Category THE DEVELOPMENT. OF PROPULSION. TECHNOLOGY. FOR U. S.. SPACE-LAUNCH. VEHICLES,. 1926-1991

Propulsion for the MX-774B, Viking, and Vanguard

Meanwhile, several American engines drew upon knowledge of the V-2 but also built upon indigenous American experience from be­fore and during World War II. The engines for the MX-774B test missile, the Viking sounding rocket, and the first stage of the Van­guard launch vehicle are examples. Although none of these engines by themselves contributed in demonstrable ways to later launch – vehicle engine technology, the experience gained in developing them almost certainly informed later developments.

MX-774 B

Reaction Motors, Inc. (RMI) developed both the MX-774B power – 112 plant and the Viking engine. The MX-774B engine (designated XLR – Chapter 3 35-RM-1) evolved from the 6000C4 engine the firm had produced

during 1945 for the X-1 rocket plane. Both were comparatively small engines, the 6000C4 yielding 6,000 pounds of thrust and the XLR – 35-RM-1 having a thrust range of about 7,600 to 8,800 pounds. Like the V-2, both engines used alcohol as fuel and liquid oxygen as the oxidizer. The use of alcohol (95 percent ethanol for the MX-774B) suggested some borrowing from the V-2, but the XLR-35-RM-1 achieved a specific impulse of 227 lbf-sec/lbm, significantly higher than that of the V-2 engine. As in the V-2, the MX-774B engine fed the propellants using two pumps operated by the decomposition of hydrogen peroxide, but the U. S. powerplant employed four separate cylinders as combustion chambers rather than the single, spherical chamber for the V-2. Like the German engine, the XLR-35-RM-1 was regeneratively cooled.

As suggested in chapter 1, the major innovations of the MX-774B that influenced launch vehicles were swiveled (not gimballed) en­gines and light, pressurized propellant tanks that evolved into the “steel balloons" used on the Atlas missile. Both of these innovations were the work of Convair (especially Charlie Bossart), the airframe contractor for MX-774B, not RMI, but the four-cylindered engine was integral to the way swiveling worked, so the engine contractor deserves some of the credit. (Each of the four cylinders could swing back and forth on one axis to provide control in pitch, yaw, and roll; a gimbal, by contrast, could rotate in two axes, not simply a single one.) According to one source, the Germans had tried gimballing on the V-2 but had discarded the idea because of the complexities of rotating the 18-pot engine, and Goddard had patented the idea. But actual gimballing of an engine was apparently first perfected on the Viking. Meanwhile, the air force canceled the MX-774B prema­turely, but it did have three test flights in July-December 1948. On the first flight, the engine performed well but an electrical-system failure caused premature cutoff of propellants. On the second flight, the missile broke apart from excessive pressure in the oxygen tank. The third flight was successful.21

THE DEVELOPMENT OF PROPULSION TECHNOLOGY FOR U. S. SPACE-LAUNCH VEHICLES, 1926-1991

THIS BOOK ATTEMPTS TO Fill A GAP IN THE LITERATURE about space-launch vehicles (and in the process, strategic missiles, from which launch vehicles borrowed much technology). There are many excellent books about rocketry. (The Note on Sources discusses many of them.) But none covers the ways in which the technology in the United States developed from its beginnings with Robert Goddard and the German V-2 project through the end of the cold war. This book concentrates on propulsion technology to keep its length manageable, but it occasionally mentions structures and guidance/control in passing, especially in chapters 1 and 2.

Besides the lack of coverage of the evolution of rocket technol­ogy in the existing literature, there is a severe problem with accu­racy of details. Apparently reputable sources differ about matters as simple as lengths and diameters of vehicles and details about thrust. I cannot claim to have provided definitive measurements, but I have tried to select the most plausible figures and have pro­vided various references in endnotes that readers can consult to find for themselves the many discrepancies.

I have been working on some aspects of this book since 1992. I initially wrote a much longer manuscript, organized by project, that covered the entire gamut of major technologies. I have organized this much shorter volume by types of propulsion with overviews in chapters 1 and 2 to provide context and cover factors in technology development that do not fit comfortably in chapters 3-7.

