energy solar flare results with Philip Dunphy are greatly appreciated. Jane Fithian expertly prepared computer files for all color and black and white figures with captions. Rosemary Raynes, Katie Makem and Erica Brown formatted the final manuscript, assisted in numerous tasks and verified references, respectively.

BIBLIOGRAPHY 1. 2. 3. 4. 5. 6. 7. 8.

Chaisson, E., and S. McMillan. Astronomy Today. Prentice Hall International, London, 1999. Abell, G.O., D. Morrison, and S.C. Wolff. The Sun—an ordinary star. In Exploration of the Universe, 5th ed., Saunders, Philadelphia, 1987. Kippenhahn, R. 100 Billion Suns: The Birth, Life, and Death of the Stars, translated by J. Steinberg. Basic Books, New York, 1983. Menzel, D.H. Our Sun, rev. ed. Harvard University Press, Cambridge, 1959. Kiepenheuer, K.O. The Sun, translated by A.J. Pomerans. University of Michigan Press, Ann Arbor, 1959. Meadows, A.J. Early Solar Physics. Pergamon, Oxford, 1970. Hufbauer, K. Exploring the Sun-Solar Science Since Galileo. The Johns Hopkins University Press, Baltimore, 1991. Carruthers, G.R. Sounding rocket experiments, astronomical. In The Astronomy and Astrophysics Encyclopedia, S.P. Maran and C. Sagan (eds). Van Nostrand Reinhold, New York, 1992, p. 650.

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10. 11. 12. 13. 14a. 14b. 15. 16. 17. 18. 19. 20. 21.

22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35.

36. 37.

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Brueckner, G.E. Solar physics, space missions. In The Astronomy and Astrophysics Encyclopedia, S.P. Maran and C. Sagan (eds). Van Nostrand Reinhold, New York, 1992, p. 646. Kundu, M., and B. Woodgate. (eds). Energetic Phenomena on the Sun. NASA-(STIB)Conf. Pub. 2439, Washington, DC, 1986. Chupp, E.L. Evolution of our understanding of solar flare particle acceleration: (1942– 1995). In High Energy Solar Physics, Woodbury, NY, AIP Conf. Proc 374: 3, 1996. Lord, D.R. Space Lab. NASA (STID), Washington, DC 1987, p. 374. Keppler, E. Ulysses finishes its first revolution around the Sun. Naturwissenschafter 85: 467 (1998). Bahcall, J.N., and R.K. Ulrich. Ap. J. 170: 593 (1971). Bahcall, J., and H. Pinsonneault. Standard solar models with and without helium diffusion, and the solar neutrino problem. Rev. Mod. Phys. 64: 885 (1992). Lang, K.R. Sun, Earth and Sky. Springer, Berlin, 1995. Chandrasekhar, S. An Introduction to the Study of Stellar Structure. Dover, New York, 1939, p. 453. Bethe, H.A. Energy production in stars. Phys. Rev. 55: 434 (1939). Davis, R., Jr. Attempt to detect the antineutrinos from 2 nuclear reactors by the Cl37 (n,e  )A37 reaction. Phys. Rev. 97: 766 (1955). Bahcall, J., and R. Davis. An account of the development of the solar neutrino problem. In Essays in Nuclear Physics. Cambridge University Press, Cambridge, 1982, p. 243. Bahcall, J.N., and G. Shaviv. Ap. J. 153: 113 (1968). Bahcall, J.N. Solar neutrinos: Solved and unsolved Problems. in J.N. Bahcall, and J.P. Ostriker (eds). Unsolved Problems in Astrophysics. Princeton University Press, Princeton, NJ, 1997, p. 195. Davis, R., Jr., D.S. Harmer, and K.C. Hoffman. Search for neutrinos from the Sun. Phys. Rev. Lett. 20: 1205 (1968). Davis, R. A review of the Homestake solar neutrino experiment. In Progress in Particle and Nuclear Physics, 32: Elsevier Science, Oxford, 1994, p. 13. Bahcall, J.N. Neutrino Astrophysics. Cambridge University Press, Cambridge, England, New York, 1989. Bahcall, J., and H.A. Bethe. A solution of the solar-neutrino problem. Phys. Rev. Lett. 65: 2233 (1990). Eddington, A.S. The Internal Constitution of Stars. Dover, New York, 1926, p. 108. Foukal, P. Solar Astrophysics. Wiley, New York, 1990. Engvold, O. The solar chemical composition. Physica Scripta 16: 48 (1977). Noyes, R.W. The Sun—Our Star. Harvard University Press, Cambridge, 1982. Zirin, H., Astrophysics of the Sun. Cambridge University Press, Cambridge, Cambridgeshire, New York, 1988. Babcock, H. The topology of the Sun’s magnetic field and the 22-year cycle. Ap. J. 133: 572 (1961). Gough, D. New from the solar interior. Science 287: 2434 (2000). Howe, R., J. Christensen-Dalsgaard, F. Hill, et. al. Dynamic variations at the base of the solar convection zone. Science 287: 2456 (2000). Lang, K.R. SOHO reveals the secrets of the Sun. Sci. Am. 276: 40 (1997). Vernazza, J.E., E.H. Avrett, and R. Loeser. Structure of the solar chromosphere. III. Models of the EUV brightness components of the quiet Sun. Ap. J. Suppl. 45: 635 (1981). Biermann, L.F. Ko¨metenschweife und solare korpuskular-strahlung. Z. Ap. 29: 274 (1950). Parker, E.N. Dynamics of the interplanetary gas and magnetic fields. Ap. J. 128: 664 (1958).

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38. Parker, E.N. Interplanetary Dynamical Processes. Interscience, New York, 1963. 39. Eddy, J.A. A new Sun – The solar results from SkyLab. NASA(STIB), Washington, DC, 1979, p. 98. 40. Bartels, J. Terestrial-magnetic activity and its relations to solar phenomena. Terrestrial Magn. Atmos. Elec. 37: 1 (1932). 41. Withbroe, G.L. Origins of the solar wind in the corona. In The Sun and Heliosphere in Three Dimensions, R.G. Marsden (ed.). Reidel, Dordrecht, 1986, p. 19. 42. Gloeckler, G.J. Geiss, H. Balsgier, et al. The solar wind ion composition spectrometer. Astron. Astrophys. Suppl. Ser. 92: 267 (1992). 43. Clerke, A.M. A Popular History of Astronomy during the 19th Century. Adam and Charles Black, Edinburgh, 1885, p. 205. 44. Pallavicini, R. The role of magnetic loops in solar flares. Philos. Trans. R. Soc. A 336: 389 (1991). 45. Pneuman, G.W. Two ribbon flares: Post (flare) loops. In Solar Flare Magnetohydrodynamics, E.R. Priest (ed.). Gordon and Breach, New York, 1981, Vol. I. p. 379. 46. Golub, L., M. Herant, K. Kalata, et al. Subarcsecond observations of the solar X-ray corona. Nature 344: 842 (1990). 47. Svestka, Z. Solar Flares. Reidel, Dordrecht, Boston, 1976. 48. Tandberg-Hanssen, E., and A.G. Emslie. The Physics of Solar Flares. Cambridge University Press, New York, 1988. 49. Somov, B.V. Physical Processes in Solar Flares. Kluwer, Dordrecht, 1992. 50. Hess, W.N. (ed.). AAS-NASA Symp. Phys. Solar Flares. NASA (STID), Washington, DC, 1964. 51. Simpson, J.A. Evidence for a solar cosmic-ray component. in The Sun, G.P. Kuiper (ed.). University of Chicago, Chicago, 1953, p. 715. 52. Morrison, P. On gamma-ray astronomy. Nuovo Cimento 7: 858 (1958). 53. Chupp, E.L., D.J. Forrest, P.R. Higbie, et al. Solar gamma ray lines observed during the solar activity of August 2 to August 11, 1992. Nature 241: 333 (1973). 54. Forrest, D.J., and E.L. Chupp, et al. Simultaneous acceleration of electrons and ions in solar flares. Nature 305: 291 (1993). 55. Biermann, L., O. Haxel, and A. Schluter. Neutrale ultrastrahlung von der Sonne. Naturforsch. 6a: 47 (1951). 56. Forrest, D.J., W.T. Vestrand, E.L. Chupp, et al. Neutral pion production in solar flares. Proc. 19th Int. Cosmic Ray Conf. 4: 146 (1985). 57. Chupp, E.L., H. Debrunner, E. Fluckiger, et al. Solar neutron emissivity during the large flare on 1982 June 3. Ap. J. 318: 913 (1987). 58. Ramaty, R. Flare physics at high energies. In Astrophysics from the Moon, M.J. Mumma and H.J. Smith (eds). AIP, New York, 1990, p. 122. 59. Lingenfelter, and Ramaty. High-energy nuclear reactions and solar flares. In High Energy Nuclear Reactions in Astrophysics. B.S.P. Shen (ed.). Benjamin, New York, 1967, p. 99. 60. Kanbach, G.O., D.L. Bertsch, C.E. Fichtel, et al. Detection of a long-duration solar gamma-ray flare on June 11, 1991 with EGRET on COMPTON-GRO. A. A. Suppl. 97: 349 (1993). 61. Dunphy, P.P., E.L. Chupp, D.L. Bertsch, et al. Gamma rays and neutrons as a probe of flare protron spectra: The solar flare of 11 June 1991. Sol. Phys. 187: 45 (1999). 62. Mandzhavidze, N., and R. Ramaty. Gamma rays from pion decay: Evidence for longterm trapping of particles in solar flares. Ap. J. Lett. 396: L111 (1992). 63. Hudson, H., and S.M. Ryan. High-energy particles in solar flares. Ann. Rev. Astron. Astrophys. 33: 239 (1995). 64. Masuda, S., T. Kusugi, H. Hara, et al. A loop-top hard X-ray source in a compact solar-flare as evidence for magnetic reconnection. Nature 371: 495 (1994).

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65. Haisch, B.M, K.T. Strong, and M.A. Rodono. Flares on the Sun and other stars. Ann. Rev. Astron. Astrophys. 29: 275 (1991a). 66. Forbes, T. Solar and Stellar Flares. Philos. Trans. R. Soc. A 258: 711 (2000). 67. Kahler, S.W. Solar Activity, Coronal Mass Ejections. In The Astronomy and Astrophysics Encyclopedia. Van Nostrand, New York, 1992, p. 631. 68. Hall, D.S. Binary Stars, RS Canum Venaticorum Type. In The Astronomy and Astrophysics Encyclopedia. Van Nostrand, New York, 1992, p. 74. 69. Hjellming, R.M. Stars, Radio Emission. In The Astronomy and Astrophysics Encyclopedia. Van Nostrand, New York, 1992, p. 780. 70. Byrne, P.B. Stars, Red Dwarfs and Flare Stars. In The Astronomy and Astrophysics Encyclopedia. Van Nostrand, New York, 1992, p. 783. 71. Pallavicini, R., L. Golub, R. Rosner, et al. Relations among stellar X-ray emission observed from Einstein, stellar rotation and bolometric luminosity. Ap. J. 248: 279 (1981). 72. Haisch, B., and J. Schmitt. The solar–stellar connection. In Sky and Telescope. Sky, Cambridge, 1999, p. 46. 73. Haisch, B.M. Solar–stellar connection. In The Many Faces of the Sun—A Summary of the Results from NASA’s Solar Maximum Mission, K.T. Strong, J.L.R. Saba, B.M. Haisch, and J.T. Schmelz (eds). Springer, New York, 1991, p. 481. 74. Shibahashi, H. What the helioseismic results mean for the solar interior. In Helioseismology and Solar Variability—Advances in Space Research, C. Frohlich and B.H. Foing (eds). Adv. Space Res. 24 (2): 137 (1999).

EDWARD L. CHUPP University of New Hampshire Durham, New Hampshire

U U.S. MANNED SPACE FLIGHT: MERCURY TO THE SHUTTLE Introduction There are many publications relating to Mercury, Gemini, Apollo, Skylab, ApolloSoyuz Test Project, and the Space Shuttle. This paper presents an objective review of some of the technical considerations of the design, development, and operation of these spacecraft. Although there were many triumphs, a number of surprises, and not a few setbacks during the 39-year period covered, such incidents will be mentioned only in the context of technical significance. A little more than 19 years elapsed between 20 February 1962, when John Glenn first rode Mercury into orbit, and 12 April 1981, when John Young and Robert Crippen inaugurated orbital flight for the Space Transportation System, or the Space Shuttle as it is commonly called. A comparison of the Mercury capsule and the spaceship Columbia reviews the extent of progress made during these first years of manned flight. During the early 1920s after approximately 19 years of development, the aviation industry was also entering the transportation phase. But there is a significant difference between the development of airplanes and the development of manned spacecraft. Early airplanes were relatively cheap and easy to build. Often, only one person and at most just a handful would be sufficient to do the entire design job. Consequently, a great number of airplanes was built and flown. This led to a rapid evolutionary process for both the design and the operation of aircraft, and a number of accidents and fatalities provided a strong cooperative influence on any misdirection. The philosophical basis for the design and operation of spacecraft has had to progress without this impartial and unerring guidance of ‘‘survival of the fittest.’’ As a substitute, we have had to rely on 756

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intensive analysis, extensive testing, imperfect simulations, and the judgment and experience of the hundreds of key people working on these programs. American manned spacecraft by type together with a number of weights, dimensions, and features that characterize them are listed in Appendix A. Mercury was a first of a kind design, and the Gemini and Apollo command and service module (CSM) followed in an evolutionary trend. The Apollo lunar module (LM) and the Space Shuttle, on the other hand, represent new and distinctive designs and have unique features for which no prior art existed.

Project Mercury The National Aeronautics and Space Administration (NASA) was established in October 1958, and within a few weeks, the first manned spacecraft program, Project Mercury, was initiated (1). To execute the program as quickly as possible, it was deemed desirable to choose an existing rocket for the launch vehicle. The Atlas, considered the most powerful that would be available during the desired time period, was chosen. Conservative estimates indicated that the Atlas could orbit approximately 900 kg (2000 lbm) of payload. The spacecraft grew almost 50% during its development period. When finally developed, the entry weight of the Mercury spacecraft was in excess of 1180 kg (2600 lbm). Fortunately, the Atlas performance also grew sufficiently during this period. The basic purpose of Project Mercury was to expose several test pilots to orbital flight and to have them evaluate the experience so that future and more substantive programs could be planned (Appendix B). To move the program rapidly and because of severe weight constraints, the design of all systems was as simple as possible and provided only the basic necessities of launch, a short orbital flight, and safe descent and landing. A simple ballistic configuration designed to minimize reentry heating was chosen. Although kept as simple as possible, each active system had at least one level of redundancy. Descent was initiated by firing three solid retrorockets; however, a safe reentry was ensured if only two of these rockets fired. A parachute system that was backed up by a reserve system of identical design provided a safe landing. Redundant sets of hydrogen peroxide monopropellant reaction control jets were used. To guard against electrical failure, one set of jets had its electrically actuated valves backed up by a set of mechanical valves directly linked to the astronaut’s rotational hand controller. Life support was provided by a pure-oxygen cabin atmosphere at one-third sea-level pressure. This was backed up by a pressure suit that would automatically inflate if the cabin atmosphere dropped to less than one-fourth sea-level pressure. Redundant voice, telemetry, and command signals could be sent and received on a number of communication channels. The Mercury capsule also carried both C-band and S-band radar beacons to assist the radars at the various ground stations that tracked the flight. Finally, power was supplied by redundant battery sets that had independent buses. Heat was dissipated from the Mercury crew compartment by a water evaporator. Figure 1 shows the Mercury spacecraft and its components. The concept for flight control was very simple. Basically, the Mercury spacecraft was inserted into a low Earth orbit using the launch vehicle guidance

758 Escape rocket

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Retropack

Pitch and Drogue chute yaw jets

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Figure 1. The Mercury Spacecraft and some of its component parts. The escape tower and rocket is shown in picture (a), the heat shield and the retrorocket pack is shown in (b) and the position of the astronaut in the space capsule. These two figures can also be seen at the following websites: http://www.howstuffworks.com/gif/mercury-capsule-drawing.jpg and http://www.howstuffworks.com/gif/mercury-capsule-drawing2.jpg. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

system. Once in orbit, no further velocity change maneuvers were required until it was time for descent. Return to Earth was accomplished by maneuvering the spacecraft to a retrorocket firing attitude where the heat shield was forward and then firing a cluster of three solid rockets strapped to the heat shield. This deflected the flight path to one that entered Earth’s atmosphere. The spacecraft was not designed to produce lift, so it followed a highly predictable ballistic entry trajectory. Consequently, the time when the retrorockets were fired determined the location of splashdown in the ocean. It was recognized that there would be a fairly large dispersion about the planned landing location. However, consideration of emergency descent or aborts during launch that could result in a landing anywhere along the flight track made survival for a period of time on the water after landing and extensive use of location aids a basic design requirement anyway. Flight control equipment on board the spacecraft was needed only for attitude control, particularly when firing the retrorockets. These functions could be done both manually and automatically. When in automatic flight, an autopilot and two horizon scanners were set up to maintain the vehicle in a fixed attitude with respect to local vertical. An onboard timer also initiated the retrofiring sequence after commanding the optimum attitude for the maneuver. This timer, which was started at liftoff, could be corrected during flight by command signals from the ground. The astronaut could also control attitude with a hand controller either by using an attitude indicator on the instrument panel or by looking through the window. He could also override automatic initiation of the retrorocket sequence. The astronaut could crudely determine his position in orbit by comparing his view of Earth using a clock-driven replica of Earth’s globe. The mission was crontrolled from the ground. Communications with the spacecraft and tracking data were obtained from a network of stations along the path of the first orbits. There were 16 different stations located to allow a maximum of 10 minutes without communication contact. The data from these stations were sent to the Mercury Control Center at the launch site in Florida. On

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the basis of these processed tracking data, the orbital ephemeris was determined. The location of intended splashdown was predetermined to accommodate postretrofiring tracking and thereby to enhance final location for recovery. During Project Mercury, six tests were made, two suborbital on a Redstone launch vehicle and four orbital on the Atlas. In preparing for these tests, 17 unmanned flight tests were made; four were made with primates aboard the spacecraft. Only 10 of these 17 tests were successful. In addition to the biomedical studies, these tests provided engineering studies of all of the modes of flight such as escape system operation, entry flight, and performance of the various onboard systems. Of the four manned orbital tests, each was of increasing duration starting with the three orbits made by John Glenn and ending with the 22-orbital flight made by Gordon Cooper. All U.S. manned spaceflights are summarized in Appendix B.