In researching and writing both manuscripts, I received help from a huge number of people. I apologize in advance for any I in­advertently neglect to mention or whose names I have forgotten. I especially want to thank Roger Launius. As my boss at the NASA History Office, he provided unfailing encouragement and support for my initial research. Now as editor for the series in which this book will appear, he has continued that support. Michael Neufeld at the Smithsonian National Air and Space Museum (NASM) shared his own research on the V-2 with me and arranged for me to consult the captured documents held by his institution. He also read many chapters in draft and offered suggestions for improvement. I am fur­ther greatly indebted to NASM for granting me the Ramsey Fellow­ship for 1991-92 and allowing me to continue my research there with the support of archivists, librarians, curators, docents, and vol­unteers, including John Anderson, Tom Crouch, David DeVorkin, Marilyn Graskowiak, Dan Hagedorn, Gregg Herken, Peter Jakab,

Mark Kahn, Daniel Lednicer, Brian Nicklas, George Schnitzer, Paul Silbermann, Leah Smith, Larry Wilson, Frank Winter, and Howard S. Wolko.

Подпись: x Preface Special thanks are due to Glen Asner of the NASA History Divi­sion, who read an earlier version of this book and offered detailed editorial advice at a time when NASA intended to publish the book. Glen’s advice was extremely valuable, as was that of three anony­mous NASA readers. Then Texas A&M University Press accepted the book for publication. Glen and Steven Dick, NASA chief histo­rian, graciously relinquished the manuscript to Texas A&M.

Chapters 1, 2, 6, and 7 contain material I published earlier in chapter 6 of To Reach the High Frontier: A History of U. S. Launch Vehicles, ed. Roger D. Launius and Dennis R. Jenkins (Lexington: University Press of Kentucky, 2002). The material in the present book results from much research done since I wrote that chapter, and it is organized differently. But I am grateful to Mack McCor­mick, rights manager, the University Press of Kentucky, for con­firming my right to reuse the material that appeared in the earlier version his press published.

A number of people read earlier versions of the material in this book and offered suggestions for improvement. They include Matt Bille, Roger Bilstein, Trong Bui, Virginia Dawson, Ross Felix, Pat Johnson, John Lonnquest, Ray Miller, Fred Ordway, Ed Price, Milton Rosen, David Stumpf, and Jim Young. Many other people provided documents or other sources I would otherwise have been unable to locate easily, including Nadine Andreassen, Liz Babcock, Scott Carlin, Robert Corley, Dwayne Day, Bill Elliott, Robert Geisler, Robert Gordon, Edward Hall, Charles Henderson, Dennis Jenkins, Karl Klager, John Lonnquest, Ray Miller, Tom Moore, Jacob Neufeld, Fred Ordway, Ed Price, Ray Puffer, Karen Schaffer, Ronald Sim­mons, Ernst Stuhlinger, Ernie Sutton, Robert Truax, and P. D. Um – holtz. Archivists, historians, and librarians at many locations were unfailingly helpful. Here I can only single out Air Force Historical Research Agency archivist Archangelo Difante and Air Force Space and Missile Systems Center historian and archivist (respectively) Harry Waldron and Teresa Pleasant; China Lake historian Leroy Doig; Laguna Niguel archivist Bill Doty; Clark University Coordi­nator of Archives & Special Collections Dorothy E. Mosakowski; NASA archivists Colin Fries, John Hargenrader, Jane Odom, and Lee Saegesser; and JPL archivists John Bluth, Barbara Carter, Dudee Chiang, Julie Cooper, and Margo E. Young for their exceptional assistance. Several reference librarians at the Library of Congress should be added to the list, but I do not know their names.

I also want to thank everyone who consented to be interviewed (included in endnote references) for their cooperation and agreement to allow me to use the information in the interviews. In addition, many people discussed technical issues with me or provided other technical assistance. These include Ranney Adams, Wil Andrepont, Stan Backlund, Rod Bogue, Al Bowers, George Bradley, Robert Cor­ley, Daniel Dembrow, Mike Gorn, Mark Grills, John Guilmartin, Burrell Hays, J. G. Hill, Ken Iliff, Fred Johnsen, Karl Klager, Franklin Knemeyer, Dennis B. Mahon, Jerry McKee, Ray Miller, Ed Price, Bill Schnare, Neil Soderstrom, Woodward Waesche, Herman Way – land, and Paul Willoughby. Finally, I offer my deep appreciation to my excellent copyeditor, Cynthia Lindlof; my in-house editor, Jen­nifer Ann Hobson; editor-in-chief Mary Lenn Dixon; and everyone else at Texas A&M University Press for their hard work in getting this book published and marketed. To all of the people above and others whose names I could not locate, I offer my thanks for their assistance.

Подпись: xi Preface It goes (almost) without saying that these people bear no respon­sibility for the interpretation and details I provide in the following pages. I hope, however, that they will approve of the uses I have made of their materials, suggestions, comments, and information.