Gemini Program Based on the success of Mercury and undisputed evidence confirmed by the Russian program that people could function comfortably and effectively in space, it became obvious that the next step would be to develop the role of people in space. Consequently, two new programs were established during 1961 to this end: Gemini, to explore the possibilities and limitations of manned operation in space, and Apollo, a bold program committed to exploration of the Moon and sending people there. In Gemini, the approach was, quickly and, where possible, directly, to capitalize on what was learned in the Project Mercury (2). However, vastly improved capability was desired to explore the feasibility of the following operations: maneuvering in orbit, rendezvous and docking with another vehicle, extravehicular activity (EVA) by the astronauts in pressure suits, guided flight during entry to precisely designated target areas, and establishing and refining flight operation methods. All of these objectives were met with highly positive results that were applied extensively in the Apollo Program and contributed greatly to the success of that program. The Gemini Program employed the new Titan II launch vehicle. This booster had roughly twice the capability of the Atlas and made it possible for the Gemini spacecraft to carry a crew of two and to incorporate a great number of design features requisite to the program requirements. The Gemini configuration was basically a scaled-up version of the Mercury vehicle. However, the center of mass of the Gemini spacecraft was offset from its centerline. This configuration caused the spacecraft to trim at a slight angle of attack, which produced a small aerodynamic force (lift) normal to the drag force. Although the lift-to-drag ratio (L/D) was only slightly greater than 0.1, it was more than sufficient to steer the entry path to a desired touchdown point. Steering was done by controlling the roll attitude of the spacecraft and thereby the lift vector to deflect the reentry path (up, down, right, left) in the desired manner. Gemini carried its adapter into orbit. This was the conical structure that connected the Gemini spacecraft to the booster. In this adapter section, which was jettisoned before entry, a great number of systems that were used during orbit were carried. A set of bipropellant rocket motors provided for both

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translational and rotational maneuvers. Fuel cells and their hydrogen and oxygen supply tanks were installed in the adapter as were radiators that provided for heat dissipation. The fuel cells and radiator made mission durations as long as 2 weeks possible. The Gemini spacecraft was equipped with an inertial platform and digital computer to aid in orbital navigation and to provide for reentry navigation and guidance. Rendezvous radar and docking equipment were installed in the nose of the Gemini capsule. The hypergolic propellants used in the Titan II launch vehicle produced a low-yield explosion in the event of a launch vehicle breakup, so ejection seats could be used instead of an escape rocket. Finally, to accommodate extravehicular activity, a hatch operable while in orbit was provided over each astronaut’s position. In contrast to the great number of unmanned flights that preceded manned operations in Project Mercury, there were only two unmanned flights in the Gemini Program. The purpose of these flights was primarily to check out system operations and the compatibility between spacecraft and launch vehicle. The Gemini Program also included a target vehicle to accommodate rendezvous operations. This target vehicle was a converted Agena propulsion stage that was launched into a rendezvous-compatible orbit by an Atlas launch vehicle. The procedure was to launch the target vehicle shortly before the Gemini was launched and then to maneuver the Gemini through a series of rendezvous maneuvers until the Gemini was brought within a few feet of the target vehicle. After inspection, the Gemini was maneuvered into a docking position, and the docking maneuver was made. After docking was completed, the Agena propulsion system could then be controlled by the crew on board the Gemini. The Agena propulsion stage could be used to make several maneuvers while the Gemini was attached, for instance, raising the orbit’s apogee to much higher altitudes than could otherwise be achieved. There were 10 manned flights of Gemini during a 20-month period. All of these flights were successfully completed, and a great deal was learned during the Gemini Program that was later used in the Apollo Program, particularly for extravehicular activity, rendezvous techniques, mission control procedures, reentry control, and postlanding recovery operations. Figure 2 shows a picture of the Gemini spacecraft along with the adapter.

Apollo Program Although Project Mercury was still in the early developmental phase, NASA started considering more advanced manned missions during the spring and summer of 1959. The most obvious and appealing prospect was for some type of lunar mission. Three types of manned missions were considered. They were, in order of increasing difficulty, circumlunar flight, lunar orbit, and lunar landing. Regardless of the ultimate goal, it was generally felt that the first flight would be circumlunar, and the spacecraft would pass within several hundred kilometers of the lunar far side. After this, orbital flights would be made using the same outbound and homeward navigational techniques proven in the circumlunar missions. Finally, a lunar landing would be made by descending from lunar orbit. By this scheme, each mission would be an extension of the previous one; thus the overall difficulty of achieving the final goal would be divided into a number of

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Gemini equipment arrangement from press reference book for gemini spacecraft number 11 revision 30 august 1966 Coolant radiators

Propellant tanks

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Orbit attitude control thrusters (typical)

Communications equipment

Retrograde rockets Coolant pumps

Ejection seats

Cryogenic oxygen tank

Reentry attitude control system Parachute landing system

Horizon sensors Rendezvous radar

Maneuver thrusters (typical) Inertial guidance system Electrical equipment

Drinking water Electrical power system Instrumentation equipment

Figure 2. A cutaway picture of the Gemini spacecraft with the adapter section that attaches the spacecraft to the Titan launch vehicle. The position of the two astronauts in the spacecraft is also shown. This figure can also be seen at the following website: http:// www.hq.nasa.gov/office/pao/History/diagrams/gemini4.gif.

incremental steps; each would have greatly reduced exposure to the unknown. Nevertheless, the first time a plan to make a manned landing by descending from lunar orbit was outlined to NASA management, several in the audience severely questioned the wisdom of not taking advantage of the experience that would be obtained from Surveyor, which was designed to go directly from Earth straight down to lunar surface. Clearly, they had not considered the excitement of the crew during a landing that started at hyperbolic velocity in a near-vertical approach and that would be fully committed before they knew whether the landing rocket would fire. This incident is mentioned to illustrate that at the same time a manned lunar landing was seriously being debated, the basic understanding of the venture was quite primitive (3). Because it required the least amount of propulsion and the least sophistication in navigational and guidance equipment on board the spacecraft and it clearly seemed the least hazardous, the first mission seriously considered was circumlunar navigation and return. This was a modest extension of orbital flight; in fact, circumlunar flight is achievable by a highly eccentric Earth orbit of the proper parameters. However, the gravity field of the Moon creates a major influence on such orbits. Consequently, even the smallest error in the state vector at the time of translunar injection could not go uncorrected for safe entry into Earth’s atmosphere at the end of the mission. The question was how to determine the error and how accurately the corrections could be made. There was real concern that a safe entry into Earth’s atmosphere at lunar-return velocity might

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be beyond the guidance and navigational ‘‘state-of-the-art’’ technology. At these velocities, the spacecraft must pull negative lift to skim along a very narrow corridor of the upper layer of Earth’s atmosphere during the initial stages of entry. If the upper boundary of this corridor were exceeded, the spacecraft would skip out of the atmosphere back into a highly eccentric orbit and perhaps expend its supplies before entering the atmosphere a second time. On the other hand, if the corridor boundary were missed on the lower side, the spacecraft would exceed the heating or load limitations of its structure. An associated concern was entry and landing point location. Because the mission was not very well understood, it was conjectured that the time of return might vary greatly from the planned time and, because of Earth’s rotation, the geographical position of the entry might have a large dispersion. For these reasons, configurations with a fairly high liftto-drag ratio appeared desirable. In summary, the thinking in 1959 was that from the standpoint of flight control, the circumlunar mission would be flown using ground-based navigation obtained from tracking data. Guidance instructions would be transmitted to the crew for the necessary midcourse corrections. A budget for velocity changes of 152 m/s (500 ft/s), it was initially estimated, needed a lift-to-drag ratio of more than 1 to provide a large maneuvering footprint while safely staying within the entry corridors. During the 1960s, enough studies had been performed by NASA and industry to understand fairly well the implications of various manned lunar mission models. Missions that involved putting a spacecraft in orbit around the Moon for a total flight duration of about 2 weeks received a great deal of interest. Such missions did not appear to be a great deal more difficult than circumlunar flight but would provide much more scientific data. Furthermore, these missions would provide the means for gaining significant flight operational experience and reconnaissance data that would support a future lunar landing. At this point, the most important roadblock to executing of a lunar mission was that an enormous new space launch vehicle would be required. Another consideration was that features of the lunar surface were not well enough known to make detailed planning of a lunar landing possible. Both the United States and the Soviet Union were conducting robotic orbiting missions that would yield high-resolution photographs to make landing site selection possible. Also, hard landers (Ranger) and soft landers (Surveyor), both U.S. missions, would reveal the properties of the lunar surface in structure and hardness. On 25 May 1961, President Kennedy, in an address to the U.S. Congress, said, ‘‘y I believe that this Nation should commit itself to achieving the goal, before this decade is out, of landing a man on the Moon and returning him safely to Earth’’ (4). This commitment initiated the Apollo Program and precipitated some firm decisions. First and foremost, the return to Earth had to be determined because the safety of the crew ultimately depended upon this. The guidance and control precision of entry into the atmosphere at lunar-return velocity was sufficiently well established to commit to an entry configuration that would have a lift-to-drag ratio of 0.5 instead of the value of 1 previously mentioned. The selected value was compatible with the use of a semiballistic entry configuration design. Such configurations could achieve the relatively low entry heating of high-drag ballistic designs with a modest amount of lift. Furthermore, these features could be embodied in an axisymmetric shape, which simplified a number

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of design, manufacturing, and test considerations. The design chosen for Apollo was a derivative of the Mercury shape (5). By offsetting the center of mass by a distance of 19 cm (7.5 in) from the centerline, the Apollo spacecraft would trim at about a 331 angle of attack, which was sufficient to produce the desired L/D of 0.5. Using this much lift, the spacecraft could be confidently guided to land 9360 km (5000 nautical miles) downrange from the entry point. As various equipment items were designed for the entry capsule, the center of mass inexorably moved toward the center of volume. Consequently, the center of mass ended up displaced only a little more than 12.7 cm (5 in) from the centerline and the resulting lift-to-drag ratio was lowered to 0.35. However, this turned out to be more than sufficient. Planned splashdown for all missions actually flown was set for 2593 km (1400 nautical miles) downrange, and the never used capability either decreased it to 1482 km (800 nautical miles) or increased it to 4074 km (2200 nautical miles). It should be mentioned that only once in all of the returns from the Moon was it deemed desirable to move the preplanned landing point. It was moved 926 km (500 nautical miles) further downrange to avoid the possibility of predicted bad weather at the previously chosen landing point. However, the decision to relocate was made early, and the change was accomplished by a propulsion maneuver during trans-Earth coast a day before entry. Thus, the actual entry was standard at a nominal downrange distance. About a year after President Kennedy gave the go-ahead for the Apollo Program, a decision was reached on which launch vehicle to use. Originally, a very large launch vehicle called Nova was in the plan; Nova was substantially larger than the Saturn V rocket system that was actually employed (6). Nova was necessary because, originally, the entire spacecraft would be put down on the Moon. The judgment was made that the development of Nova would be too costly and would take long enough that President Kennedy’s deadline for the lunar landing could not be met. Using the less capable Saturn launch vehicle, two spacecraft would be necessary because a rendezvous maneuver would be employed to reduce the weight of the system that would have to be lifted to the Moon (Fig. 3). A debate ensued in 1962 whether the rendezvous between the two spacecraft should take place in Earth orbit or in an orbit around the Moon. The advantage of the former was that a spacecraft would remain in Earth orbit following the Apollo missions as a legacy, and many thought that was important. The principal advocate of the lunar orbit rendezvous was Dr. John Houbolt of the NASA-Langley Research Center (7). Eventually he made a compelling argument that only by adopting the lunar orbit rendezvous method would it be possible to meet President Kennedy’s deadline of 1970 for the landing on the Moon. From the standpoints of mission planning and guidance and navigation, lunar orbit rendezvous was completely compatible with all of the work that had been done by the time the decision was taken. The requirement of the rendezvous in lunar orbit did have a major impact on the equipment and the operational techniques that would be necessary for rendezvous navigation. The experience gained during the Gemini Program was extremely valuable in that connection. Furthermore, the general rules of the interplay between the mission control center in Houston and the astronauts were also developed during the Gemini Program.

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1 Pitch motor (solid) 13 300 newtons thrust 1 Tower jettison motor (solid) 178 000 newtons thrust Launch escape system 1 Launch escape motor (solid) 667 000 newtons thrust 379 Liters monomethylhydrazine (reaction control system) 227 Liters Nitrogen tetroxide (reaction control system) 9500 Liters Nitrogen tetroxide 8000 Liters hydrazine/unsymmetrical Dimethyl hydrazine Lunar module 3800 Liters Nitrogen tetroxide (Lunar module ascent/descent stage) 4500 Liters hydrazine/unsymmetrical dimethyl hydrazine (lunar module ascent/descent stage)

253 200 Liters liquid hydrogen

92 350 Liters liquid oxygen 95 Liters Nitrogen tetroxide (auxiliary propulsion system) 114 Liters monomethylhydrazine (auxiliary propulsion system)

Apollo command module 12 Control engines (liquid) 390 newtons thrust each 16 Control engines (liquid) 445 newtons thrust each Service module 1 Engine P-22KS (liquid) 97 400 newtons thrust 16 Attitude control engines (liquid) 445 newtons thrust each 1 Ascent engine (liquid) 15 700 newtons thrust 1 Descent engine (liquid) 4670 to 46 700 newtons thrust (variable) Instrument unit

Third stage 6 Attitude control engines (liquid) 654 newtons thrust each 2 Ullage motors (solid) 15 100 newtons thrust each 2 Ullage engines (liquid) 320 newtons thrust each 4 Retromotors (solid) 158 800 newtons thrust each 1 J-2 Engine (liquid) 889 600 newtons thrust

1 000 000 Liters liquid hydrogen

Second stage

101.6 meters 331 000 Liters liquid oxygen

8 Ullage motors (solid) 101 000 newtons thrust each 5 J-2 Engines (liquid) 889 600 newtons thrust each (later uprated to 1 023 000 newtons)

1 311 100 Liters liquid oxygen First stage

810 700 liters RP-1 (kerosene) 8 Retro motors (solid) 391 000 newtons thrust

5 F-1 Engines (liquid) 6 672 000 newtons thrust each (later uprated to 6 805 000 newtons)

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The two spacecraft required for the lunar orbit rendezvous were the command and service module (CSM) for flight to lunar orbit and return and the lunar module (LM) for descent to the lunar surface and return to lunar orbit (Fig. 4). Recognizing that the duration of the Apollo mission would necessarily be extended, a crew size of three was chosen, partly to accommodate the 4-hour-on, 8-hour-off duty cycle wherein one crew member would be on duty at all times. However, based on Gemini experience, the crew stated a decided preference to sleep, eat, and be on duty at the same time. Accommodating this preference to the mission was not considered unsafe. Ground controllers could easily monitor the condition of the spacecraft systems during the crew off-duty periods. Nevertheless, the choice of a three-man crew was well suited to the purposes of the program, and, as a matter of fact, it would have been extremely difficult to have performed the necessary tasks with less than three. Based on this number, one crewman was left in lunar orbit aboard the CSM while the other two members descended to the lunar surface in the LM. It was deemed highly desirable to have two crewmen on the lunar surface, particularly during the extravehicular activity, because this accommodated the buddy system, which materially added to the safety of that operation. The arrangement of the CSM was quite similar to that of the Gemini spacecraft. The command module (CM) housed the crew and was the only vehicle designed to reenter the atmosphere. Like Mercury, it was equipped with a launch escape rocket to be used in case of a mishap. Although the configuration was similar to that of Mercury, the conical afterbody was blunter to minimize heating on this portion at the 331 design angle of attack. A docking mechanism and tunnel were located at the apex of the cone to accommodate docking with the LM and crew transfer. The command module was much roomier and considerably more comfortable than the Gemini vehicle. (During other periods of flight, the center couch was removed to increase room for crew activity.) There was sufficient room under the remaining couches for two crew members to sleep. The service module, like the Gemini adapter, provided for propulsion and electric power during the mission. It was equipped with 16 reaction control thrusters arranged in four identical modules. These provided the thrust for rotational and minor translational maneuvers. The service propulsion system was equivalent to a propulsion stage. It was used during the mission for several minor maneuvers and two major ones: insertion into lunar orbit and departure from lunar orbit into trans-Earth flight. The service module also contained a thermal radiator and a high grain S-band dish antenna (Fig. 5). The LM (Fig. 6) was really a two-stage vehicle. The descent stage was equipped with a throttleable descent engine, a landing radar, and a four-legged

Figure 3. The Saturn V launch vehicle that was used during the Apollo program to take people to the Moon. There were two stages to take the spacecraft that would go to the Moon into near Earth orbit. A third stage, the Saturn IV B, propelled the spacecraft, consisting of the command module, the service module, and the lunar module, into a lunar trajectory. The Saturn IV B was jettisoned following the trajectory insertion, and it eventually crashed into the Moon. This figure also can be seen at the following website: http:// www.cr.nps.gov/history/online_books/butowsky4/images/space12c.jpg.

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Launch escape system

Command module

Service module

Spacecraft / lunar module adapter

Lunar module

Launch vehicle

Apollo launch configuration for lunar landing mission

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Nose cone and "Q - ball" Canard assembly Pitch control motor

Docking mechanism Drogue parachutes (2)

Main parachutes (3) Tower jettison motor

Hatch Aft compartment (tanks, reaction control engines, wiring, plumbing)

Crew compartment Launch escape system

Launch escape motor Electrical power system radiator panels (8)

Command module Service module

Fuel cells (3) Reaction control thruster assembly (4 locations) Cryogenic oxygen and hydrogen storage tanks

Helium tanks (2)

Reaction control system assembly (4 locations) (shown open)

VHF scimitar antenna (2) Environmental control system radiator panels (2) Launch escape tower

Service propulsion system tanks (4)

Forward boost protective cover

Service propulsion engine nozzle

Aft boost protective cover

High - gain (deep space) antenna Apollo command and service modules and launch escape system

Figure 5. A detailed cut-away drawing of the Apollo Command and Service Modules. The positions of the three astronauts in the Command Module are indicated. The launch escape system is also shown. This figure also can be seen at the following NASA website: http://www.hq.nasa.gov/office/pao/History/diagrams/ad004.gif.

landing gear. The legs had an extra large spread to accommodate the low impact stability associated with the very low gravity of the Moon. Each leg was tipped with a large dish-shaped pad, which was designed to accommodate the unknown bearing pressure of the lunar soil. The pad was socket-mounted to allow skidding at touchdown in any direction with little chance of tripping. The descent stage was designed to act as the launch platform for the ascent stage at the time of departure from the lunar surface. The ascent stage housed the ascent propulsion system and the LM cabin. There were accommodations for two crewmen, who were positioned at their flight

Figure 4. A picture of the Apollo spacecraft in the launch configuration. The Command Module, the Service Module and the Lunar Module are clearly illustrated. The legs of the Lunar Module folded so that the vehicle fit into the conical shroud on top of the Saturn V spacecraft. The escape tower is also shown. This figure also can be seen at the following NASA website: http://www.hq.nasa.gov/office/pao/History/diagrams/ad003.gif.

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Lunar module Antenna Eva Docking window antenna Antenna

Overhead hatch Docking target VHF antenna Thrust chamber assembly cluster (4) Plume deflector (4) Docking light

Rendezvous radar antenna Docking lights Tracking light Forward hatch

RTG fuel cask Landing probe (4)

Mesa Ladder

Landing gear Egress platform

Rover

Figure 6. A drawing of the Lunar Module in the landing configuration. The shaded portion is the descent stage which contains the rocket motor used for the landing. This portion of the spacecraft is left on the lunar surface following the completion of the extravehicular activity (EVA). The ascent stage also has a propulsion system that moves this vehicle back to lunar orbit where it then mates with the Command and Service Module. This figure also can be seen at the following NASA website: http://www.hq.nasa.gov/office/ pao/History/diagrams/ad014.gif.

stations in the standing position to maximize downward visibility during the landing maneuver. The crewmen slept in hammocks while on the lunar surface. Like the Mercury and Gemini spacecraft, the CSM and the LM had a pureoxygen atmosphere at a pressure of one-third sea-level pressure. The ascent stage was equipped with a rendezvous radar, 16 reaction control thrusters, and a high-grain S-band dish antenna. The LM was battery-powered and water-cooled. Ascent propulsion was used to launch the ascent stage from the lunar surface into a rendezvous orbit. It was also available to abort the landing maneuver into a rendezvous orbit if the landing could not have been made. After the CSM had established the desired orbit about the Moon, the LM pilot and the commander transferred to the LM and powered it up. The LM was then separated and the CM pilot was left alone in the CSM until the other crewmen returned from the lunar surface. The LM then made a phasing maneuver, by which its orbit relative to the CSM was lowered. This maneuver placed the LM sufficiently ahead of the CSM so that an aborted powered descent would terminate into a rendezvous-compatible orbit. At the proper time, powered descent was initiated with the descent propulsion system at nearly full throttle. Slight throttle adjustments were made during descent to keep the descent

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trajectory on target. As the targeted landing area was approached, the LM was pitched over while, at the same time, the engine was throttled back. This allowed the crew to achieve visual contact with the landing area several minutes before touchdown. At this time, the commander could determine the touchdown point that was being targeted by the onboard guidance system. He did this by looking through a scale painted on his window, called the landing point designator. He could then override the guidance system until the landing point designator indicated a desirable landing location. During this period, altitude and descent rate were sensed by the landing radar, and control by the guidance system was continued. Hovering flight was achieved several hundred feet above the lunar surface. The final descent and touchdown were then made under control of the commander. As the LM approached the surface, a considerable amount of dust was blown up, but some visibility remained, and good landings were made in all cases. It should be mentioned that several turbofan-powered hovering vehicles designed to mimic the landing characteristics of the LM were built. The crew used the lunar landing training vehicles extensively to sharpen their proficiency before each lunar landing mission. While on the surface of the Moon, the crew used extravehicular mobility units (EMUs) to move about. The EMUs each consisted of a space suit pressurized to 25 kPa (0.25 atm) and a battery-powered backpack. In many ways, the EMUs functioned like miniature manned spacecraft. They provided a cooled, revitalized atmosphere for the astronaut to breathe. The astronaut was kept at a comfortable temperature level by an actively cooled garment through which water was circulated at a regulated temperature. The EMU was equipped with redundant two-way communication links that provided voice communication between the two astronauts and, using the LM communication equipment as a relay station, communication with the Mission Control Center on Earth. Ground controllers could also monitor the physical condition of the astronauts and the performance of the equipment by telemetry from the EMU. The suits provided sufficient mobility to accommodate a range of activity on the lunar surface such as setting up science stations, taking photographs, and performing short geological traverses for sample selection. On later missions, lunar surface exploration was greatly enhanced by the Rover, an electric-powered four-wheel vehicle that could carry the astronauts and a considerable amount of equipment. The Rover carried two-way communications equipment and a high-gain antenna that provided sufficient bandwidth for a direct link to Earth carrying color or television transmissions as well as relaying the communications from the astronauts to Earth. It was recognized that both the CSM must be made considerably more reliable than previous spacecraft for several reasons. Compared to an orbital flight, a journey to the Moon implied a much greater exposure to the hazards of space. As a consequence of the complexity and the time required to abort a lunar mission, both the LM and the CSM had backup navigational and guidance systems, adequate reserve propellant, and levels of redundancy that were carefully determined to be sufficient. All possible failure modes were analyzed so that single-point failures were identified and eliminated when possible or safely accommodated. Furthermore, operational methods were worked out to circumvent failures or to ameliorate the conditions that might be brought about by failures.

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Although tracking from the ground was chosen as a primary method of navigation during the mission, it was decided that the spacecraft should be capable of autonomous navigation and guidance to return safely if all communication links were lost. The onboard navigation and guidance system was also used in verifying the accuracy of various maneuvers. Navigation in space, as on the ocean, requires precise determination of position. For this purpose, the Apollo command module was equipped with a sextant to measure the angle between a number of preselected landmarks on both Earth and the Moon and certain cataloged celestial bodies. Sightings and the time of sightings were fed into the onboard computer, which was preprogrammed to solve the navigational problem. The spacecraft was also equipped with inertial measurement units, which, together with the computer, performed the automatic guidance function. In addition to the launch vehicle and LM and CSM, there were two other vital elements in successful execution of the lunar missions. These were the Manned Space Flight Network (MSFN) and the Mission Control Center, including all of the mission procedures and time lines that were worked out by Mission Control Center personnel. The MSFN was modeled after the Deep Space Information Facility (DSIF) developed by the Jet Propulsion Laboratory. The MSFN consisted of three 25.9-m (85-ft) diameter antennas located at approximately 1201 intervals of longitude around Earth to provide coverage of the mission. Both the MSFN and the hardware onboard Apollo were capable of producing highly accurate navigational data. Data from both sources were compared before making any velocity change maneuver. Also, immediately after the maneuver was made, data were again cross-checked. Navigation done on the ground had the benefit of a large complex of powerful computers. Furthermore, at least two S-band trackers were always available as data sources. The data from the S-band tracker were extremely accurate. In addition to providing a Doppler count for velocity, the carrier signal was also phase-modulated with a pseudorandom noise (RN) code for range measurement. This digital signal, which was nonrepetitive for 5.5 s, was turned around and retransmitted on another carrier by a transponder on the spacecraft. Distance measurements as accurate as about 10 m could be obtained. Velocity measurements were much more useful. Highpowered data processing techniques could produce an accuracy better than a millimeter per second by smoothing Doppler data across a period of 1 min. Using such data, extremely accurate state vectors could be obtained not only on translunar and trans-Earth flight but also while Apollo was in lunar orbit. This capability was important because lunar gravitational anomalies and venting from the spacecraft continually perturbed the orbit. Computational techniques were developed to the point at which tracking data obtained from the lunar module during its landing descent burn could be processed to serve as a sufficiently accurate ‘‘tie-breaker’’ if onboard primary and backup computations produced unexplainable differences. The general approach to mission planning was to break the mission down into a number of discrete events and periods. Each of these was analyzed in great detail, and a complete model of the mission at great precision was constructed before flight. When flown, the missions would usually duplicate the plan in exact detail. A feature of the planning was the inclusion of time allowances for

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unexpected events so that the preplanned schedule could be maintained. The advantage was that almost every event or phase of the basic mission was extremely well understood and exercised. In addition to the basic mission plan, a great number of contingency plans was available to cover every rational problem. All missions were planned to accommodate midcourse corrections both outbound and on return. Specific times were set aside for these maneuvers: four translunar midcourse correction events were allowed and three trans-Earth. However, if the error to be corrected was sufficiently small, the maneuver would not be made, and, as a matter of fact, many missions needed only one corrective maneuver each way. When first considering translunar flight in 1959, the budget of 142.4 m/s (500 ft/s) for midcourse corrections was established. The estimate was down to 91.4 m/s (300 ft/s) when the program was initiated several years later. When actual flight began, the ‘‘three sigma’’ estimate was 23.8 m/s (78 ft/s). Actually, most flights required less than 6.1 m/s (2.1 ft/s). For example, on the last flight, Apollo 17 executed only one correction maneuver each way: translunar, it was 3.2 m/s (10.6 ft/s); for trans-Earth, only 0.6 m/s (2.1 ft/s) was needed, which shows the great improvement in navigational systems. The robotic missions that preceded the Apollo landings, Rover, Surveyor, and Lunar Orbiter have already been mentioned. The Ranger spacecraft was a probe that transmitted a few close-up images of the lunar surface just before colliding with the Moon at high velocity. The Surveyor was a soft lander that made five successful landings on the lunar surface. After landing, the Surveyor transmitted pictures of the surrounding features of the moonscape that provided extremely useful information on surface roughness as well as the quantity and size of rocks. Just as important, engineering data obtained from the Surveyor landings were extremely valuable in verifying the firmness of the lunar surface for landing the lunar module. The unmanned Lunar Orbiter flights, however, were every bit as valuable to the Apollo Program. They provided high-resolution photographs of the lunar surface that were extremely useful in selecting landing sites. The cartographic quality of the photographs was more than sufficient to make accurate maps of the lunar surface that could be used for orbital navigation and for visual recognition by the astronauts in the terminal phase of their descent. Just as important, analysis of orbital tracking data greatly improved the accuracy of the lunar gravitational constant and provided valuable data on lunar gravitational anomalies; all of them facilitated translunar and lunar-orbit navigation on the Apollo missions. The accuracy of navigational techniques that were ultimately developed made it possible for Apollo 12, the second landing mission, to come to rest within a short walking distance of Surveyor II, which had landed on the Moon 2 12 years previously. The Surveyor’s location had been identified on a Lunar Orbiter photograph by patient and meticulous study. Twelve unmanned test flights were made in the Apollo development program. All but one was successful. Six of these were devoted to qualifying the launch escape and parachute deployment system for the command module. The others were concerned with systems tests of the spacecraft and with compatibility of the launch vehicle and the launch environment. In one of these tests, the service propulsion system was used to drive the command module back into the atmosphere at the velocity and in the flight path expected during lunar return.

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Eleven manned flights were made in the Apollo lunar program. Nine of them were journeys to the Moon, and six were lunar landing missions. These flights are summarized in Appendix B. On 27 January 1967, the program underwent a critical setback during a countdown rehearsal for what was planned to the first manned Apollo flight. The interior of the command module suddenly burst into flames, and trapped the crew. Astronauts Gus Grissom, Ed White, and Roger Chaffee were killed, and the spacecraft was destroyed. Although the immediate source of the fire was never determined, the general cause was associated with the use of many materials in the cabin interior that were flammable in the pure-oxygen cabin atmosphere, which, during prelaunch conditions, was at sea-level pressure. It took the program more than a year to recover from the fire (8). Extensive changes in materials were made, and in cases for which suitable replacement materials could not be found, fireproof coatings and coverings were used. Special test programs were made to certify the fireproofing program. It was found that, although good fire protection was achieved for an oxygen atmosphere at the reduced pressure of orbital flight, there were no practical solutions for the prelaunch conditions when the cabin would be at sea-level pressure. Therefore, the oxygen in the cabin was diluted with 40% nitrogen during prelaunch activities. After the cabin pressure dropped to one-third of sea-level pressure during ascent, there was still sufficient oxygen for breathing. Later, when the astronauts were ready to depressurize further to spacesuit pressure, the nitrogen content of the atmosphere had been sufficiently purged to preclude a bends problem. When astronauts James Lovell, Jack Swigert, and Fred Haise were on board Apollo 13, which was to have been the third lunar landing mission, another accident occurred. During translunar flight as the Moon was being approached, one of the two tanks used to store cryogenic oxygen in the service module failed. The failure was sufficiently violent to cause the oxygen in the other bottle to begin leaking, most probably through an external line. Except for gaseous oxygen stored in the command module for life support during the short period of reentry flight, there was no other oxygen supply for the CSM. The lost oxygen was to be used both for life support and to power the fuel cells. The failure was eventually traced to a thermostat switch that had been frozen shut because of an overcurrent applied during the preflight checkout on the launch pad. As a result, the heater wire inside the oxygen tank became too hot, and the fluorocarbon insulation melted causing a spark. This ignited the insulation which readily burned in the pressurized oxygen tank, and the resulting pressure rise caused the tank to fail. This desperate situation was countered by moving the crew to the LM cabin and powering down the CSM. Because the LM was battery-powered for only a planned 3-day period, it was configured for minimum power consumption. After getting everything in order, the crew made a velocity change maneuver to correct the flight path to a ‘‘free-return’’ trajectory by which the spacecraft would swing around the Moon and head back to Earth on a path consistent with reentry targeting. Another major propulsive maneuver was made after the spacecraft was in trans-Earth flight. That maneuver changed the trajectory to shorten the return time. Although there was sufficient oxygen for the return flight, an improvised method for adapting the command module’s lithium hydroxide canisters

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used to remove carbon dioxide for use in the LM environmental control system had to be devised. This was done by experimenting on the ground and then communicating instructions to the crew. Fortunately, the crippled spacecraft returned safely with an unharmed but disappointed crew. The Apollo program continued for four additional and successful lunar landings.

The Skylab Program After the completion of Apollo 17, the Apollo CSM and the Saturn launch vehicles were used for the Skylab Program (Fig. 7). The third-stage structure of the Saturn V was converted to a space laboratory. The hydrogen tank of the Saturn IV B stage was divided into laboratory, sleeping, eating, recreation, and hygiene compartments. Storage lockers, an environmental control system, and other equipment were added for the comfort and physical well-being of the crew. The oxygen tank was converted to an oversize trash container. A deployable micrometeoroid shield and passive thermal control coatings were added to the skin of the tank. A battery of telescopes designed to study the Sun were mounted on a high-precision pointing platform. There was also an airlock to accommodate EVA and a docking module to which the CSM could dock. In addition to the solar telescopes, a great variety of experimental equipment was carried, including a number of Earth-pointed remote-sensing instruments and cameras.

Lunar module/Apollo Telescope Mount

Service module Command module

ATM solar array panels Orbital workshop

OWS solar array panels Instrumentation unit Spacecraft LM adapter (fixed) Airlock module Multiple docking adapter

Structural transition section

Figure 7. The Skylab orbital workshop configuration. The Apollo Command and Service Module is on the left, the center section is the Apollo Solar Telescope Mount, and the lefthand section shows the Modified Saturn IV B that became the living space for the three astronauts during the three missions when they worked in this temporary space station. This figure also can be seen at the following NASA website: http://www.hq.nasa.gov/office/ pao/History/SP-4011/p144.htm.

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The Skylab orbital workshop was stabilized using three control moment gyros. These could be desaturated from time to time by reaction control jets using nitrogen as a propellant. Electric power for the Skylab was provided by two photovoltaic solar-powered systems. One system consisting of four folding arrays was attached to the solar telescope mount. The other system consisted of two wing arrays extending from the converted fuel tank (10). The Skylab orbital workshop was launched on 14 May 1973 into a 501 inclination orbit. The nearly circular orbit ranged in altitude from 496 to 498 km (268 to 269 nautical miles). During launch, one of the wing solar cell arrays and the micrometeoroid shield were carried away. The other wing was jammed so that it could be only partly deployed. The micrometeoroid shield also was important in the passive thermal control system; its absence left the tank skin exposed. This was coated with a highly reflective coating which happened to be biased hot. The result was that the laboratory quickly became overheated and reached internal temperatures higher than 501C. The launch of the first crew was delayed while special repair equipment was designed, constructed, and tested on the ground. Eleven days following the launch of the Skylab orbital workshop, astronauts Pete Conrad, Joe Kerwin, and Paul Weitz were launched on Skylab 2. They deployed a parasol-like sunshield over the exposed skin which reduced the interior temperature to 241C. They also freed the jammed wing of the solar cell array. Although the missing wing reduced the power available from that expected, there was sufficient power to perform the planned activity. The Skylab orbital workshop was occupied by three different crews (See article on Skylab elsewhere in this Encyclopedia). Each successive crew stayed a longer period of time. A great number of experiments and science investigations was performed before the Skylab orbital workshop was abandoned on 8 February 1974. It was left in a condition that would allow partial reactivation in a possible visit from the Space Shuttle. Unfortunately, its orbit decayed before that was possible. Skylab reentered the atmosphere on 11 July 1979, and came to Earth mostly in the Indian Ocean, with some pieces landed in western Australia.

Apollo-Soyuz Project After a series of meetings in 1971 and 1972, the United States and the Soviet Union agreed to a joint manned space mission as part of the ‘‘de´tente’’ policy pursued by the Nixon administration. This mission became known as the ApolloSoyuz Test Project (ASTP) (Fig. 8). Its basic purpose was to produce hardware and operational procedures that would provide the means for one country to work with the other in future manned space missions. Particular emphasis was placed on assistance and rescue missions. Rather than merely trade design and mission control information, it was agreed that the only way to be sure assistance and rescue missions would really be feasible would be to execute an actual joint mission (11). It was agreed that the American Apollo and the Russian Soyuz would rendezvous and dock during a cooperative mission. To accomplish this objective, a number of technical problems had to be addressed. The primary problems were

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New compatible docking system

Soyuz

Docking module

Apollo

Apollo -Soyuz rendezvous and docking test project

Figure 8. The Apollo-Soyuz spacecraft. The Apollo Command and Service Module, the Docking Module, and the Soyuz capsule are all clearly shown. Two Soviet cosmonauts and three American astronauts performed the docking maneuver that led to the first meeting in space of people from each nation. Experiments and tests performed by the crew led to naming this effort the Apollo-Soyuz Test Program. This figure can also be seen at the following NASA website: http://www.hq.nasa.gov/office/pao/History/diagrams/ astp/pk69.htm.

compatible communications both in space and on the ground, compatible mission control procedures, compatible docking hardware, and accommodations for the different atmospheric constituents in the two spacecraft. The Soyuz cabin was maintained at sea-level pressure with the same 80%/20% mix of nitrogen and oxygen found in air. The Apollo cabin was at one-third sea-level pressure using an atmosphere of 100% oxygen. This incompatibility was solved by using a docking module that would be carried with the CSM during launch. It was carried in the adapter in a manner similar to that used to carry the LM during lunar missions. On one end of the docking module was the international docking system, which had been agreed on with the Soviet Union. The other end, like the LM, had the portion of the Apollo probe and drogue docking system. The docking module had an interval volume sufficient for as many as three people at the same time. It was also equipped with bottled supplies of both nitrogen and oxygen. Consequently, it could serve as an airlock between the two spacecraft. The ATSP mission was successfully completed in July 1975 by the launch of cosmonauts Alexei Leonov and Valery Kubasov in Soyuz followed by astronauts John Stafford, Vance Brand, and Deke Slayton in Apollo. After rendezvous, both the American and the Russian mechanisms performed successful dockings. After the final undocking, both crews spent additional time in space performing individual experiments.