THE DEVELOPMENT OF PROPULSION TECHNOLOGY FOR U. S. SPACE-LAUNCH VEHICLES, 1926-1991

Delta Upper Stages

Подпись:One example of this influence was a second stage used on the Delta launch vehicle featuring an Aerojet engine designated AJ10-118F. This was the F model in a series of engines originally derived from the Vanguard second stage. This particular engine was similar to the Transtage propulsion unit (AJ10-138), but the two were not identical. Both used a fiberglass combustion chamber impregnated with resin and an ablative lining for cooling. Like the Titan II en­gines, both used a mixture of 50 percent hydrazine and 50 percent UDMH as the fuel, igniting hypergolically with a nitrogen-tetroxide oxidizer. This replaced the IRFNA used on earlier versions of the AJ10-118 and increased the specific impulse from more than 265 to upward of 290 lbf-sec/lbm. The F version of the engine had a thrust of between 9,235 and 9,606 pounds, well above the roughly 7,575 pounds of the earlier versions, and it was capable of up to 10 starts in orbit. The new engine completed its preliminary design in 1970 but did not fly until July 23, 1972.59

Another Delta upper stage that used storable propellants was TRW’s TR-201 second-stage engine. Design of this engine began in October 1972, its combustion chamber made of quartz phenolic, with cooling by ablation as had been true of the AJ10-118F, but the TR-201 weighed only 298 pounds in contrast to the Aerojet engine’s reported dry weight of 1,204 pounds. Both engines burned nitrogen tetroxide with a 50/50 mixture of hydrazine and UDMH, igniting hypergolically, but the TRW propulsion unit yielded 9,900 pounds of thrust, whereas the Aerojet unit yielded a maximum of 9,606 pounds. The newer system did provide only 5 instead of 10 restarts.60

Titan I and Titan II

Simultaneously with the development of Polaris and then Minute – man, the air force continued work on two liquid-propellant missiles, the Titans I and II. The Titan II introduced storable propellants into the missile inventory and laid the groundwork for the core portion of the Titans III and IV space-launch vehicles. Titan I began as es­sentially insurance for Atlas in case the earlier missile’s technology proved unworkable. The major new feature of the first of the Titans was demonstration of the ability to start a large second-stage engine at a high altitude.93 The WAC Corporal had proved the viability of the basic process involved, and Vanguard would develop it further (after Titan I was started). But in 1955, using a full second stage on a ballistic missile and igniting it only after the first-stage engines had exhausted their propellants seemed risky.

The air force approved development of Titan I on May 2, 1955. Meanwhile, the Western Development Division had awarded a 44 contract on January 14, 1955, to Aerojet for engines burning liq – Chapter 1 uid oxygen and a hydrocarbon fuel for possible use on Atlas. These soon evolved into engines for the two-stage missile. Even though the Aerojet engines burned the same propellants as Atlas, there were problems with development, showing that rocket engineers

still did not have the process of design “down to a science." Despite the change in propellants, the Titan II used a highly similar design for its engines, making Aerojet’s development for that missile less problematic than it might otherwise have been (although still not without difficulties), with technology then carrying over into the Titans III and IV core launch vehicles. Meanwhile, the air force de­ployed the Titan Is in 1962. They quickly deactivated in 1965 with the deployment of Minuteman I and Titan II, but Titan I did provide an interim deterrent force.94

The history of the transition from Titan I to Titan II is compli­cated. One major factor stimulating the change was the 15 minutes or so it took to raise Titan I from its silo, load the propellants, and launch it. Another was the difficulty of handling Titan I’s extremely cold liquid oxygen used in Titan I inside a missile silo. One solution to the twin problems would have been conversion to solid propel­lants like those used in Polaris and Minuteman, but another was storable propellants. Under a navy contract in 1951, Aerojet had begun studying hydrazine as a rocket propellant. It had good perfor­mance but could detonate. Aerojet came up with a compromise so­lution, an equal mixture of hydrazine and unsymmetrical dimethyl hydrazine, which it called Aerozine 50. With nitrogen tetroxide as an oxidizer, this fuel mixture ignited hypergolically (upon contact with the oxidizer, without the need for an ignition device), offering a much quicker response time than for Titan I.95 As a result of this and other issues and developments, in November 1959 the Depart­ment of Defense authorized the air force to develop the Titan II. The new missile would use storable propellants, in-silo launch, and an all-inertial guidance system.96