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The Space Shuttle Program The Space Shuttle system had been developed to improve accessibility greatly to space and thereby to facilitate many opportunities for space applications that would not be possible otherwise (12). Through improved operational capability and flexibility, the Shuttle would bring about a wide variety of missions and activity in space that previously had not been considered. The Shuttle missions primarily fall into three general categories: 1. Missions lasting from one to several weeks and employing one or more of the Spacelab modules developed by the European Space Agency. These missions consist of a variety of experiments and observations requiring extensive participation by the crew. 2. Missions employing one or more additional propulsion stages to carry spacecraft to orbits beyond the performance capabilities of the Shuttle alone. Most of these missions are aimed at geosynchronous orbit; some planetary exploration and military missions were also launched by the Shuttle. 3. Missions in which the Shuttle deploys satellites directly into orbit. The Shuttle can retrieve or service some of these satellites, thereby adding a new dimension of utility. In some cases, the satellites will be equipped with modest propulsive capability allowing them to move back and forth to orbits beyond the operating altitude of the Shuttle. The first example of this capability was the retrieval and repair of the ‘‘Solar Max’’ satellite in 1984 and later, the spectacular success of the in-orbit repair of the Hubble Space Telescope’s optics in 1993. The most recent operation was the repair and refurbishment of the Hubble Space Telescope in 2002 (13). Most Shuttle missions probably have been a combination of these general types to enable more comprehensive use of the cargo load of each flight. When launched from Cape Kennedy, orbits with inclinations from 28.51 (the latitude of the launch site) to 561 will be achievable. The Shuttle was initially capable of carrying as much as 29,000 kg (32 tons) of payload into low-inclination, low-altitude orbits. It was capable of returning as much as 14,500 kg (16 tons) from orbit. However, various restrictions have been placed on the payload capability of the Space Shuttle since operations of the vehicle were initiated in 1981. The Space Shuttle in the launch configuration consists of four major elements: an Orbiter, an external tank, and two solid-rocket boosters (Fig. 9). The boosters provide the majority of the thrust for the first 2 minutes of flight. Each booster weighs about 567,000 kg (1,250,000 lbm) and produces a peak thrust in excess of 11.1 MN (2,500,000 lbf). The boosters are jettisoned after burnout and are recovered for reuse after descending by parachute into the ocean near the launch site. The external tank contains 530 cubic meters (140,000 gal) of liquid oxygen and 1438 cubic meters (380,000 gal) of liquid hydrogen. These propellants, with a combined weight in excess of 680,000 kg (1,500,000 lbm), feed the three main engines in the Orbiter. Eight minutes after liftoff, when main propulsion burnout

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Figure 9. The configuration of the Space Shuttle system with the Orbiter, the two solid rocket motors, and the external tank for the fuel and the oxidizers for the Space Shuttle main engines mounted in the Orbiter. This figure also can be seen at the following NASA website: http://www.hq.nasa.gov/office/pao/History/rogersrep/v1p3.jpg.

occurs, the Shuttle is just short of attaining orbital velocity. The tank is then jettisoned and follows a long, shallow trajectory that ends in a remote portion of the Indian Ocean. The tank is destroyed during reentry. After the tank is jettisoned, the Orbiter employs the two orbital maneuvering engines to propel itself into orbit. The orbital maneuvering engines (OMES) are also used for other maneuvers in space during the remainder of the mission. The Orbiter is the most complex part of the Shuttle and provides control of the boosters and the external tank during launch. The Orbiter is designed to be reusable; only a minimum amount of refurbishment and maintenance are required between flights. Almost all of its systems should function reliably after a hundred or more flights. Insofar as practical, development of new systems or technology was avoided in designing of the Orbiter. The Orbiter’s electric power is produced from hydrogen and oxygen in fuel cells that are an improved version of those used during the Apollo Program. The basic structure of the vehicle is primarily aluminum; skins, stringers, frames, and ribs are in an arrangement typical of that of an airplane. The landing gear and the cockpit arrangement are also quite similar to those found on conventional aircraft. Hydraulic power is used to move flight control surfaces and for ground steering and braking. The Orbiter is a key element of a highly complex launch vehicle. It is a unique and highly versatile spaceship that has astronauts on board to operate the vehicle and to perform experiments. It is the combination of these functions—

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launch vehicle, spacecraft, and airplane—that make the Orbiter a most complex machine. Consequently, major advancements in the state of the art were necessary in several areas. These advances were most pronounced in flight control, thermal protection, and liquid-rocket propulsion technology. (See the article on Space Shuttle Orbiter elsewhere in this Encyclopedia.) The external tank is 45.7 m (150 ft) long and 8.5 m (28 ft) in diameter. The structure of this huge tank weighs only slightly more than 31,750 kg (70,000 lbm). Nevertheless, the tank is the central element of the launch configuration, which has a combined weight of nearly 2,000,000 kg (4,500,000 lbm). The consequence is the creation of a number of easily excited low-frequency structural modes. Therefore, the flight control system employs sophisticated mode-suppression techniques in gimbaling the rocket engines as the vehicle is steered during launch. Flight control during entry is even more complex. Unlike previous manned spacecraft such as Apollo, the Orbiter actually is flying during atmosphere entry and depends on active aerodynamic control surfaces to maintain a stable attitude and control of its flight path. When aerodynamic flight begins, it is flying four times faster than the X-15, the previous speed record holder. The Orbiter’s configuration features a double delta wing, a single, high, vertical fin, a huge cargo compartment, and a large aft end containing an assembly of rocket engines. This configuration is based on a compromise satisfying a large number of requirements. The cargo bay size was derived from a survey of potential traffic that clearly indicated that space cargo would in general consist of low-density objects. Large main propulsion engines were desirable to minimize the size of the boosters, which represent a major part of the cost per flight. The result was a center of mass that is unusually far aft. Depending on mass and location or cargo, the Orbiter occasionally encounters neutral to slightly negative static stability about the pitch axis at low speeds. For control during aerodynamic flight, the Orbiter is equipped with very large elevons on the wing trailing edge. A large flap extends from the rear of the fuselage to trim the vehicle. A combination rudder and speed brake on the trailing edge of the vertical fin is used for directional and speed control. Return from orbit is initiated by making the typical retrograde propulsion maneuver using the OMES engines. The Orbiter then loses altitude and gradually enters the atmosphere. Initial control, similar to that of Apollo, is by rolling about the velocity vector at a very high angle of attack using the reaction control system. As aerodynamic forces gradually increase, the control of the vehicle becomes manageable using aerodynamic control surfaces, and, except for those controlling yaw, the reaction control jets are deactivated. The rudder, which might be expected to control yaw, is ineffective as a control surface on the Orbiter because of the high angle of attack. Both yaw jets and the rudder were active for yaw control from Mach 3.5 to Mach 1.0. Below that speed regime at a lower angle of attack, the yaw jets were deactivated, and only aerodynamic control surfaces were employed in the flight control system. Reentry at the end of the flight is made using a preselected angle-of-attack (a) profile. An angle of attack of 401 is maintained until the Orbiter decelerates to M ¼ 11. From that point on, a is decreased continually and passes through M ¼ 1 at a ¼ 81. The approach and landing maneuvers in the subsonic flight regime

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ranged in a from 41 to 101. The acceptable reentry altitude corridor is approximately plus or minus 1500 m (5000 ft) of the nominal flight path. The Orbiter is guided to stay within this corridor, while flying on the preselected a profile, by varying the bank angle. The nominal angle at the start of reentry is 801. This value gradually diminished to about 451. During the first flight, the landing site was essentially straight downrange. Consequently, some portions of reentry were flown banked left and others banked right. In this case, bank-reversal maneuvers were made near numbers of M ¼ 18, 9, 5, and 2.5. In future flights, longer portions of reentry may be spent banked in one direction or the other to accommodate landing at sites which may be displaced as much as 2037 km (1100 nautical miles) on either side of the straight downrange track. Downrange navigation is achieved by flying nearer to the lower or upper bound of the corridor. This can be accomplished by using a slightly greater or less than nominal bank angle, as the case may require. During the first Shuttle flight, John Young went to control-stick steering during the last two bank-reversal maneuvers. This allowed him to enter the bank-reversal maneuver more gradually than possible when the maneuver was programmed in the autopilot. Recent simulations had shown that the stability margins could be increased by decreasing the roll-rate acceleration during this flight regime. The decision to have the crew make this maneuver avoided modifying mature software. The entry is flown by computer until subsonic flight when the pilot uses control-stick steering and makes the approach and landing. The Orbiter configuration was subjected to more hours of wind tunnel testing than any other flying machine. This was necessary because the Orbiter flies across a greater range of Mach numbers, Reynolds numbers, dynamic pressures, and angles of attack than previous machines. More importantly, during its first flight, it was totally committed to successfully negotiating this wide and diverse range of flight conditions before it could land. This was a unique situation not previously encountered in modern aviation flight-testing. High-performance airplanes and in fact virtually all newly designed airplanes are extensively flighttested before maximum performance is approached. Each flight test is a cautious extension of previously encountered conditions and is followed by thorough analysis of test results from which the next safe increment in the flight-test program can be defined. As a substitute for testing in actual flight, the Orbiter’s flight control system had been extensively tested in computer-driven simulation facilities. Mathematical models resident in the facility software mimic the atmospheric conditions of flight, including gusts and crosswinds. Other models represent the wind-tunnelderived aerodynamic responses to the simulated flight environment. However, wind-tunnel data have limited precision. Results obtained in different facilities sometimes were significantly different. Furthermore, experience has shown that data obtained in flight often exceed the bounds of data scatter of wind-tunnel data. Extensive statistical analyses of these effects on the stability and the control surface effectiveness were made to determine possible worst case aerodynamic qualities. The Orbiter’s flight control system successfully ‘‘flew’’ simulated entries in which such worst case conditions were modeled. The heart of the Orbiter’s flight control system is a set of five identical general-purpose computers. Each computer in the set has the capacity to control

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the entire flight, including the guidance and navigation functions, without assistance from the others. Under control of the central computers are a great number of subordinate data processors that have specialized functions. The central computers formulate steering commands for the gimbal actuators during launch, the reaction control jets during space flight, and the yaw jets and aerodynamic surfaces during aerodynamic flight. These steering commands can originate in response to stick inputs from the crew or be generated within the central computers in the fully automated flight mode. Redundant sets of rate gyros, accelerometers, inertial measurement units, star trackers, radio navigational aids, radar altimeters, and dynamic and static pressure sensors all feed data to the central computers. The more critical sensors are quadruply redundant and allow at least two and possibly three sensors in a particular set to fail without degrading the performance of the flight control system. The five centralized computers are divided into a redundant set of four and the fifth computer is held in reserve as an independent backup. Normally the redundant set will control the flight. Each of the four computers in this set is loaded with identical software, and all work in parallel with each other processing data from the sensors and transmitting commands to the controls. If one of the sensors should fail, each computer will detect it, and the sensor will be deactivated. Upon completion of every computation cycle, the redundant computers make a simple arithmetic comparison of their output. Any computer that fails to agree with its colleagues for two successive cycles is considered to have faulted and is deactivated. The computers also have built-in fault-detection features which should cause them to self-deactivate in most failure cases. Simple logic would indicate that this system with its depth of redundancy and its faultdetection and deactivation features should prove extremely reliable. Nevertheless, there is concern that despite all precautions, an inherent weakness could exist in the software that would ripple through the computers and deactivate them one after another. Such should an event occur, the crew would switch control to the backup computer, which is loaded with software that, although coded differently, can replace all of the functions of the redundant set. The use of digital computers in the Orbiter for flight control provides a major improvement in versatility and precision over other systems. The digital system facilitates redundancy management, fault detection, and fault isolation and, when proven, should provide a major improvement in flight safety. Although the inherent complexity of the Shuttle flight control task forced the development of an unusually elaborate flight control system, this pioneering effort proved to be a major step toward the production of more economical and safer aircraft. The propellant combination of liquid hydrogen and oxygen is the most energetic considered practical for use. The Shuttle, like the Saturn launch vehicle upper stages, employs these propellants for its main propulsion. However, a new high-performance engine was developed for the Shuttle (Fig. 9). It features a combustion chamber pressure of 22,409 kPa (3250 psi) at full power, or about four times higher than that of the Saturn engines. It also incorporates a new cycle that circumvents the wasteful use of some of the propellant just to power the turbine-driven propellant pumps. It might be mentioned that the total horsepower required to pump propellant to the Orbiter at full power is greater than the total propulsive power of a Forrestal-class carrier. The Shuttle rocket engines

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produce 7% more propulsive energy per pound of propellant than the hydrogenfueled rocket engines used on Saturn. They also have a higher thrust-to-weight ratio. A consequence of these factors is a net gain of 18,144 kg (40,000 lbm) of payload compared to what could have been obtained using Saturn-type engines. The penalty for this performance is a considerably more complex engine that was difficult to develop and test. The Orbiter is easily the largest man-made object to be recovered from space. Seventeen times heavier and with 50 times the surface area, it greatly exceeds the size of the Apollo command module, which previously held this distinction. Not surprisingly, the creation of a thermal protection system suitable for reuse after multiple reentries required a new material technology. Fundamentally, the choice lay between insulating the inner structure from a hot external skin or using external insulation that would keep the metallic skin cool enough to be part of the primary structure, as in conventional aircraft. The second approach was chosen as both lighter and less costly to produce. The cost savings were primarily associated with the straightforward use of conventional structural materials and manufacturing techniques commonly used in modern large airplanes. An equally important benefit was the large database from which the cost and weight of the major structural assemblies could be estimated. It also allowed additional time to analyze the thermal input because of the isolation of the structural and thermal protection system. Except for a few regions in which the surface temperature will not exceed 644 K (7001F), the majority of the Orbiter’s external surface is covered with lightweight tiles. There are more than 32,000 tiles, most of which are 15.2- or 20.3-cm (6 or 8 in) squares. The basic material of the tiles consists of a random matrix of fine quartz fibers. The compaction of the fibers and therefore the density is closely controlled during manufacture. Each tile is coated with a pigmented glaze that provides water proofing, abrasion resistance, and high heat radiation. The density of most of the tiles is comparable to that of balsa wood, although some tiles of a higher density are located in regions where expected physical loads or abuse may be greater. Because the tiles have a lower coefficient of expansion than the aluminum skin to which they are bonded, a felt pad is inserted between the tile and the skin to isolate strains due to the relative motion between skin and tile. The tiles should be suitable for 100-mission reuse if they never exceed a surface temperature of 1533 K (23001F). If heating rates are higher, fewer reuses will be possible. One of the very good properties of the tile material is its capability of withstanding heating rates almost double its design value and still survive one flight. This was particularly reassuring during the first flight, for which heating rates could only be predicted based on a combination of theoretical analysis and wind-tunnel data. The worst case heating predictions fell well below the one-time capability of the material. On 28 January 1986, the space shuttle, ‘‘Challenger,’’ was destroyed by a failure of the seal on one of the solid rocket motors (SRM) that is intended to boost the vehicle into Earth orbit. Six astronauts and a schoolteacher acting as a ‘‘payload specialist’’ were on board ‘‘Challenger.’’ All were killed. There is no doubt that this accident was the single most traumatic event in the effort to put people in space. The failure of the SRM occurred when the joint between the lowest and the next-to-lowest segment of the rocket stack failed and released the hot gases

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inside the rocket (the location of this joint—called a ‘‘field joint’’ because it is made up in the field—is shown in Fig. 10). These hot gases rapidly melted the fixtures that attached the SRM to the fuel/oxidizer tank and caused the SRM to deflect and puncture the tank. The release of the hydrogen and oxygen quickly created a flammable mixture that ignited and destroyed the Orbiter. A commission to investigate the accident chaired by former Secretary of State William P. Rogers was established. The commission found that the proximate cause of the accident was that the O-ring seal of the joint opened when subjected to the pressure of the gases inside the rocket. This was attributed to a design flaw in the O-ring seal that was exacerbated by the cold temperature on the day of the launch, which embrittled the O-ring material of the seal. The report of the commission also listed as possible contributing causes the problems encountered in the assembly of the rocket stack and the very high wind shear experienced at approximately the same time as the joint failed. The commission recommended that space shuttle flights be suspended until the seal of the joint was redesigned, manufactured, and tested so that failures of this kind would not be possible (14). It took more than 2 years to implement this recommendation, and Shuttle flight operations resumed in September 1988. The ‘‘Challenger’’ was replaced by a new space shuttle, ‘‘Endeavor,’’ which flew for the first time in May 1992. Since September 1998, over one hundred space shuttle missions have been successfully executed. During these flights, many important in-orbit capabilities have been demonstrated, including the repair and retrieval of satellites by astronauts working outside the Space Shuttle cabin [extravehicular activities (EVA)], the execution of many onboard experiments Nozzle and thrust vector control system

Cutaway view of the solid rocket booster showing solid rocket motor propellant and aft field joint Separation motors (4) 22,050 Ib thrust each Solid rocket motor aft field joint

Aft skirt and launch support Booster-external tank attachment ring, aft avionics and sway braces

Solid propellant

Dimensions Length ............ 149.16 ft (45.46 m) Diameter ............ 12.17 ft (3.70 m)

Main parachutes (3) Separation motors (4) 22,050 Ib thrust each

3.8 m (12.4 ft) outside diameter

SRB-external tank thrust attachment

Drogue chute Forward skirt Nose fairing

Rate gyro assemblies (3), separation avionics, operational flight instrumentation, recovery avionics, and range safety system

Figure 10. A cutaway drawing of the solid rocket motor of the Space Shuttle system. The joint that failed during the launch of ‘‘Challenger’’ is the ‘‘aft field joint’’ labeled on this drawing. The picture also shows the configuration of the rocket with the nozzle, the four segments of the solid fuel grain, the separation motors, and the location of the control system in the nose section of the rocket. This figure also can be seen at the following NASA website: http://www.hq.nasa.gov/office/pao/History/rogersrep/v1p56.jpg.

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using the European Spacelab modules that were built for the Shuttle by the German space agency, docking with the Russian MIR space station, and launching important scientific satellites. The Space Shuttle fleet is currently providing the principal means for lifting many of the component parts of the International Space Station and all of the people who will assemble the station into Earth orbit. The Shuttles ‘‘Atlantis,’’ ‘‘Discovery,’’ and ‘‘Endeavor’’ have been uprated to deliver payloads to the higher inclination orbits of the International Space Station. The main engine is being modified and provisions such as docking hardware are being added. The external tank has also been redesigned primarily using a lithium aluminum material to increase payload capability. When the International Space Station is completed, the Space Shuttle will, for the first time, actually become what it has been called for so many years: the vehicle that will ‘‘shuttle’’ people and supplies back and forth to the International Space Station. The Space Shuttle represents a bold major development in the technology of human spaceflight. It has made possible activities and achievements in space that would not have been done otherwise. It has been and will continue to be the major new stepping-stone to a new era in space-operations.

Conclusions The enterprise of putting humans is space has been controversial from the very beginning. Shortly after the Mercury project was initiated in 1959, President Kennedy’s science advisor, Dr. Jerome B. Wiesner, recommended giving serious consideration to canceling the program. In a report he submitted to President Kennedy shortly before his inauguration in January 1961, Dr. Wiesner argued that human spaceflight was very dangerous and could damage the new administration’s prestige if some people were killed by accidents in space. Furthermore, he said that putting people in space is an expensive proposition and that the new administration’s scientific and military objectives could best be achieved using robotic spacecraft. There is a kernel of truth in this argument, which is why it is still being heard now even after four decades of successful human spaceflight. What the people who make this argument either forget or ignore is the public impact of human spaceflight. The world craves heroes, and there is no doubt at all that our astronauts are modern heroes. John Glenn is a recent example. He was the first American to orbit the Earth in 1962 and flew again on the Space Shuttle, ‘‘Discovery,’’ in 1999 when he was 77 years old. Successive administrations have been advised by very prestigious people to abandon the program to put people in space. Fortunately, this advice has been roundly ignored, and successive Presidents since John Kennedy have strongly supported the program to put people in space. A good case can be made that 500 years from now, the only thing that will be remembered about the twentieth century is that on 20 July 1969, human beings first set foot on another body in our solar system. The lunar landing in 1969 is likely to be remembered just the way 1492 is the only year in the fifteenth century that the vast majority of people around the world remember. Christopher Columbus’ landing on an island in the Bahamas has had the same impact as Neil Armstrong’s landing on the Moon. Both of these events opened new horizons for human beings to aim at, and this is why they are both historically significant.

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BIBLIOGRAPHY 1. Swenson, L.S., J.M. Grimwood, and C. Alexander. This New Ocean: A History of Project Mercury. 2. Hacker, B.C., and J.M. Grimwood. On the Shoulders of Titans: A History of Project Gemini. 3. McDougall, W.A. The Heavens and the Earth: On the Shoulders of Titans: A History of Project Gemini, NASA SP-4203, U.S. Government Printing Office, Wasington, DC, 1977. 4. Sidey, H. Time Magazine, Nov. 14, 1983. 5. Ertel, I.D., and M.L. Morse. The Apollo Spacecraft: A Chronology. NASA SP-4009, US Government Printing Office, Washington, DC, 1969. 6. Bilstein, R.E. Stages to Saturn: A Technological History of the Apollo/Saturn Launch Vehicles. NASA SP-4206, U.S. Government Printing Office, Washington, DC, 1980. 7. Hansen, J.R. Spaceflight Revolution: NASA Langley Research Center from Sputnik to Apollo: A History of Manned Lunar Spacecraft. NASA SP-4205, U.S. Government Printing Office, Washington, DC, 1995. 8. Brooks, C.G., J.M. Grimwood, and L.S. Swenson, Jr. Chariots for Apollo: A History of Manned Lunar Spacecraft. NASA SP 4205, U.S. Government Printing Office, Washington, DC, 1979. 9. Lovell, J., and J. Kluger. Lost Moon: The Perilous Voyage of Apollo 13. Houghton Mifflin, Boston, New York, 1994. 10. Compton, W.D., and C.D. Benson. Living and Working in Space: A History of Skylab. NASA SP-4208, U.S. Government Printing Office, Washington, DC, 1983. 11. Foehlich, W. Apollo-Soyuz. NASA EP-109, U.S. Government Printing Office, Washington, DC, 1976. 12. Heppenheimer, T.A. The Space Shuttle Decision: NASA’s Search for a Reusable Space Vehicle. NASA SP-4221, U.S. Government Printing Office, Washington, DC, 1999. 13. Chaisson, E.J. The Hubble Wars. Harper Collins, New York, 1994. 14. Report of the Presidential Commission on the Space Shuttle ‘‘Challenger’’ Accident. Washington, DC, June 6, 1986, Vol. I.

MAXIME A. FAGET MILTON A. SILVEIRA Formerly with the Engineering and Development Directorate NASA-Johnson Space Center Houston, Texas

URANUS AND NEPTUNE

Uranus and Neptune, were the first planets discovered by telescopic observation. They are Jovian planets of similar size and much more massive than Earth (by factors of 14.5 and 17) but are still far less massive than Jupiter (by factors of 22 and 18). Both have deep atmospheres dominated by hydrogen, but unlike Jupiter and Saturn, most of their mass consists of rocky and icy components, which are

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in fluid form due to high interior temperatures. Both have ring systems, a diverse system of satellites, unusual tilted and offset magnetic fields, and similar atmospheric circulations. Their blue colors distinguish them from the pale tans of Jupiter and Saturn. But they also differ significantly from each other; they have very different obliquities (spin axis inclinations relative to their orbital planes), vastly different internal heat fluxes, and remarkably different weather patterns. Their orbital and physical parameters are summarized in Table 1.

Discovery Uranus, at visual magnitude 5.5, is about as bright as Jupiter’s brightest moon and is close to the limit of visual detectability (magnitude 6). Neptune (magnitude 7.85) is far too dim (by a factor of 30) to be seen by a visual observer. Thus, it is not surprising that the discovery of both planets awaited the development of advanced telescopes. Although observed as early as 1690 by Flamsteed (4) and assumed to be a star, it was not until 13 March 1781 that Uranus was discovered by William Herschel, the best telescope maker of his time. Its extended image

Table 1. Orbital and Physical Parameters of Uranus and Neptune Uranus Orbital Parameters: Mean orbital radius, AUa,d Orbital eccentricitya,e Orbital sidereal Period (Y) from ssd.jpl.nasa.gov (May 22, 2001 update) Orbital inclination to ecliptica,f Planetary physical parameters: Massg (Earth masses, Earth ¼ 5.974  1027 g) Equatorial radius at 1 bar level, kmb Polar radius at 1 bar level, kmb Ellipticity (Requator/Rpole  1) Obliquity (tilt of rotational pole to orbital pole)c Sidereal rotation period of interiorb Mean densityg, g/cm3 Equatorial surface gravity, m/s2, 1 bar levelg (Earth ¼ 9.78) a

Neptune

19.19126 0.04717 84.017

30.06896 0.00859 164.791

0.771

1.771

14.535 25,559 24,973 0.0229 97.861 17.240 h 1.318 8.69

17.141 24,766 24,342 0.0171 29.561 16.11 h 1.638 11.00

Ref. 1., p. 316, Table 5.8.1. Ref. 2. c Here the rotational pole and orbital pole vectors are both defined by the right hand rule: the pole direction is such that when the pole vector points at the observer, the observer will see counterclockwise motion. In the case of Uranus, the rotational pole points opposite to the planet’s North Pole, as defined by the International Astronomical Union (IAU). Atmospheric dynamicists often find it more convenient to treat the rotational pole as the North Pole. d One AU ¼ 149.60  106 km. e Eccentricity ¼ distance between ellipse foci divided by 2  semimajor axis. f Ecliptic plane ¼ plane of earth’s orbit. g Ref. 3. b

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implied to Herschel that it was not a star, though at first he believed that it was a comet. Herschel was an amateur at the time, but later became the first president of the Royal Astronomical Society. This was the first planet discovery that can be attributed to a specific individual at a particular time. All of the brighter planets were known to the ancients. Neptune’s discovery was even more unusual. It was the first planet discovered with the aid of mathematical predictions. Prediscovery observations and 40 years of post discovery observations of Uranus showed that its motion was being perturbed from its expected elliptical orbital path. John Couch Adams of England and Urbain Jean Joseph Le Verrier of France were independently able to use these perturbations to infer the location of a previously unknown planet (Neptune). Le Verrier was the more successful in publishing his calculations and convincing an observer to look in his predicted direction (5). Following receipt of Le Verrier’s coordinates, John Galle and his assistant, Heinrich D’Arrest of Germany, first located Neptune on 23 September 1846 and confirmed its position and disc-shaped image the following night, more than a year after the first prediction had been made, but only five days after Le Verrier sent his request to Galle. Various prediscovery observations of Neptune were subsequently identified, the earliest by Galileo in 1612 (5)!

Formation According to current theories (6,7), the primordial nebula from which the solar system formed is comprised of 98% hydrogen and helium. In the part of the nebula where the outer planets formed, the remaining 2% was dominated by water, ammonia, methane, and rock, all of which were probably in condensed form in the vicinity of Uranus and Neptune. Most of the condensates were ‘‘ices,’’ a term applied to methane and ammonia, as well as to water, even when they are not actually in solid form. About a quarter of the solids were rock. This very small fraction of condensed solids played the key role of accreting into asteroid-sized planetesimals (of the order of a kilometer in size), which subsequently accreted into planetary cores. When these cores grew to a critical mass, they were then able to attract significant amounts of nebular gas (mostly hydrogen and helium). The process probably proceeded more slowly for Uranus and Neptune than for Jupiter and Saturn due to lower nebular densities and slower orbital velocities, and thus they were able to capture less nebular gas before the nebula was dissipated by the intense solar wind that developed during the Sun’s early evolution. Uranus and Neptune were thus formed by materials of a higher average density than the materials that formed Jupiter and Saturn. The rocky cores of Uranus and Neptune are each about one-sixth of the planet’s radius and are surrounded by icy material out to 75–80% of the radius; the outer 20–25% is primarily a hydrogen and helium gaseous envelope. During accretion of planetesimals, a planet will gain angular momentum as well as mass. A planet in an eccentric orbit will gain the greatest angular momentum at the edges of the accretion zone and will tend to accumulate prograde angular momentum (8). During the formation of both Uranus and Neptune, the last accreted objects might have been relatively large and had sufficient angular momentum to shift the spin axis significantly away from its ‘‘natural’’ spin direction, which is

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perpendicular to the ecliptic plane. The last large object that hit Uranus may have been the size of Earth and might have played a role in generating debris from which the satellites may have formed (8). The inclined angular momentum of the Neptune system is also a probable result of impacts; the largest impactor was between 0.1 and 0.5 Earth masses (9). When satellites form from debris created by a late large impactor, most of the debris material is from the impactor, which thus determines the satellite composition. The regular satellites are those that have nearly circular orbits close to the equatorial plane of the planet. These must have formed after the final large impact that shifted the angular momentum of the planet, and they might have formed from the debris generated by that impact. Any existing material in orbit near the planet at the time of the impact would be perturbed into inclined orbits that would result in collisions and breakup, which would then also contribute to accretion into satellite systems. Irregular satellites that have highly inclined, retrograde orbits are indicative of captured objects. Neptune’s largest satellite, Triton, is plausibly a captured object. It could have been captured during a close approach to Neptune if it had impacted a regular satellite of about 1% of its mass, which would then slow Triton sufficiently to keep it in orbit around Neptune (8).

Interior Structure The planets’ masses, shapes, rotational rates, and gravitational moments place constraints on their interior structures. Models of both planets can be constructed using three shells (10). Some models use an inner core of rocky material, an intermediate shell of icy material, and an outer layer of gas. The ratio of ice to rock in these models is about 15, and they require an atmosphere enhanced in volatiles by about a factor of 20 compared with solar fractions. Neptune’s rock–ice boundary in these models is about 15–20% of the planet’s radius. The ice– atmosphere boundary is about 80% of the planet’s radius. Successful models have also been created by using a gradual transition between the ice shell and the hydrogen-rich outer shell. Considerable uncertainty remains concerning the composition of the deep Neptune atmosphere and of the size of the rocky core (it might be about one Earth mass or considerably smaller). It is thought that the icy shells on both Uranus and Neptune are largely chemically homogeneous and consist of a mixture of ice, rock, hydrogen, and helium (10). Compared to Neptune, Uranus is somewhat less dense and somewhat more centrally condensed. As shown in Fig. 1, this structure has three main differences from that of Jupiter. Jupiter has a much larger gaseous molecular envelope, a large region of metallic hydrogen, and a smaller volume of ices. Neptune loses internal heat at a rate 30 times larger (in power/unit mass, or luminosity) than what could be provided by radioactive decay of trace elements (the Earth’s internal heat source). Neptune’s heat source, like Jupiter’s, is thought to be primordial heat left from the process of formation. The capture of solid and gaseous materials generated heat that raised the interior temperatures to very high levels (thousands K). That interior heat is still being released into space by Neptune, which emits 2.6 times as much heat as it absorbs from the Sun (12). This means that internal heat loss contributes 1.6 times as much as the

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molecular envelope gas

allic

met

He

H

ices rock

Figure 1. Approximate interior structures of Uranus and Neptune compared to Jupiter (11). Rocky and icy material may be mixed on Uranus and Neptune. The gas layer on Uranus is probably thicker than that on Neptune because Neptune is slightly denser. The radius at which metallic hydrogen is formed on Jupiter has been revised recently. Hubbard (7) now places it at about 0.8 RJ.

reradiated heat absorbed from the Sun. It is somewhat of a mystery that no comparable heat emission is measured for Uranus, which is so similar to Neptune in most other respects. The lack of an internal heat source explains why Uranus’ effective temperature (59 K) is the same as Neptune’s even though it is much closer to the Sun (where sunlight is about 2.5 times as intense). Estimates of the heats of formation for Uranus and Neptune and the estimated cooling since formation suggest that the remaining primordial heat is more than sufficient to explain the present luminosity of Neptune. One explanation for the low luminosity of Uranus is that there are compositional gradients in the interior of Uranus that lead to suppressed convection (10), which reduces heat transfer efficiency and thus the temperature of its outer atmospheric layers. Such gradients might have arisen in connection with the giant impact that is presumably responsible for Uranus’ spin axis inclination. An alternate theory (13) suggests that different external solar forcing produced by the high obliquity of Uranus has made the heat transfer process on Uranus much more efficient than on Neptune and resulted in a more rapid loss of internal heat. The generation of an offset dipole magnetic field, that has a significant quadrupole component requires the existence of a conducting fluid layer in convection. This layer might be bounded at the deepest level by the point below which the interior is stably stratified, and thus not convecting, and the point above which the fluid interior is not electrically conducting. If the entire interior

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fluid of Neptune exhibited the same angular rotation as that observed by clouds at the top of the atmosphere, then the J4 gravitational expansion coefficient would be positive, whereas the observed value is negative. This implies that the differential rotation of the upper atmosphere is a superficial effect that does not involve a significant fraction of the planet’s mass. However, this analysis does not place a well-defined boundary on how deep that flow could extend.

Atmospheres Neptune and Uranus have similar compositions, similar tropospheric temperature structures, and similar styles of zonal circulation, but different cloud patterns and great differences in weather activity. The basic atmospheric parameters for Uranus and Neptune are summarized in Table 2. Composition. Hydrogen, helium, and methane are the most prominent components of the atmospheres of Uranus and Neptune. Ammonia and probably H2S and water are present in layers below the region accessible by optical spectroscopy. Hydrocarbons of various types are observed in the upper atmosphere, a result of UV-induced chemistry involving the breakdown of CH4. The presence of hydrogen in any outer planet atmosphere was established first for Uranus, using the pressure induced 3-0 S(1) line of molecular hydrogen at 825.8 nm, which was first measured by Kuiper (15), but first identified by Herzberg (16). Molecular hydrogen occurs in two forms, the ortho form in which the nuclear spin vectors of the two atoms are parallel, and the para form in which the nuclear spin vectors are antiparallel. At high temperatures, the two forms reach an equilibrium concentration of three ortho molecules to one para molecule. That is called normal hydrogen. Hydrogen convected to lower temperatures will retain the normal mixing ratio for a long time because of the very weak interaction between the two nuclei in the absence of a catalyst. Normal hydrogen was expected on Jovian planets because of rapid mixing that should overwhelm the slow conversion process. With an effective catalyst and sufficient time, the two forms will reach an equilibrium distribution that depends on temperature, and with sufficiently rapid conversion, large effects on specific heat, atmospheric buoyancy, and temperature lapse rate can occur. When conversion is slow, the

Table 2. Atmospheric Compositiona Uranus Atmospheric composition (percent of total volume or molecular number) H2 (Sun ¼ 84) He (Sun ¼ 16) H2O (solar O ¼ 0.15) CH4 (solar C ¼ 0.07) NH3 (solar N ¼ 0.02) H2S (solar S ¼ 0.003) a

Ref. 14.

83 15 ? 2 (30  solar) ? ?

Neptune

79 18 ? 3 (40  solar) ? ?

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two forms behave like independent gases, and the specific heat at constant pressure (Cp) is lower. This leads to a steeper adiabatic lapse rate, given by dT/dz ¼ g/Cp, where g is the local gravitational acceleration. Observational constraints are conflicting (17). The hydrogen quadrupole spectrum and the collision-induced dipole spectrum measured in the infrared are both approximately consistent with thermal equilibrium for the two forms of hydrogen at the temperature at which the spectral lines are formed. But the measured temperature lapse rate is more consistent with that expected for normal hydrogen. This conflict might be resolved with the concept of stratified convection layers (18) in which a given gas parcel resides in a thin layer long enough to reach ortho–para equilibrium, even though the convective overturn time is relatively short. It is suggested that condensation of CH4 might produce a stepwise stratification of mean molecular weight that would increase stability, playing the same role as salinity variations in generating layered convection in terrestrial oceans. An alternative suggestion by Flasar (19) is that the lapse rate for P4700 mb is actually stable because of methane condensation that introduces a buoyancy gradient capable of supporting what otherwise would appear to be a superadiabatic (highly unstable) gradient for equilibrium hydrogen. Flasar suggests that the close agreement between the measured lapse rate and the adiabatic gradient for normal hydrogen is just a coincidence. The He/H2 ratio is constrained best from the analysis of far-infrared spectra. The collision-induced dipole absorption of H2-H2 and H2-He proves most of the continuum opacity at long wavelengths and is sufficiently well understood theoretically that it is possible to infer the ratio by fitting the thermal IR spectra that are also consistent with temperature profiles determined by Voyager 2 radio occultation measurements. The inferred helium mole fraction (number or volume fraction rather than mass fraction) in the upper atmosphere of Uranus is 15.2%73.3% (20). Neptune’s is a little higher at 19.0%73.2% (21). As expected from formation theories, both are within errors of the solar value of 16%. This contrasts with Saturn, which has only about one-fourth of the solar fraction, presumably due to rainout of He in Saturn’s deep interior where it is thought that He becomes insoluble in metallic hydrogen. Spectroscopic observations indicate that methane (CH4) is enhanced relative to solar abundance values by a factor of 25–30 in regions where it ought to be well mixed (22). However, because methane condenses to form clouds in both atmospheres, methane abundance above the condensation level can be greatly reduced, variable, and difficult to estimate. Initially, it was thought that the tropopause would act as a cold trap, limiting the stratospheric methane mixing ratio to values no greater than the saturation mixing ratio at the tropopause. Yet Voyager Ultra-Violet Spectrometer data for Neptune imply that stratospheric mixing ratios are at least 10 times this limit (23). This excess, termed oversaturation, suggests stronger vertical mixing on Neptune compared to Uranus, where stratospheric oversaturation is not observed. The large tropospheric enhancement of methane relative to the solar mixing ratio for both Uranus and Neptune is a likely consequence of the planetary formation process. The large fraction of icy materials accreted by these planets should also have resulted in similar enhancements of water and ammonia, even though little has so far been observed in the atmospheres.