Подпись: 45 German and U.S. Missiles and Rockets, 1926-66 On April 30, 1960, the Air Force Ballistic Missile Division’s de­velopment plan for Titan II called for it to be 103 feet long (compared to 97.4 feet for Titan I), have a uniform diameter of 10 feet (whereas Titan I’s second stage was only 8 feet across), and have increased thrust over its predecessor. This higher performance would increase the range with the Mark 4 reentry vehicle from about 5,500 nauti­cal miles for Titan I to 8,400. With the new Mark 6 reentry vehicle, which had about twice the weight and more than twice the yield of the Mark 4, the range would remain about 5,500 nautical miles. Because of the larger nuclear warhead it could carry, the Titan II served a different and complementary function to Minuteman I’s in the strategy of the air force, convincing Congress to fund them both. It was a credible counterforce weapon, whereas Minuteman I served primarily as a countercity missile, offering deterrence rather than the ability to destroy enemy weapons in silos.97

In May 1960, the air force signed a letter contract with the Mar­tin Company to develop, produce, and test the Titan II. It followed this with a contract to General Electric to design the Mark 6 reentry vehicle. In April 1959, AC Spark Plug had contracted to build an in­ertial guidance system for a Titan missile, although it was not clear at the time that this would be the Titan II.98

Although the Titan II engines were based on those for Titan I, the new propellants and the requirements in the April 30 plan necessi­tated considerable redesign. Because the new designs did not always work as anticipated, the engineers had to resort to empirical solu­tions until they found the combinations that provided the necessary performance. Even with other changes to the Titan I engine designs, the Titan II propulsion system had significantly fewer parts than its Titan I predecessor, reducing chances for failure during operation. Despite the greater simplicity, the engines had higher thrust and higher performance, as planned.99

Flight testing of the Titan II had its problems, complicated by plans to use the missile as a launch vehicle for NASA’s Project Gem­ini, leading to the Project Apollo Moon flights. However, the last 13 flights in the research-and-development series were successful, giving the air force the confidence to declare the missile fully opera­tional on the final day of 1963. Between October and December 1963, the Strategic Air Command deployed six squadrons of nine Titan IIs apiece. They remained a part of the strategic defense of the United States until deactivated between 1984 and 1987. By that time, fleet ballistic missiles and smaller land-based, solid-propellant ballistic missiles could deliver (admittedly smaller) warheads much more accurately than could the Titan IIs. Deactivation left the former operational Titan II missiles available for refurbishment as space – launch vehicles.100

Development of Titan I and Titan II did not require a lot of new technology. Instead, it adapted technologies developed either ear­lier or simultaneously for other missile or launch-vehicle programs. Nevertheless, the process of adaptation for the designs of the two Titan missiles generated problems requiring engineers to use their fund of knowledge to find solutions. These did work, and Titan II became the nation’s longest-lasting liquid-propellant missile with the greatest throw weight of any vehicle in the U. S. inventory.

The Sergeant Missile Powerplant

Meanwhile, the first major application of the technologies devel­oped for the RV-A-10 was the Sergeant missile, for which JPL began planning in 1953 under its ORDCIT contract with Army Ordnance. JPL submitted a proposal for a Sergeant missile in April 1954, and on June 11, 1954, the army’s chief of ordnance programmed $100,000 for it. At the same time, he transferred control of the effort to the com­manding general of Redstone Arsenal. Using lessons learned from the liquid-propellant Corporal missile, JPL proposed a co-contractor for the development and ultimate manufacture of the missile. In February 1956, a Sergeant Contractor Selection Committee unani­mously chose Sperry Gyroscope Company for this role, based on JPL’s recommendation and Sperry’s capabilities and experience with other missiles, including the Sparrow I air-to-air missile system for the navy. In April 1954, the Redstone Arsenal had reached an agreement with the Redstone Division of Thiokol to work on the solid-propellant motor for the Sergeant, with the overall program to develop Sergeant beginning in 1955.30

There is no need to provide a detailed history of the Sergeant missile here. It took longer to develop than originally planned and was not operational until 1962. By then the navy had completed the far more significant Polaris A1, and the air force was close to field­ing the much more important Minuteman I. The Sergeant did meet a slipped ordnance support readiness date of June 1962 and became a limited-production weapons system until June 1968. It did equal its predecessor, Corporal, in range and firepower in a package only

FIG. 6.3

The Sergeant Missile PowerplantTechnical drawing of the Jupiter C (actually,

Juno I, including scaled-down Sergeant upper stages) with America’s first satellite, Explorer I, showing the latter’s characteristics. (Photo courtesy of NASA)

The Sergeant Missile Powerplant

half as large and requiring less than a third as much ground-support equipment. Its solid-propellant motor could also be readied for fir­ing much more quickly than the liquid-propellant Corporal.31

The Sergeant motor was a modification or direct descendant of the RV-A-10’s motor. The latter (using the TRX-110A propulsion formulation) employed 63 percent ammonium perchlorate as an oxidizer, whereas the TP-E8057 propellant for the Sergeant motor (designated JPL 500) had 63.3 percent of that oxidizer and 33.2 per­cent LP-33 liquid polymer in addition to small percentages of a curing agent, two reinforcing agents, and a curing accelerator. At a nozzle expansion ratio of 5.39, its specific impulse was about 185 lbf-sec/lbm, considerably lower than the performance of Polaris A1. It employed a five-point-star grain configuration, used a case of 4130 steel at a nominal thickness of 0.109 inch (almost half that of the RV-A-10 case), and a nozzle (like that of the RV-A-10) using 1020 steel with a graphite nozzle-throat insert.32 Ironically, perhaps, the main contributions the Sergeant made to launch-vehicle tech­nology were through a scaled-down version of the missile used for testing. These smaller versions became the basis for upper stages in reentry test vehicles for the Jupiter missile and in the launch vehicles for Explorer and Pioneer satellites.33

VIKING

Подпись: 113 Propulsion with Alcohol and Kerosene Fuels, 1932-72 Whereas MX-774B was an army air forces/air force project, the navy sponsored the Viking, with Milton W. Rosen responsible for the de­velopment and firing of the rocket. Reaction Motors designed the engine, drawing on its own experience as well as data from the V-2, with the Glenn L. Martin Company designing and building the over­all rocket. Viking’s pioneering development and use of gimballing were the responsibility of Martin engineers. Like the MX-774B powerplant, Viking drew on the V-2 technology, but Rosen and his engineers designed it specifically for upper-atmospheric research. Rosen’s specifications called for a thrust of 20,000 pounds compared with 56,000 for the V-2. Under a contract initiated in September 1946, RMI designed an engine (XLR-10-RM-2) with a single cylin­drical thrust chamber.

As with the V-2, the U. S. rocket’s propellants were alcohol and liquid oxygen, pumped into the combustion chamber by turbines driven by decomposed hydrogen peroxide. Whereas the V-2 had used alcohol at 75 percent strength and hydrogen peroxide at 82 percent, the Viking used 95 percent ethyl alcohol and 90 percent hydrogen peroxide. Edward A. Neu did the detailed design work on the com­bustion chamber and injector. Tests caused parts to fail and be re­placed. Burnthroughs of the steel combustion-chamber liner (inner

wall) led to the substitution of pure nickel, the first known use of this metal for such a purpose. Its superior thermal conductivity and higher melting point solved the cooling problem in conjunction with the regenerative cooling in the original design. One injector caused an explosion, so new designs were necessary. Valves were a problem until M. E. “Bud" Parker borrowed valve designs from the MX-774B engine, which thus did influence at least the Viking design.

After the first launch of a Viking rocket from White Sands, New Mexico, on May 3, 1949, the vehicle experienced component failure, leading to subsequent improvements. As a result, each of the dozen Viking rockets fired through the last launch on February 4, 1955, differed from its predecessor. Rosen thought this was the most im­portant aspect of the program. One example was the growth of the thrust of the various Vikings from 20,450 pounds on the first launch to 21,400 on two others. Even though the engine itself was generally successful, it made no known contributions to engine technology per se other than the experience gained by RMI, Martin, and navy engineers. The real contribution of Viking lay in the gimballing sys­tem for steering, not pure propulsion.22

INTRODUCTION

ALTHOUGH ROCKETS BURNING BLACK POWDER HAD existed for centuries, only in 1926 did Robert H. Goddard, an Amer­ican physicist and rocket developer, launch the first known liquid – propellant rocket. It then took the United States until the mid – 1950s to begin spending significant sums on rocket development. The country soon (January 1958) began launching satellites, and by the end of the cold war (1989-91), the United States had developed extraordinarily sophisticated and powerful missiles and launch ve­hicles. From the Atlas to the Space Shuttle, these boosters placed an enormous number of satellites and spacecraft into orbits or trajecto­ries that enabled them to greatly expand our understanding of Earth and its universe and to carry voices and images from across the seas into the American living room almost instantaneously. What allowed the United States to proceed so quickly from the compara­tively primitive rocket technology of 1955 to almost routine access to space in the 1980s?

This book provides answers to that question and explains the evolution of rocket technology from Goddard’s innovative but not fully successful rockets to the impressive but sometimes problem­atic technology of the Space Shuttle. Although propulsion technol­ogy has often challenged the skills and knowledge of its developers, by and large, its achievements have been astonishing.