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Microwave atmospheric observations imply a significant depletion of NH3 relative to solar values, by a factor of 100–200 in the 150–200 K region of Uranus’ atmosphere (24). A possible explanation is that NH3 is lost to an extensive NH3– H2O solution cloud or that NH3 is lost to the formation of a cloud of NH4SH (ammonium hydrosulfide). The depletion might be completely accomplished by formation of an NH4SH cloud if the H2S/NH3 ratio is enhanced by a factor of 4 compared to the solar value of 0.2. It seems to require very large enhancement factors of water for the NH3–H2O solution cloud to play a major role in the depletion of NH3. The necessary enhancement of H2S might have resulted from accretion of chondritic meteorites, in which sulfide minerals are found but nitrogen incorporation is not significant. An alternative hypothesis by Lewis and Prinn (25) is that Uranus never acquired much nitrogen because uncondensed N2 and CO were the dominant chemical forms of N and C in the solar nebula (rather than NH3 and CH4) in the region of the terrestrial planets and also, as a result of rapid mixing in the outer parts of the nebula where thermal equilibrium compounds would be NH3 and CH4. So far, there is no direct observation of H2O or NH3 on either Uranus or Neptune, nor of H2S on any Jovian planet except Jupiter. If H2O is present at solar mixing ratios, condensation clouds might be found at pressures of the order of 100 bars. If enhanced 65 times solar, water would condense at a temperature of 647 K, but would not condense at all beyond that enhancement (24). Spectral Characteristics of Uranus and Neptune. Methane absorption dominates the visible and near-IR disk-averaged spectra of both Uranus and Neptune (upper part of Fig. 2). At wavelengths below 1 micron, there are numerous methane bands of increasingly greater absorption from the green to the red and near IR. The absorption of red light by methane contributes to the blue colors of Uranus and Neptune. Neptune is bluer than Uranus because of increased absorption in the window regions between the methane bands. This might mean that the visible cloud layer that controls the amount of light reflected backward is thicker and brighter on Uranus, whereas the corresponding cloud layer on Neptune is more transparent and allows more of the light to be absorbed by the deeper atmosphere. Alternatively, the cloud itself might provide the extra absorption needed on Neptune (22); if so, the cloud must absorb even more effectively at wavelengths between 1 and 2.5 microns (29). The basic disk-averaged characteristics of Neptune and Uranus at near-IR wavelengths are illustrated in the upper right panel of Fig. 2, which displays the ground-based observations of Fink and Larsen (27). The relatively low diskintegrated albedos of both planets are due to the significant amount of methane overlying the visible cloud layer. The albedo peaks that occur in windows of relatively weak absorption, for example, 1.3 and 1.6 mm, are narrower for Uranus than for Neptune, indicating that Uranus has a clearer atmosphere with fewer high altitude hazes to reflect photons that would otherwise be absorbed by underlying methane. In the 1–1.8 mm region methane is the dominant absorber, whereas hydrogen collision-induced absorption (CIA) is the major absorber in the 1.85–2.2 mm range. The effects of these absorbers on the penetration depths of photons into the atmospheres of these planets are illustrated in the middle panel of Fig. 2. This displays the wavelength-dependent pressure levels at which a unit albedo re-

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m

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Figure 2. (Top) Reflectivity spectra of Nepture (26–28). (Middle) Penetration depth of photons into Neptune’s atmosphere (30). (Bottom) Appearance of Neptune as a function of wavelength as recorded by HST imaging of WFPC2 and NICMOS (30).

flecting layer would have apparent albedos of 37%, 13.5%, and 1.8% (e  1, e  2, and e  4) (30). This spectrum of sensing depths was computed for Neptune using a tropospheric methane mixing ratio of 0.022 and a stratospheric mixing ratio of 3.5  10  4 (22), the saturation mixing ratio is assumed wherever it is less than either of the others. (Penetration depths for Uranus are roughly the same as those for Neptune.) In Fig. 2, we see that the 1.27 mm and 1.59-mm windows probe all the way into the putative H2S visible cloud deck near 3.2 bars. Were that cloud composed of high-albedo particles, we would expect albedo values near unity, rather than values in the 0.05–0.08 range. On the basis of the near-IR spectrum alone, the cloud is very dark, more transparent, or deeper than indicated by other methods. The exact levels sensed in the strongly absorbing 2.3-mm methane band depend on stratospheric methane mixing ratios, which are not well constrained by current observations. Sample images of Neptune obtained by the Hubble Space Telescope Wide Field/Planetary Camera 2 and Near Infrared

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Camera and Multi-Object Spectrometer are shown in the bottom panel of Fig. 2 (30). These illustrate the dominance of Rayleigh scattering at short wavelengths and the effects of methane absorption at long wavelengths. Neptune has both dark and bright discrete features, but the only discrete features on Uranus seem to be bright features. Dark features on Neptune have the greatest contrast at blue wavelengths (see lower left of Fig. 2), but it is typically a rather small 2–10%. The identity of the blue absorber that creates this contrast is unknown. The contrast between bright clouds and the background atmosphere can also be rather small at short wavelengths, where Rayleigh scattering is important, but it is dramatically enhanced at wavelengths where methane absorption is strong. On Neptune, the contrast can exceed 500:1 at 2 microns (30). The maximum contrast for bright clouds on Uranus is seen at about 1.9 microns and is about 1.8:1 in raw images, but estimated at about 10:1 at high spatial resolution (31). Relatively less contrast is observed on Uranus because its bright cloud features do not reach to pressures as low as those on Neptune. Temperature Structure. The temperature structures of Uranus and Neptune are very similar, as illustrated in Fig. 3 by dotted and solid curves, respectively. There is remarkably little latitudinal variation in this structure. Given the 981 obliquity (pole inclination relative to orbital normal) of Uranus and Neptune’s 291, the absence of significant pole-to-pole gradients (for Uranus) or equator to pole gradients (for Neptune) implies some heat transport or compensation effect. The effective emission temperatures on both planets are very low (about 59 K), leading to very long radiative time constants of about 5  109 seconds at 400– 500 mb (32). This is the ratio of the thermal energy content of an atmosphere to

0.001

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=? CH4 BRT clouds ? H2S (?)

10.000 50 100 150 200 Temperature, K, Neptune cloud structure

Uranus cloud structure

Figure 3. Vertical structure of temperatures and cloud and haze layers. The Neptune temperature profile (solid) differs from the Uranus profile (dotted) mainly in the stratosphere where a larger population of hydrocarbon hazes on Neptune leads to more heat absorption and a warmer stratosphere.

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the radiative cooling rate for one local scale height. This radiative time constant is twice as long as Uranus’ orbital period. Thus, seasonal variations on Uranus are strongly damped and phase shifted (delayed) by close to 1/4 year. In fact, even though the North Pole (IAU convention) of Uranus had been in darkness for 20 years, the Voyager IRIS instrument found no measurable difference in temperature structure between the two polar regions (33). Voyager 2 arrived close to the Southern Hemisphere solstice, a time when hemispheric thermal contrast should be at a minimum because of the phase lag, and thus it was not able to measure the amplitude of the seasonal response. Because of Uranus’ obliquity, the average solar heating for an entire Uranian year is greater at the poles than at the equator. However, seasonal model predictions of a somewhat cooler equator were not confirmed by Voyager observations. The main latitudinal variation in temperature on both Uranus and Neptune is consistent with the thermal wind equation. This equation is a proportionality between vertical wind shear and horizontal temperature gradients, which is valid for vertical hydrostatic balance (between pressure and gravity) and horizontal balance between coriolis forces and horizontal pressure gradients. The sign of the derived vertical wind shear indicates that the zonal winds decay as height increases (discussed later). Cloud Structure. Our current limited understanding of cloud structures on Uranus and Neptune is illustrated in Fig. 3. The saturation vapor pressure curve for a CH4 mixing ratio of 2% suggests a methane cloud base near 1.4 bars on Neptune and 1.2 bars on Uranus, although nucleation at somewhat higher altitudes is likely because it would probably require some degree of supersaturation (34). Radio occultation observations of refractivity gradients in this region agree with the 1.2 bars on Uranus (35) but suggest a cloud base near 1.9 bars on Neptune (36). Neptune’s stratospheric hazes (34,37,38), which significantly reduce Neptune’s shortwave albedo, have a relatively low total equatorial optical depth of B0.02 at 0.75 mm (39). Larger optical depths found for the global average might be due to contributions from unresolved high-altitude, isolated, bright methane clouds. The optical depth of the methane cloud at red wavelengths is relatively small; on Neptune, it ranges from about 0.05 near the equator to 0.3 near 251 S (22), and on Uranus from 0.4 (40) to 1.3–1.5 (41,42); there is evidence for lower values (several tenths) for central disk observations, and higher values are inferred from disk-integrated values (43). Much larger opacities are possible for discrete features. The presence of an H2S cloud is inferred from microwave spectra that probe the deep atmosphere; enhancement by a factor of 30 relative to the solar mixing ratio is not directly measured but is inferred to explain the very low abundance of NH3 (44,45). In this scenario, the formation of a very deep NH4SH cloud consumes the excess NH3. When the corresponding H2S condensation curve is computed, we find that it intersects the temperature profile in the 6–9 bar region, above which we might expect to see a cloud of H2S. The presence of an opaque cloud at pressures deeper than 3.6 to 3.8 bars is inferred from hydrogen quadrupole line widths (40). However, attempts to find direct evidence of H2S in optical spectra have failed (46). Thus, we cannot completely rule out the possibility that both NH3 and H2S are very depleted, as advocated by Romani et al. (37) and that the visible cloud deck is a thin ammonia ice cloud or that no cloud at all is present in the 3–10 bar region, though that would conflict with the quadrupole results. An ammonia cloud, if present, might

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also form near the 7-bar level. Current models also have difficulty explaining thermal infrared observations of Neptune, which seem to require that the methane cloud be much more opaque at a significantly higher altitude or horizontally heterogeneous (21). Uranus is clearer than Neptune, perhaps because it lacks internal heat to drive the mixing needed to keep particles suspended for long times. The global average methane haze layer on Neptune is relatively thin (B0.1), but the optical depths of discrete bright cloud features are probably much greater. On Neptune, these features reach 50–100 km above the mean cloud level, reaching to pressures of 100–200 mbar. On Uranus, discrete bright clouds do not rise much above 500 mb (31). Horizontal cloud structures for Uranus and Neptune are illustrated by sample Voyager and HST images in Figs. 4–7. Voyager imaging of Uranus (Fig. 4) in 1986 displayed a rather bland appearance. An approximate true-color image (left image in Fig. 4) displays neither banding nor discrete cloud features. The extremely enhanced false-color image in the middle of Fig. 4 shows that there was a latitudinal variation in the amount of UV-absorbing haze in the atmosphere; the greatest absorption occurred near the visible pole of Uranus (IAU south) where it appears as relatively orange. Discrete cloud features are illustrated in the time sequence of images at the right of Fig. 4, which were made using Voyager’s orange filter. Hubble Space Telescope images of Uranus made in 1997, 11 years after the Voyager encounter, reveal the first discrete cloud features in its Northern Hemisphere. One feature is barely visible near the right-hand limb in the upper left

V2 orange V2 green V2 blue

V2 orange V2 violet

V2 UV

Figure 4. [JPL P29478]. Voyager 2 images of Uranus taken on 17 January 1986 using blue, green, and orange filters to make the true-color composite (left), which displays a virtually featureless disk. UV, violet, and orange filtered images were shown as blue, green, and red components in the extremely enhanced false-color image (middle), which reveals polar bands of UV-absorbing haze particles, centered on the South Pole of Uranus (IAU convention). Even in this view, no discrete cloud features are apparent. [Voyager 2 JPL P29467.] The right-hand time sequence of orange-filtered images from 14 January 1986 shows the motion of two small bright streaky clouds that were the first discrete features ever seen on Uranus. Uranus is rotating counterclockwise in this view, as are the clouds, though more slowly than Uranus’ interior, revealing that low latitude winds on Uranus are retrograde, as are the winds on Neptune. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

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image, which was made with a 619-nm filter (the wavelength of a weak methane band). Much better contrast between discrete cloud features and the background atmosphere was obtained at near-IR wavelengths using the HST NICMOS camera, as illustrated in the middle and right-hand images of Fig. 5.

Belinda Rosalind

Puck

Epsilon ring Portia

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Desdemona 547 nm

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Cressida

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Figure 5. (Left) [STScI PRC97-36b, NASA and H. Hammel]. HST WFPC2 images of Uranus on 31 July and 1 August 1997. Although little contrast is seen at 547 nm (blue), a banded structure and the first discrete Northern Hemisphere cloud are visible at 619 nm (upper image, colored red). (Middle) [Space Telescope Science Institute STSCI-PRC97-36A and E. Karkoschka]. This false-color 1997 image is a composite of near-IR images taken by the Near Infrared Camera and Multi-Object Spectrometer (NICMOS) at wavelengths of 1.1 mm (shown as blue), 1.6 mm (shown as green), and 1.9 mm (shown as red). Absorption by methane gas limits the depth at which reflected sunlight can still be seen at 1.1 and 1.6 mm, and absorption by hydrogen is most significant at 1.9 mm. The blue exposure probes atmospheric levels down to a few bars, responding to scattering by aerosols and by atmospheric molecules above this level. The green component is least sensitive to methane absorption, sensing down to 10 bars or more, but sees much less Raleigh scattering per bar than the blue component, so that a dark absorbing cloud near 3 bars would result in more blue than green in regions that were clear above the cloud, perhaps accounting for the blue color at midlatitudes. The red component can only sense down to about the 2-bar level, and sees the least contribution from Rayleigh scattering, so that very little red is seen in regions that are clear to 3 bars. The green color around the South Pole suggests significant local haze opacity at pressures near 2–3 bars. The red color of the discrete features near the northern (right) limb indicate relatively high-altitude clouds that reflect sunlight before much absorption has taken place. The curved arcs in the central image indicate motions in 90 minutes of cloud features and eight of the 10 small satellites discovered by Voyager 2. The area outside the rings was enhanced to make the satellites more visible. The images also show the bright epsilon ring, which is wider and brighter in the upper part of the image, and two fainter inner rings. (Right) [STScI-PR98-35, NASA and E. Karkoschka]. This Hubble Space Telescope near-IR image of Uranus on 8 August 1998 shows a number of new cloud features in both hemispheres. The false-color image was created in a manner similar to the first, except that the rings and satellites were not separately brightened. This figure is available in full color at http://www.mrw.interscience. wiley.com/esst.

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Voyager 2 images of Neptune in 1989 (Fig. 6) provided a rich bounty of detail concerning horizontal cloud structure and revealed many discrete bright and dark cloud features, as well as bright and dark bands. Many of these structures were unique to Neptune and have not been seen before or since. HST imaging of Neptune beginning in 1994 showed some cloud banding similar to the 1989 Voyager images, but many changes were found, including the disappearance of Neptune’s Great Dark Spot (discussed later).

Weather Phenomena on Neptune Dark Spots. Voyager identified two prominent dark oval features in Neptune’s Southern Hemisphere; both are thought to be manifestations of anticyclonic eddies, as illustrated in Figs. 6 and 7. (Anticyclones rotate counterclockwise in the Southern Hemisphere.) The Great Dark Spot (GDS) was about an Earth diameter long. It was seen for about 8 months during Voyager’s approach to Neptune but has not been seen since. It exhibited several unusual dynamic features; some of them are illustrated in the middle panel of Fig. 6. Its

Figure 6. (Left) [JPL P34606]. Voyager 2 image of Neptune. Color composite formed from green, blue, and red filtered images taken on 18 August 1989. South is down, and the South Pole is tipped toward Earth. The Great Dark Spot (GDS) seen at the center of the image is about 13,000 km (about the diameter of Earth) by 6,600 km, and at a latitude of 181S at the time of this image. The bright cloud to the south of the GDS, termed the companion, is at relatively high altitude compared to the blue features, and was very prominent in ground-based images made at methane-band wavelengths (51). The bright clouds at the edge of the dark circumpolar band, called south polar features (SPF), are at a latitude of about 671S; these are highly variable on short timescales and exhibit vertical relief of about 75–150 km, as inferred from shadows seen in Voyager images. (Right) [JPL P34668]. This image of Neptune’s south polar regions near 681S, made on 23 August 1989, shows the first cloud shadows ever recorded by Voyager on any planet (the Sun is to the left). (Middle) [JPL P34610]. This time sequence of remapped images of Neptune’s GDS, from top to bottom, provides views at intervals of about 18 hours, revealing its strange wobble and shape changes, which occur in a period of 193 hours. Its mean dimensions were approximately 381 in longitude and 141 in latitude and had modulation amplitudes of 7.4 and 1.51, respectively (47). It also drifted toward the equator at a rate of about 11 per month. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

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Figure 7. (Left) [Space Telescope Science Institute PRC-98-34 and H. Hammel]. HST composite color image made from images at blue, orange, and near-IR wavelengths and corrected for limb darkening. (Middle) [STScI and L. Sromovsky]. This quartet of falsecolor images made using HST shows persistent banded structure and an increase in the number of bright clouds between 1996 and 1998. (Right) [STScI and L. Sromovsky]. Views of dark spots seen by Voyager (GDS in top row) and HST (NGDS-32 and NGDS-15 in bottom row). The HST images (48) are composites of several images taken at slightly different times. This figure is available in full color at http://www.mrw.interscience. wiley.com/esst.

shape and orientation varied cyclically in a period of 193 hours (47). The longitudinal width varied from 31 to 45o, and its latitudinal extent varied from 12.5 to 15.51. This behavior was successfully reproduced using a Kiva vortex model (49). This model predicts anticyclonic circulation (counterclockwise in the Southern Hemisphere), although no direct measurements have been able to test this prediction. The GDS also drifted toward the equator at a steady rate of 1.241/ month (47); this would have placed it on the equator in November 1990, although theoretical models imply that it would have dissipated before that point (50). This behavior is in marked contrast to that of Jupiter’s Great Red Spot, which has remained at essentially a fixed latitude for centuries. The GDS was accompanied by a prominent bright companion cloud positioned at its southern boundary (Figs. 6 and 7). In ground-based methane band imaging at the time of the Voyager encounter, the companion was the brightest feature on the planet (51) and seems to be similar to orographic clouds. It is likely caused by vertical deflection of flow around the GDS, producing methane condensation during the upward part of the flow and evaporation on the downward return leg (52). The first Hubble Space Telescope images that had sufficient image quality to show the subtle contrast of the GDS at blue wavelengths and sufficient longitudinal coverage were made in 1994, after the Hubble repair mission. The GDS was not seen at that time, and no other southern dark spot has been seen since. The second prominent dark spot observed by Voyager (DS2 in Fig. 7) was much smaller than the GDS (about 201 in longitude by 61 in latitude), and had several unique dynamic characteristics of its own (47). Its latitudinal position oscillated at an amplitude of 2.41 about its mean latitude of 52.51 S in a period of