This combination of complexity and sophistication caused some inventive soul to coin the term “rocket science." But often in the history of rocketry, so-called rocket scientists ran up against prob­lems they could not fully understand. To solve such problems, they often had to resort to trial-and-error procedures. Even as understand­ing of many problems continually grew, so did the size and perfor­mance of rockets. Each increase in scale posed new problems. It turns out that rocketry is as much art as science. As such, it best fits the definitions of engineering (not science) that students of technol­ogy, including Edwin Layton, Walter Vincenti, and Eugene Ferguson, have provided. (Besides engineering as art, they have discussed the discipline’s emphasis on doing rather than just knowing, on artifact design instead of understanding, and on making decisions about such design in a context of imperfect knowledge.)1

In the light of their findings and the details of rocket development discussed in the present book, I will argue that designers and de­velopers of missiles and space-launch vehicles were fundamentally engineers, not scientists, even though some of them were trained

as scientists. For instance, Ronald L. Simmons received a B. A. in chemistry from the University of Kansas in 1952 and worked for 33 years as a propulsion and explosive chemist at Hercules Powder 2 Company, a year with Rocketdyne, and 13 years for the U. S. Navy, Introduction contributing to upper stages for Polaris, Minuteman, and other mis­siles. He considered himself to be a chemist and as such, a scientist, but admitted that he had done “a lot of engineering." Unconsciously underlining points made about engineering by Vincenti and Layton, he added that it was “amazing how much we don’t know or under­stand, yet we launch large rockets routinely. . . and successfully." He believed that we understood enough “to be successful. . . yet may not understand why."2

This is not to suggest any lack of professional expertise on the part of rocket engineers. Rocketry remains perplexingly complex. In the early years, engineers’ knowledge of how various components and systems interacted in missiles and launch vehicles was necessar­ily limited. But quickly data, theory, and technical literature grew to provide them a huge repository of information to draw upon. Some processes nevertheless remained only partially understood. But when problems occur, as they still do, the fund of knowledge is great, permitting designers and developers to focus their efforts and bring their knowledge to bear on specific kinds of solutions.

Yet often there are no clear answers in the existing literature. En­gineers must try out likely solutions until one proves to be effective, whether sooner or later. In the chapters that follow, I sometimes refer to this approach as cut-and-try (cutting metal and trying it out in a rocket) or trial-and-error. Neither term implies that practition­ers were experimenting blindly. They brought their knowledge and available literature (including science) to bear on the problem, mea­sured the results as far as possible, and made informed decisions. Limited funding and rigorous schedules often restricted this process. Given these circumstances, it is remarkable that they succeeded as often and well as they did. Not rocket science, this cut-and-try methodology was part of a highly effective engineering culture.

This book is about launch-vehicle technology. Because much of it originated in missile development, there is much discussion of missiles. These missiles launch in similar fashion to launch ve­hicles. But they follow a ballistic path to locations on Earth rather than somewhere in space. Their payloads are warheads rather than satellites or spacecraft. Especially in the discipline of propulsion, they have employed similar technology to that used in launch ve­hicles. Many launch vehicles, indeed, have been converted missiles or have used stages borrowed from missiles.

Both types of rockets use a variety of technologies, but this book focuses on propulsion as arguably more fundamental than such other fields as structures and guidance/control. The book starts with Goddard and his Romanian-German rival Hermann Oberth. It fol­lows the development of technology used on U. S. launch vehicles through the end of the cold war. Because the German V-2 influenced American technology and was a (not the) starting place for the Saturn launch vehicles in particular, there is a section on the Ger­man World War II missile and its developers, many of whom, under Wernher von Braun’s leadership, came to this country and worked on the Saturns.

Подпись: IntroductionChapters 1 and 2 provide an overview of missile and rocket de­velopment to furnish a context for the technical chapters that fol­low. Chapters 3 through 7 then cover the four principal types of chemical propulsion used in the missiles and launch vehicles cov­ered in chapters 1 and 2. Chapter 8 offers some general conclusions about the process of rocket engineering as well as an epilogue point­ing to major developments that occurred after the book ends at the conclusion of the cold war. (There is no discussion of attempts at harnessing nuclear [and other nonchemical types of] propulsion— sometimes used in spacecraft—because funding restraints and tech­nical risks precluded their use in production missiles and launch vehicles.)3

The book stops about 1991 because after the cold war ended, de­velopment of launch vehicles entered a new era. Funding became much more restricted, and technology began to be borrowed from the Russians, who had followed a separate path to launch-vehicle development during the Soviet era.