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865 hours. During this oscillation, its zonal speed increased in step with the zonal wind shear, resulting in a 481 longitudinal oscillation amplitude. Bright clouds formed near the center of the dark spot during the northern half of the oscillation and declined during the southern half. Strangest of all, its mean drift rate relative to Neptune’s interior matched to great precision the mean drift rate of the south polar features (discussed later); the two mean positions remained on opposite sides of Neptune and at mean latitudes differing by 201. The DS2 feature has not been seen since the Voyager encounter. Two new dark spots were discovered in HST images (Fig. 7); both were in the Northern Hemisphere. NGDS-32 was discovered first in 1994 images (53) at about 321 N latitude, and a bright companion cloud was seen at its equatorward edge. This dark spot is comparable in size to the GDS but is harder to observe because the latitude circle where it is found is close to the northern limb. NGDS15 was discovered first in 1996 HST images (54) close to latitude 151 N. It is somewhat smaller than NGDS-32 and spans about 201 in longitude and 111 in latitude. It is unusual in having no bright companion cloud. NGDS-15 is also unusual in being close to the latitude at which modeled dark spots tend to dissipate rapidly. Both NGDS-32 and NGDS-15 differ from the Voyager GDS by exhibiting no measurable latitudinal drift, a characteristic that is baffling to modelers because model dark spots always drift equatorward on Neptune. The modeled drift rate depends on the latitudinal gradient in the zonal wind speed. If that gradient is locally perturbed in the vicinity of the dark spot latitudes, it may be possible to get consistency between model and measured behavior. South Polar Features. Rapidly varying small bright features clustered near 62–701 latitude, termed south polar features, were the first clouds Voyager observed that had shadows (see Fig. 6). These features were observed almost exclusively in a region less than 1801 wide, and though the mean position where they were seen drifted only slightly relative to the interior, the individual cloud elements actually moved relatively quickly at a prograde rate of about 200 m/s. Their lifetimes were so short that they could not be tracked during a full rotation of Neptune. At low resolution, the features blend together so that they sometimes appear to form a plume (the southernmost prominent bright cloud in the left image of Fig. 6). But in high-resolution images, they can be seen as small individual elements (see the right image of Fig. 6). The vertical relief implied by shadows is 50–150 km (51). Waves and Bands. Cloud patterns on Neptune suggest the existence of atmospheric wave motions. Voyager imaging showed that near 201 S latitude, there was a clustering of bright cloud features at two regions spaced approximately 1801 apart in longitude; one region was in the vicinity of the GDS. This suggests a wave that has two complete oscillations within 3601 of longitude (wave number 2). Neither the pattern nor the GDS have been seen since the Voyager encounter. The phase-locked motions of the region of SPF formation and DS2 suggest a similar wave interaction. A more obvious wave example is the dark band between 551 S and 651 S, visible in blue-filtered images (right group in Fig. 7). When viewed in polar projection, the band appears as a circle offset from Neptune’s South Pole, leading to a 21 sinusoidal modulation in latitude of the wave boundaries as a function of longitude. The wave can be seen as a difference in tilts of the band in the two images shown for August 1998 in the middle group

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of Fig. 7 where the band appears greenish because of the false-color scheme. Although DS2, which was nestled in one of the northern excursions of the wave in 1989, is no longer visible (at least in HST images), the wave structure appears to have persisted from 1989 through at least 1998. Secular Variations on Uranus and Neptune. In recent years, Uranus has revealed greater weather activity as its Northern Hemisphere continued its emergence out of decades of darkness. First seen in HST images of Uranus (31,55), bright Northern Hemisphere clouds have now become visible from the ground (56). Over a longer time period, Uranus has exhibited rather strong diskaveraged albedo variations, including a 14% increase from 1963 to 1981 (during which the sub-Earth latitude varied from near zero to 681 S) (57). Additional ground-based observations of a peak blue reflectivity in 1985 and a 7% decline from 1985 to 2001 are reported by Karkoschka (58), who also showed that there is a hemispheric asymmetry in brightness; the Northern Hemisphere is darker, and much of the recent decline in brightness has to do with geometric effects as more of the darker Northern Hemisphere comes into view. He also showed that additional physical effects played a role in earlier brightness increases and suggested that significant physical effects would also appear in the near future. Recently, Neptune also exhibited a relatively steady brightening at visible wavelengths of 1.4–1.9% from 1996 to 1998 (30) and almost 10% from 1990 to 2000 (59); this dramatically breaks from the previous inverse correlation with solar UV variations. Neptune’s atmosphere also exhibits dynamic activity during decade-long timescales, as established by ground-based imaging of changing distributions of bright features and ground-based photometry of varying light curves. During 1976 and 1987, it appears that much brighter cloud features were present than have been seen during or since Voyager (60). Ground-based observations of bright cloud features at latitudes where none were seen by Voyager, large changes in light curve amplitudes from year to year (61), and factor of 10 changes in near-IR brightness (62) also suggest major developments or large latitudinal excursions during a several year period. Sromovsky et al. (29) showed that the spectral character of the 1977 ‘‘outburst’’ could be matched by a factor of 7 increase relative to August 1996 in the fractional area covered by high-altitude, bright, cloud features. Zonal Mean Circulation and Its Stability. Uranus and Neptune provide strong evidence that rotation dominates solar radiation and internal heat flux in determining the form of a planet’s zonal circulation. Given that Uranus has had one pole in sunlight for B20 years, it would have been natural to expect strong meridional circulation between the two hemispheres. Yet, the observed temperature difference between Uranus’ long dark hemisphere and that heated by the Sun is very small and opposite in sign to this expectation. Part of the explanation was already discussed in the section on thermal structure, that is, the long radiative time constant removes most of the interhemispheric thermal contrast. The form of the circulation turned out to be zonal (parallel to lines of constant latitude). Voyager’s inability to find many cloud features on Uranus in 1989 hampered efforts to characterize fully its mean zonal circulation. Yet, even with sparse sampling, it was clear that the circulation was retrograde near the equator (atmospheric parcels fell behind the planet’s rotation) and prograde at high

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latitudes. Radio occultation measurements provided one low latitude measurement (Fig. 8) that was critical in confirming this picture. Uranus’ cloud features have remained difficult to observe since that time, until recently. Improved nearIR imaging capabilities (NICMOS and Keck adaptive optics imaging) and the emergence of the Northern (IAU convention) Hemisphere into daylight have brought many new cloud features into view. New measurements, summarized by Hammel et al. (53) and shown in Fig. 8, suggest a possible small difference between hemispheres. To first order, however, the new data are consistent with the Voyager results and suggest a relatively stable and symmetrical circulation that bears a strong resemblance to Neptune’s in form but is considerably weaker in amplitude. The most detailed definition of Neptune’s zonal mean circulation is provided by Voyager 2 imaging observations (64). A strong retrograde equatorial jet at 400 m/s (895 mph) and a weaker prograde jet at 250 m/s are the main features of the circulation (Fig. 8). The prograde jet on Uranus moves at about 200 m/s and at a lower latitude (601 vs. 70–751 for Neptune). The Uranian equatorial jet is much weaker than Neptune’s (about 100 m/s retrograde vs. 400 m/s for Neptune) and covers a narrower latitude range (7251 vs. 7501 for Neptune).

Planetographic latitude, °

Uranus

Neptune

50

50

0

0

−50

−50 Voyager Post-voyager

100 200 300 −600 −200 −100 0 Eastward wind speed, m/s

−400

0 200 −200 Eastward wind speed, m/s

400

Figure 8. Zonal mean circulations of Uranus and Neptune. Voyager results (filled circles) and post-voyager results (open circles) are roughly consistent with an unchanging symmetrical circulation for both planets, although somewhat better fits to the Uranus observations are obtained with slightly asymmetrical profiles. Data points are from the Uranus compilation by Hammel et al. (63), and the Neptune Voyager observations of Limaye and Sromovsky (64) and HST observations by Sromovsky et al. (30). The postVoyager results for Uranus are a combination of HST and Keck imaging from 1997–2000, the post-Voyager Neptune results are entirely from HST imaging from 1994–1998.

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From the wind measurements displayed in Fig. 8, it seems that there is much less variability in the measurements of Uranus. This is in part the result of very rapid evolution of cloud features on Neptune, which reduces the typical time interval over which a cloud feature can be tracked and thus reduces the accuracy in speed that can be achieved. It is also true that part of the variability in the Neptune measurements comes from true eddy motions. The cloud features tracked on Uranus, on the other hand, seem to be more stable, allowing longer observation periods. This might be so because the cloud features are less numerous and less visible on Uranus, so that only the more intense and long-lasting features can be seen at all. It is also likely that due to far lower internal energy flux on Uranus, there may be less eddy activity to start with, as indicated by the small number of high altitude clouds and the relatively clear upper troposphere. The equatorial retrograde winds on both planets might be a consequence of angular momentum conservation in the presence of an axisymmetric meridional circulation. Because distance to the spin axis decreases with latitude, so does the angular momentum of atmospheric parcels at rest with respect to the rotating planet. But if atmospheric mass at midlatitudes were to move equatorward, it would tend to slow its rotational speed to conserve its angular momentum, leading to a retrograde wind near the equator at the level of the equatorward flow. Midlatitude gas parcels moving poleward would tend to speed up as they get closer to the spin access, leading to a prograde circulation at high latitudes. The reverse flow at return levels would tend to produce the opposite effects. This meridional flow may be consistent with the observed zonal circulations, but there are no direct observations that could confirm or deny the existence of the meridional circulation. Such a model would not explain the equatorial prograde jets of Jupiter and Saturn that clearly require transport of angular momentum by eddies. The solar irradiance at Uranus is about 3.8 W/m2, and the planet-wide average is about 0.65 W/m2 in absorbed flux. Neptune, on the other hand, is exposed to a solar irradiance of only 1.5 W/m2 and absorbs an average of about 0.26 W/m2 compared to its internal heat flux of 0.4 W/m2 (65). Thus, the two planets actually have about the same total heat flux available for generating atmospheric motions. Their different atmospheric motions might be due to differences in the latitudinal distributions of the fluxes. Based on its modest spin axis inclination, Neptune’s equator will receive much more solar heat than its poles, even when averaged for a year, whereas the 981 inclination of Uranus results in its poles receiving 50% greater solar input than the equator. Based on calculations for Saturn, which has almost the same obliquity as Neptune, latitudes poleward of 601 should receive only half of the average solar flux received near the equator. Despite these different heat input distributions, there is little latitudinal variation in atmospheric temperatures on either planet, as noted in the discussion of thermal structure. To equilibrate the temperatures across all latitudes would require different magnitudes of horizontal heat transport and (even different directions) on Uranus and Neptune. This may account for some differences in circulation, but the point made by Ingersoll (66) is worth noting: there is not much connection between energy sources and speeds and patterns of outer planet circulations. Jupiter has 20 times the available power to drive circulations, yet it has only one-third the wind speeds. Although baroclinic eddies might

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provide the needed horizontal heat transport on Uranus, Friedson and Ingersoll (67) suggest that the internal heat flux on Neptune might reduce latitudinal gradients, acting as a sort of thermostatic control, efficiently providing extra heat in regions that are slightly cooler and reducing heat in regions that are slightly warmer. For this mechanism to work, the solar energy absorption must occur in regions within or close to the free convection zone that extends into the interior. But how can the high winds of Uranus and Neptune and Neptune’s highly variable weather phenomena be maintained? The answer seems to be that these atmospheres have very low dissipation and thus take very little power to keep them running, much like a well-lubricated ball bearing. We don’t know very much about the winds below or above the level of the visible cloud features. We can use Voyager 2 measurements of horizontal temperature gradients to estimate the vertical wind shear by using the thermal wind equation discussed earlier. We find that the winds on both Uranus and Neptune decay as height increases, suggesting frictional dissipation in the stratosphere (33,66). The vertical shear is relatively low and requires about 10 scale heights (the pressure drops by a factor of 1/e for each scale height) to damp to zero velocity. To what depth below the clouds the winds continue to increase is unknown. We do know from studies of Neptune’s gravity field (discussed earlier) that its winds cannot maintain the same speed throughout the interior but must damp to low values within a small fraction of the planet’s radius. The Voyager measurements of Neptune’s zonal circulation (Fig. 8) are at least roughly consistent with earlier ground-based and subsequent HST observations (53,54). However, it is not clear that this circulation is stable in detail, or whether there is more latitudinal variation than Voyager observations have indicated. According to the theory of LeBeau and Dowling (50), the detailed curvature of the zonal wind latitudinal profile is important because it determines the drift rate and lifetime of Great Dark Spots. Thus, measurements of both zonal wind and discrete feature latitudinal drifts put strong constraints on such theories. Recent HST observations (30) are beginning to indicate a consistent pattern of small deviations from the Voyager profile.

Satellite Systems Uranus and Neptune, like Jupiter and Saturn, have systems of regular satellites in prograde orbits that lie near the planets’ equatorial planes. The two planets also have irregular satellites. Neptune’s moon Triton is the most significant. Because of their orbital similarities and prograde orbits, regular satellites, it is thought, formed with the planet in a common process, rather than being captured after the planet’s formation. During the early stages of giant planet formation, the local environment becomes hot enough to vaporize the constituents that later cool to form the solids, which subsequently accrete into satellites. The thermal history of each planet and the timing relative to the blowing out of nebular gas by the intense solar wind generated during the Sun’s T Tauri phase are thought to determine the characteristics of the satellite systems. Satellites of Uranus. Uranus has 20 known satellites; five were known prior to Voyager’s 1986 encounter with the Uranian system (Fig. 9), and 11 were

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Miranda

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Ariel

Umbriel

Titania

817

Oberon

Figure 9. [JPL P30054]. Voyager 2 montage of the five largest satellites of Uranus in order of increasing distance from Uranus (Miranda, Ariel, Umbriel, Titania, Oberon) and correct relative sizes and brightness. Similar to Fig. 15. p. 52, Science 233. Imaging Team Report. This figure is available in full color at http://www.mrw.interscience.wiley.com/esst.

discovered in Voyager images (Table 3). The names of the Uranian satellites are derived from the writings of Shakespeare and Pope. The four largest regular satellites (Ariel, Umbriel, Titania, and Oberon) have low inclination and low eccentricity. Most formation models require that the satellites formed after the Table 3. Satellites of Uranus Satellite

Cordelia Ophelia Bianca Cressida Desdemona Juliet Portia Rosalind Belinda 1986U10 Puck Miranda Ariel Umbriel Titania Oberon Caliban Stephano Sycorax Prospero Setebos a

Orbital radius (km) 49,750 53,760 59,160 61,770 62,660 64,360 66,100 69,930 75,260 76,000 86,000 129,780 191,240 265,970 435,840 582,600 7,169,000 7,948,000 12,213,000 16,568,000 17,681,000

Radius Mass Sidereal Densitya Orbital Geomet- Discovery (km)a (1021 g) period ric alinclinadatec b a tion, deg bedo (days) 20 21 26 40 32 47 68 37 40 40 77 236 579 585 789 761 49 10 95 15 15

? ? ? ? ? ? ? ? ? ? ? 0.063 1.27 1.27 3.49 3.03 ? ? ? ? ?

0.335 0.376 0.435 0.464 0.474 0.493 0.513 0.558 0.624

1.3 1.3 1.3 1.3 1.3 1.3 1.3 1.3 1.3

0.762 1.414 2.520 4.144 8.706 13.463

1.3 1.21 1.67 1.40 1.72 1.63 1.5 1.5 1.5 1.5 1.5

0.08 0.10 0.19 0.01 0.11 0.07 0.06 0.28 0.03

0.07 0.07 0.07 0.07 0.07 0.07 0.07 0.07 0.07

0.32 4.22 0.31 0.36 0.10 0.10 139.68

0.07 0.32 0.39 0.21 0.27 0.23 0.07 0.07 0.07 0.07 0.07

152.67

1986a 1986a 1986a 1986a 1986a 1986a 1986a 1986a 1986a 1999b 1985a 1948c 1851d 1851d 1787e 1787e 1997f 1999f 1997g 1999h 1999i

Ref. 68: From ssd.jpl compilation on 24 April 2002. Ref. 69. c Discoveries are dated by data acquisition date (a ¼ Voyager 2, b ¼ Karkoschka, c ¼ Kuiper, d ¼ Lassell, e ¼ Herschel, f ¼ Gladman, g ¼ Nicholson, h ¼ Holman, i ¼ Kayelaars). b

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impact event that tilted Uranus’ spin axis and that they evolved to a state of synchronous rotation in which the satellite always keeps the same side facing the planet. Synchronous rotation has been approximately confirmed for all five of the largest satellites by Voyager 2 imaging (70). These satellites are generally denser and darker than the moons around Saturn and have a rocky fraction of 50% or more. One theory is that a giant impact generated a disk that was ice-poor because the shock energy converted methane and ammonia into largely uncondensable CO and N2. An alternative theory is that the nebular gas in the vicinity of Uranus was more primitive and unprocessed, so that C and N were more tied up in CO and N2 and less in CH4 and NH3. The impact theory has the advantage of possibly generating amorphous solid carbon, which might help to explain the relatively dark color of the satellites. Oriel and Umbriel have a dense population of large impact craters; the population is especially dense for diameters of 50 to 100 km, similar to that observed for many of the oldest and most heavily cratered objects in the solar system. Titania and Ariel have very different crater populations; there are far fewer large craters, and numerical frequency increases in smaller crater sizes, indicative of a secondary impact population and younger surfaces. Miranda’s crater distribution looks like that of Oberon and Umbriel but has an average factor of 3 greater number of any given size (70). Umbriel and Ariel have similar size and mass but dramatically different surface characteristics, indicating differences in evolution or composition. Umbriel (see enlarged view in Fig. 10) is by far the darker, displays a weaker spectral signature for water ice, exhibits very little albedo contrast across most of its surface, and shows no evidence of crater rays at Voyager resolution. The

Miranda

Umbriel

Figure 10. (Left) [USGS P30230]. South polar view of Miranda, produced by a mosaic of nine images obtained by Voyager 2 on 24 January 1986. Older, heavily cratered terrain appears to have low albedo contrast, whereas the younger complex terrain is marked by bright and dark bands, scarps, and ridges. (Right) [JPL P29521]. Southern Hemisphere of Umbriel imaged by Voyager 2 on 24 January 1986. This darkest of the five large moons of Uranus has an ancient, heavily cratered surface and little albedo contrast except for the strange bright ring near the top of the image, which may be a frost deposit associated with an impact crater.

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global albedo pattern suggests a young fresh surface, but the crater population suggests a very ancient surface. The surface of Ariel, on the other hand, is much brighter and seems to be geologically younger. Old, large population I craters have been lost, presumably by some combination of viscous relaxation, as indicated by the slumped configuration of its largest existing crater, and extrusion of material over the surface (indicated by smooth plains). Flows on Ariel (and also Titania) are more likely to be a mixture of ammonia and water, methane, or CO clathrates, due to the lower melting points of these mixtures. Miranda (see enlarged view in Fig. 10) is the smallest of Uranus’ large satellites, the closest to Uranus, and is thus most affected by gravitational focusing of external impactors. Its surface is composed of two very different terrain types: an old, heavily cratered terrain without much albedo contrast and a young, complex terrain that has scarps, ridges, and bright and dark bands. The 11 inner satellites are small and very dark (Table 3). The increasing ice content (inferred from lower density) of satellites closer to Uranus might be due to the higher temperatures closer to Uranus that promote the conversion of CO and N2 to CH4 and NH3. Most of the satellites have nearly circular orbits close to the equatorial plane of Uranus, but the outer four are much more elliptical. Additional small satellites are probably present near the rings and control the narrow rings. Small Satellites of Neptune. Neptune has eight known satellites (Table 4). Voyager was able to resolve surface features clearly on three of the four largest: Triton, Proteus, and Larissa (Fig. 11), but never got close enough to Nereid to obtain a detailed image. The six inner satellites, all discovered by Voyager in 1989, range in size from 29 to 208 km in radius. They all have low geometric albedos of 0.06–0.08 at visible wavelengths and are gray in color. These are darker than Nereid, which has an albedo of 0.155 (71). The inner satellites are in nearly circular orbits within five planetary radii, wheras Nereid is on a distant and very eccentric orbit that extends from 57 to 385 Neptune radii, suggesting Table 4. Satellites of Neptune Satellite

Mean orbital radius, kmb

Naiad 48,227 Thalassa 50,075 Despina 52,526 Galatea 61,953 Larissa 73,548 Proteus 117,647 Triton 354,760 Nereid 5,513,400 a

Radius, kma

Mass, Sidereal Density, Orbital Geometric Discovery 1021 ga period, g/cm3 a inclinaalbedoa dated c days tion, deg

29 0.13 0.296 40 0.34 0.312 74 2.25 0.333 79 2.7 0.429 96 4.8 0.554 208 49 1.121 1,353 21,400 5.877 170 31 360.16

1.3 1.3 1.3 1.3 1.3 1.3 2.07 1.5

4.7 0.2 0.1 0.1 0.2 0.6 156.8 27.6

0.060 0.060 0.059 0.063 0.056 0.061 0.756 0.155

Ref. 68. Ref. 3. c Ref. 69. d Discoveries are dated by data acquisition date (a ¼ Voyager2, b ¼ Lassel, c ¼ Kuiper). b

1989a 1989a 1989a 1989a 1989a 1989a 1846b 1949c

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Proteus

Larissa

Figure 11. Voyager imaging of Triton and newly discovered satellites 1989N1 and 1989N2. (Left) [JPL P34687]. This Voyager 2 image has a resolution of 10 km. The South Pole, which is sunlit throughout the current season, is at bottom left. The absence of large impact craters suggests that Triton’s surface has been renewed within the last billion years. (Upper right) [JPL P34727]. Voyager 2 image of Neptune’s satellite 1989N1 (Proteus) at a resolution of 2.7 km. Its average diameter is 208 km. Its albedo is only 6%, compared to Triton’s 76%, and its color is gray. (Lower right) [JPL P34698]. Voyager 2 image of 1989N2 (Larissa), Neptune’s fourth largest satellite (mean radius 95 km), at a resolution of 4.2 km. It also has a low albedo (about 5%) and seems to have craters 30– 50 km in diameter. This figure is available in full color at http://www.mrw.interscience. wiley.com/esst.

that it may be a captured object rather than having formed in place. Nereid, though smaller than Proteus, is somewhat brighter. It was discovered by Kuiper in 1949. Nereid’s photometric properties suggest that it has a surface of dirty frost. Of these comparably sized satellites, only Proteus has been imaged well enough to permit even a crude map of surface features. Its most prominent feature is an impact basin (72) about 210 km in diameter (larger than the 208-km radius of Proteus). The capture of Triton (see next section) and its evolution would have greatly perturbed any inner satellites, inducing mutual collisions and breakup. Catastrophic disruption would also be likely from external sources. Thus, the inner satellites are likely to be reaccreted debris. It remains unexplained why Nereid is brighter than the inner satellites. One possibility is that its greater distance prevented it from acquiring a veneer of dark particles lost from the rings. Among the known satellites, only Galatea seems to have any dynamic influence on Neptune’s rings, which is to confine the ring arcs azimuthally (73), as discussed in a later section. Triton. Neptune’s satellite, Triton (Figs. 11 and 12), is one of the most peculiar of all satellites. It is the only satellite that has a retrograde orbit (it orbits opposite to the direction of Neptune’s rotation). It was discovered by William Lassel, an amateur astronomer, less than a month after the discovery of Neptune — no

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Figure 12. Close-up views of Triton. (Left) [JPL P34714]. Voyager 2 1989 image of the south polar terrain of Triton, showing 50 dark plume deposits or ‘‘wind streaks’’ on the icy surface. The plumes originate at dark spots a few miles in diameter, and some deposits stretch for more than 100 miles. A few active plumes were observed during the Voyager encounter. (Lower right) [JPL P34690]. Voyager 2 image of irregular dark patches on Triton’s surface. (Upper right) [JPL P34722]. Voyager 2 image of Triton’s cantaloupe-like terrain at a resolution of about 750 m. This terrain form of roughly circular depressions separated by rugged ridges is unique to Triton and covers large areas in its Northern Hemisphere.

mean feat given that it is 200 times fainter than Neptune! Prior to the Voyager encounter with Neptune, there was considerable speculation about the size of Triton and whether or not it had an atmosphere. From its deflection of Voyager’s orbit and imaging of its surface, the size and mass were finally established in 1989. Its diameter of 2710 km is larger than Pluto’s 2300 km, though smaller than Titan’s 3150 km. Its surface albedo of 72% at visible wavelengths (74) was that of an icy surface, but its density of 2.05 grams/cm3 implied that rocky material also had to be a significant component. This density is comparable to that of the Pluto/Charon system. Only a very tenuous atmosphere of mainly nitrogen was observed and a surface pressure of only about 15 microbars. Voyager observations revealed slightly pinkish bright regions and slightly bluish to gray dark regions, but the composition of the surface materials could be obtained only from ground-based spectroscopic observations. In 1978 observations of Triton, Cruikshank and Silvaggio (75) identified the spectral signature of methane gas. Later observations identified spectral features indicating that frozen methane (NH4), solid molecular nitrogen (N2), carbon monoxide (CO), carbon dioxide (CO2), and water (H2O) were present on the surface. Because of their extreme volatility, the existence of N2, CO, and CH4 in solid form on Triton’s

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surface means that Triton must be extremely cold. In fact, Triton’s high reflectivity and great distance from the Sun make it one of the coldest places in the solar system; it has a surface temperature of only 38 K. The small amount of vapor sublimed into the atmosphere at these temperatures (mainly N2) is balanced by condensation from the atmosphere. The equilibrium point at which condensation equals sublimation occurs at a pressure of only 15 microbars (15 millionths the pressure at the surface of Earth). Slightly warmer and slightly colder regions upset this local balance and produce atmospheric pressure variations, winds, and a redistribution of surface materials. As Triton’s seasons lead to variations in the distribution of solar heating, the distribution of Triton’s surface materials is likely to change over time, along with its mean albedo, its light curve (its brightness variation as a function of rotational angle), and perhaps its color. In fact, Triton’s color in 1989 seems to have been considerably less red than it was in 1979 (51). Voyager images of Triton’s limb revealed discrete clouds and thin hazes at altitudes up to 30 km (51). The haze particles are possibly CH4 ice and/or more complex organics created by interactions with UV radiation. Organic compounds may also be responsible for some of the slightly darker material on the surface, although rocky materials are also plausible contributors to surface features. Some of the most remarkable images of Triton were those of the active geyserlike plumes. The geysers were about 8 km tall, rising vertically from a small spot on a surface then abruptly ending in a dark cloud that extended in a narrow (about 5 km wide) plume downwind for as much as 150 km before disappearing (51). The cloud form suggests that horizontal wind speed increases abruptly at 8 km. The plumes are a plausible cause of the numerous dark streaks seen in the south polar region (Fig. 12). Venting of nitrogen gas, or perhaps methane, induced by solar heating (most active plumes are near subsolar latitudes) or localized geothermal energy, seems to entrain small dark particles that form the plume. The solar driven model involves dark material deposited under a layer of transparent frozen nitrogen that becomes heated by the Sun to a point at which the nitrogen in contact with the dark material is vaporized, builds in pressure, and eventually is released through a crack or rupture in the surface. Triton’s seasons are extreme as a result of its inclined and precessing retrograde orbit. Neptune’s spin axis is tipped 291 relative to its orbital plane normal (compared to Earth’s 23.51 tilt relative to its orbit normal). If Triton orbited within Neptune’s equatorial plane, it would have seasonal forcing similar to Neptune’s (and similar to Earth’s as well, though much reduced in amplitude because of its great distance from the Sun). Triton’s orbit, however, is inclined 211 from Neptune’s equatorial plane, and because of the torque exerted by Neptune’s oblate mass distribution, it precesses in a period of 688 years (76). That inclination can either add to or subtract from Neptune’s. At one extreme, Triton’s rotational (and orbital) axis can be tipped by 291 þ 211 ¼ 501 away from the Neptune orbit normal, putting much of one of Triton’s hemisphere in constant darkness throughout its 5.9-day orbital period and leading to heating at latitudes constantly exposed to the Sun and cooling at latitudes that are in constant darkness. This results in a transfer of volatile materials from the heated region to the cooled region, forming a polar cap of fresh ice at one pole and eroding the cap that had previously formed at the other pole. At another extreme, Triton’s

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rotational axis can precess to a point at which it has only an 81 (291  211) inclination to the Neptune orbit normal and thus would receive a much more uniform exposure to solar heating in Northern and Southern Hemispheres. As Neptune orbits the Sun every 164.8 years, the subsolar latitude on Triton will be further modulated between positive and negative values of the current orbital inclination angle, leading to a two-component modulation of seasons on Triton; the shortest period of modulation is the 164.8 year orbital period of Neptune, which has an amplitude envelope modulated by the 688 year period of precession of Triton’s orbital plane about Neptune. The summer solstice of 2000 will be followed by an equinox in 2041, during which Triton’s surface should undergo considerable change. Triton’s surface is unlike that of any other satellite. Much of it appears roughened like the surface of a cantaloupe (upper right image of Fig. 12), formed by multitudes of circular dimples called cavi, which are about 25–30 km in diameter. The largest impact crater is only 27 km in diameter. From the small number of craters observed, Triton’s surface appears to be relatively young geologically. It has large plains, apparently flooded by cryovolcanic fluids. Linear ridges 12–15 km wide indicate cracking of the surface and upwelling of material within the cracks (Figs. 11 and 12). Peculiar irregular blotches that have bright aureoles (Fig. 12, lower right) are of unknown origin.

Ring Systems Uranus has a much more extensive and massive ring system than Neptune, perhaps a consequence of the additional debris generated by the larger impact event that presumably tilted Uranus’ spin axis. Rings of Uranus. The basic characteristics of the rings of Uranus are summarized in Table 5. Rings were first discovered by stellar occultation (dimming of a star’s brightness as rings pass between a star and an observer). By measuring the star’s brightness as a function of time, it is possible to determine ring

Table 5. Rings of Uranusa Ring 1986U2R 6 5 4 Alpha Beta Eta Gamma Delta 1986U1R Epsilon a

Distance, km

Width, km

Optical depth

Albedo

38,000 41,840 42,230 42,580 44,720 45,670 47,190 47,630 48,290 50,020 51,140

2,500 1–3 2–3 2–3 7–12 7–12 0–2 1–4 3–9 1–2 20–100

o0.001 0.2–0.3 0.5–0.6 0.3 0.3–0.4 0.2 0.1–0.4 1.3–2.3 0.3–0.4 0.1 0.5–2.1

0.03 0.03 0.03 0.03 0.03 0.03 0.03 0.03 0.03 0.03 0.03

Ref. 69. Distance is from Uranus’ center to the ring’s inner edge.

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locations and optical depth. In 1977, nine rings were discovered and characterized using stellar occultation (77). These are unofficially named 6, 5, 4, alpha, beta, eta, gamma, delta, and epsilon, in order of increasing distance from Uranus. Samples of the 1986 Voyager imaging of the Uranian rings are displayed in Fig. 13. The left image compares the image in reflected light on approach to Neptune as the image is in forward scattered light, taken after Voyager passed Uranus. Rings that have a significant component of small particles (of the order of a micron) appear much brighter in forward scattering (much like dust on a car windshield). Large particles dominate the appearance in backscattered light. (See the figure caption for further details of this comparison.) The Uranian rings are narrow and sharp-edged and have optical depths of 0.3 or more. Most are very narrow (no more than 10 km wide), inclined to the equatorial plane of Uranus, and eccentric. Exceptions include the gamma and epsilon rings, which are not inclined, and the eta ring, which is not inclined and nearly circular. The epsilon ring is the widest and brightest (the most prominent ring in Fig. 5) and is also the most eccentric. It varies in distance from Uranus by about 800 km, and varies in width from 20 km where it is closest to Uranus, to 100 km where it is furthest. Its variation in radial distance is five times that of the next most eccentric ring (ring 5). It is somewhat of a mystery why orbital speed differences across the epsilon ring do not spread the ring material radially, though it is likely that some satellite resonances are responsible. Two shepherd satellites have been identified for the epsilon ring (right image of Fig. 13). These provide forces tending to confine the ring particles radially. The alpha and beta rings also vary systematically in width, from 5 km to 12 km; extremes are offset

Figure 13. (Left) (Fig. 16–14 of Ref. 78). Voyager 2 images of the Uranian ring system first in backscattered light (upper half) one day before passing by Uranus in Janaury 1986 and second, in forward scattered light (lower half), taken after passing by Uranus and looking backward. The nine labeled rings (upper half) are those discovered by stellar occultation measurements. The forward scattering view dramatically enhances the visibility of micron-sized particles, revealing structures not otherwise visible. Also note the mismatch of the two views of the epsilon ring, a consequence of its significant eccentricity. (Right) [JPL P29466]. Voyager 2 discovered two moons (1986U7, named Cordelia, and 1986U8, named Ophelia) that are shepherd satellites. The inner moon pushes ring particles outward, and the outer moon pushes them inward. This prevents the narrow rings from spreading out.

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by about 301 in orbital longitude from the closest and most distant positions (periapsis and apoapsis). The most inclined rings (6, 5, and 4) deviate from Uranus’ equatorial plane by only 24 to 46 km. A much more complex series of rings is seen in forward scattered light. In Fig. 13, there is an obvious lack of correlation between regions of high dust density (bright in forward scattering due to wavelength-sized particles) and regions of large particles (seen in backscatter and occultation observations). The structure and lack of correlation between dust features and large-particle features are reminiscent of Saturn’s D ring. Rings of Neptune. Neptune’s rings (Table 6 and Fig. 14) are the least understood. Following the discovery of rings around Uranus, stellar occultations were also used to search for rings around Neptune. After several failures, the first detection in 1984 puzzled astronomers by showing rings on only one side of the planet. Voyager imaging in 1989 revealed that though Neptune’s rings did completely encircle the planet, Neptune’s ring particles were not uniformly distributed along the rings. Instead, much of the ring mass was clumped in restricted ring arcs. The outermost (Adams) ring, though continuous, contained three main arcs of much higher particle density of the order of 101 wide in longitude. The confinement of material in the ring arcs, is thought to be the result of gravitational interactions with the moon Galatea (80). One resonance interaction confines the material radially, producing narrow rings, and a second resonance interaction produces clumping of the ring material longitudinally, although the theory predicts regular spacing of clumps. It is not clear why the material is not periodically clumped. Only two rings are prominent in images taken on the sunlit side. The ring material is very dark and probably red (80). The Neptune ring particles are as dark as those in the rings of Uranus, and have a single scattering albedo of about 0.04. A plausible composition is ice mixed with silicates and/or some carbon-bearing material. The Adams and Le Verrier rings contain a significant fraction of dust, comparable to the fraction in Saturn’s F ring or Jupiter’s ring; both are significantly dustier than the main rings of Saturn or Uranus.

Table 6. Rings of Neptune Ring

Distance (km)

Galle

41,900

Le Verrier

53,200

Lassell Aragob Unnamedb 1989N1R Adams

a

53,200–57,200 57,200 61,950 62,930

Width (km) 15a B2000b 15a B110b 5800a B4000b oB100b o50a B50b

Optical depthb B0.00008

B0.015

0.01–0.02a B0.002b 0.0001a B0.00015b

B0.015

0.01–0.1a 0.0045b 0.12(arcs)b

Ref. 69. Distance is from Uranus’ center to the ring’s inner edge. Ref. 79.

b

Albedob

B0.015

B0.015 0.04 (arcs Egalite´ 1, Egalite´ 2)

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Figure 14. (Left) [JPL P35060]. Three clumps, or ring arcs, are visible in this view of Neptune’s outermost Adams ring, imaged by Voyager 2 in August 1989. (Right) [JPL P35023]. This pair of Voyager 2 clear-filter images shows the ring system at the highest sensitivity and at a phase angle of 1341. The brighter, narrow rings are the Adams and Le Verrier rings. Extending out from the Le Verrier ring is the diffuse Lassell ring. The inner medium width ring is the Galle ring. The ring arcs seen in the left-hand image were in the blacked out region between these more sensitive images.

Satellites disrupted by collisions provide a plausible model for the source of Neptune’s ring material. Meteorite collisions seem to be sufficient to explain the dust in the main Uranian rings (81), and in the diffuse Galle and Lassell rings (impacting some unseen parent body), but a more prolific source (perhaps interparticle collisions) is needed to account for the high abundance of dust in the Le Verrier and Adams rings.

Magnetic Fields and Magnetospheres Jupiter’s magnetic field was obvious from the ground because of the synchrotron radiation that it generated, but the magnetic fields of Uranus and Neptune were substantiated only by Voyager observations during close approaches to the planets. The supersonic solar wind particles interact with planetary fields to form a bow shock, and inside the bow shock, the solar wind encounters and is deflected by the planetary field, forming a boundary called the magnetopause. For Uranus, the detached bow shock was observed as a sudden increase in magnetic field intensity upstream at 23.7 Uranus radii (or 23.7 RU), and the magnetopause boundary was seen at 18 RU (82). For Neptune, the corresponding distances (in Neptune radii) were 34.9 RN and 26.5 RN. The magnetospheres are not complete barriers to solar wind particles, however, especially in the polar regions where particles are able to impact the atmosphere and generate an aurora. The UV spectrometer of Voyager 2 found auroral emissions from both Uranus and Neptune. The large offsets and tilts of the magnetic fields (discussed later) lead to auroral zones far from the poles of the rotational axes. The solar wind particles also contribute charged particles that form radiation belts. The Uranian rings and moons are embedded deep within the magnetosphere and thus play a role as absorbers of trapped radiation belt particles, as confirmed for Uranus by depressions in electron counts at magnetic latitudes swept by Miranda, Ariel, and Umbriel (83). A similar situation exists for Neptune’s rings and satellites, except

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for Nereid, which is outside Neptune’s magnetosphere when it is on the sunward side of the planet. The magnetic fields of Uranus and Neptune have unusually large inclinations relative to their rotational poles: the north dipole inclination is 58.61 for Uranus (84) and 471 for Neptune (85). The dipole moments are 48 and 25 times that of Earth (which is 7.9  1025 gauss cm3). For comparison, Jupiter’s moment is 20,000 times that of Earth, Saturn’s 600 times, and respective inclinations are 9.61 and less than 11, compared to Earth’s 111. Besides inclination relative to the spin axis, these magnetic dipoles are also unusual in having large offsets from the planet centers (Fig. 15) by 0.3 RU for Uranus and 0.55 RN for Neptune. However, the magnetic field of Uranus is not purely that of a dipole. It also has a strong quadrupole component, which is most significant close to the planet and contributes to the large variability of the magnetic field near the cloud tops that ranges from 0.1 gauss on the 1986 sunlit hemisphere to 1 gauss at a point on the 1986 dark hemisphere (Earth’s field is about 0.3 gauss near the equator). The rotational periods of the magnetic fields are used to define the rotational periods of the interiors, which are given in Table 1. The primary mechanism for producing magnetic fields in major planets, it is thought, is the dynamo mechanism, which appears to have three basic requirements (86): (1) planetary rotation, (2) a fluid electrically conducting region of the interior, and (3) convection within the conducting fluid. For Jupiter and Saturn, the rotation is rapid, the conducting fluid is metallic hydrogen, and primordial

10°