Most readers of this book presumably have watched launches of the Space Shuttle and other space-launch vehicles on television, but maybe a discussion of the fundamentals of rocketry will be useful to some. Missiles and launch vehicles lift off through the thrust pro­duced by burning propellants (fuel and oxidizer). The combustion produces expanding, mostly gaseous exhaust products that a nozzle with a narrow throat and exit cone cause to accelerate, adding to the thrust. Nozzles do not work ideally at all altitudes because of changing atmospheric pressure. Thus, exit cones require different angles at low and higher altitudes for optimum performance. For this reason, rockets typically use more than one stage both to al­low exit cones to be designed for different altitudes and to reduce the amount of weight each succeeding stage must accelerate to the required speed for the mission in question. As one stage uses up its propellants, it drops away and succeeding stages ignite and assume

the propulsion task, each having less weight to accelerate while taking advantage of the velocity already achieved.

Подпись: IntroductionMost propellants use an ignition device to start combustion, but so-called hypergolic propellants ignite on contact and do not need an igniter. Such propellants usually have less propulsive power than such nonhypergolic fuels and oxidizers as the extremely cold (cryo­genic) liquid hydrogen and liquid oxygen. But they also require less special handling than cryogenics, which will boil off if not loaded shortly before launch. Hypergolics can be stored for comparatively long periods in propellant tanks and launched almost instantly. This provided a great advantage for missiles and for launches that had only narrow periods of time in which to be launched to line up with an object in space that was moving in relation to Earth.

Solid-propellant motors also allowed rapid launches. They were simpler than and usually not as heavy as liquid-propellant engines. Solids did not need tanks to hold the propellants, high pressure or pumps to deliver the propellant to the combustion chamber, or ex­travagant plumbing to convey the liquids. Normally, rocket firms loaded the solid propellant in a case made of thin metal or compos­ite material. Insulation between the propellant and the case plus an internal cavity in the middle of the propellant protected the case from the heat of combustion, the propellant burning from the cav­ity outward so that the propellant lay between the burning surface and the insulation. The design of the internal cavity provided op­timal thrust for each mission, with the extent of the surface fac­ing the cavity determining the amount of thrust. Different designs provided varying thrust-time curves. Solid propellants did pose the problem that they could not easily accommodate stopping and re­starting of combustion, as liquids could do by using valves. Con­sequently, solids usually served in initial stages (called stage zero) to provide large increments of thrust for earth-escape, or in upper stages. For most of the period of this book, the Scout launch vehicle was unique in being a fully solid-propellant vehicle.

Liquid propellants typically propelled the core stages of launch vehicles, as in the Atlas, Titan, Delta, and Space Shuttle. Upper stages needing to be stopped and restarted in orbit (so they could insert satellites and spacecraft into specific orbits or trajectories af­ter coasting) also used liquid propellants, as did stages needing high performance. But in liquid-propellant engines, the injection of fu­els and oxidizers into combustion chambers remained problematic in almost every new design or upscaling of an old design. Mixing the two types of propellants in optimal proportions often produced instabilities that could damage or destroy a combustion chamber.

This severe problem remained only partly understood, and although engineers usually could find a solution, doing so often took much trial and error in new or scaled-up configurations. Solid propellants were by no means immune to combustion instability, although the problems they faced were somewhat different from those occurring in liquid-propellant engines. And often, by the time solid-propellant instabilities were discovered, design was so far along that it became prohibitively expensive to fix the problem unless it was especially severe.

Подпись: IntroductionBesides propulsion, missiles and launch vehicles required struc­tures strong enough to withstand high dynamic pressures during launch yet light enough to be lifted into space efficiently; aerody­namically effective shapes (minimizing drag and aerodynamic heat­ing); materials that could tolerate aerothermodynamic loads and heating from combustion; and guidance/control systems that pro­vide steering through a variety of mechanisms ranging from vanes, canards, movable fins, vernier (auxiliary) and attitude-control rock­ets, and fluids injected into the exhaust stream, to gimballed en­gines or nozzles.4

With these basic issues to deal with, how did the United States get involved in developing missiles and rockets on a large scale? What sorts of problems did developers need to overcome to permit a rapid advance in missile and launch-vehicle technology? The chap­ters that follow answer these and other questions, but maybe a brief summary of how the process worked will guide the reader through a rather technical series of projects and developments.

Launch-vehicle technology emerged from the development and production of missiles to counter a perceived threat by the Soviet Union. In this environment, heavy cold-war expenditures to de­velop the missiles essentially fueled progress. In addition, many other factors (not always obvious to contemporaries) helped further the process. No short list of references documents the complex de­velopment discussed in this book, but one element of the effort was an innovative and flexible engineering culture that brought together a variety of talents and disciplines in a large number of organiza­tions spanning the nation. People from different disciplines joined together in cross-organizational teams to solve both unanticipated and expected problems.

Likewise, supporting problem solving and innovation was a grad­ually developing network that shared data among projects. Although military services, agencies, and firms often competed for roles and missions or contracts, the movement of people among the compet­ing entities, actual cooperation, professional organizations, partner-

ing, federal intellectual-property arrangements, and umbrella or­ganizations such as the Chemical Propulsion Information Agency promoted technology transfers of importance to rocketry. At the 6 same time, the competition spurred development through the urge Introduction to outperform rivals.

A further factor helping to integrate development and keep it on schedule (more or less) consisted of numerous key managers and management systems. In some instances, managers served as heterogeneous engineers, managing the social as well as the tech­nical aspects of missile and launch-vehicle development, stimulat­ing support for rocketry in general from Congress, the administra­tion, and the Department of Defense. By creating this support, they practiced what some scholars have defined as social construction of the technologies in question. At times, managers engaged in both technical direction and heterogeneous engineering, while in other cases technical managers and heterogeneous engineers were sepa­rate individuals.5

Although rocket technology is complex, I have tried to present it in a way that will be comprehensible to the general reader. The primary audience for this book will tend to be scholars interested in the history of technology or propulsion engineers seeking an over­view of the history of their discipline. I have included many ex­amples of problems encountered in the development of missiles and launch vehicles and explained, as far as I could determine, the way they were resolved. Even though I have not written in the techni­cally rigorous language of engineering (or in some cases because of that), I hope my discussion of the evolution of propulsion technol­ogy will engage the interest of everyone from rocket enthusiasts to technical sophisticates.

SPACE-LAUNCH – VEHICLE TECHNOLOGY

Подпись: German and U.S. Missiles and Rockets, 1926-66evolved from the development of early rockets and missiles. The earliest of these rockets that led to work on launch vehicles themselves was Robert Goddard’s in 1926, generally regarded as the first

liquid-propellant rocket to fly. But it was not until the mid-1950s that significant progress on large missiles occurred in the United States, greatly stimulated by the cold war between the United States and the Soviet Union. (Of course, the Germans had already devel­oped the A-4 [V-2] in the 1940s, and the United States launched a se­ries of reconstructed German V-2s in the New Mexico desert from 1946 to 1952.) Missile development was especially important in furthering the development of launch vehicles because many mis­siles became, with adaptations, actual stages for launch vehicles. In other cases, engines or other components for missiles became the bases for those on launch vehicles. By 1966 large, powerful, and comparatively sophisticated launch vehicles had already evolved from work on early missiles and rockets.

Centaur Propulsion

Подпись:Before the defense establishment transferred the technology from Project Suntan to rocketry, it had to be nudged by a proposal from Convair’s Krafft Ehricke. Called to service in a Panzer division on the western and then eastern fronts during World War II, the young German was still able to earn a degree in aeronautical engineering at the Berlin Technical Institute. He was fortunate enough to be as­signed to Peenemunde in June 1942, where he worked closely with Thiel. Although he came to the United States as part of von Braun’s group and moved with it to Huntsville, Ehricke was a much less conservative engineer than von Braun. Whether for that reason or others, he transferred to Bell Aircraft in 1952 when it was working on the Agena upper stage and other projects. Not happy there either by the time he left (when he believed interest had shifted away from space-related projects), he heeded a call from Karel Bossart to work at Convair in 1954.5

At the San Diego firm, Ehricke initially served as a design spe­cialist on Atlas and was involved with Project Score. By 1956, he was beginning to study possible boosters for orbiting satellites but could find no support for such efforts until after Sputnik I. Then, General Dynamics managers asked him to design an upper stage for Atlas. (Consolidated Vultee Aircraft Corporation merged into General Dynamics Corporation on April 29, 1954, to become the Convair Division of the larger firm.) Ehricke and other engi­neers, including Bossart, decided that liquid hydrogen and liquid oxygen were the propellants needed. Ehricke worked with Rock – etdyne to develop a proposal titled “A Satellite and Space Devel­opment Plan." This featured a four-engine stage with pressure feeding of the propellants, neither Rocketdyne nor Ehricke be­ing aware of Pratt & Whitney’s pumps. In December 1957, James Dempsey, vice president of the Convair Division, sent Ehricke and another engineer off to Washington, D. C., to pitch the design to the air force.6

The air service did not act on the proposal, but on February 7,

1958, Ehricke presented it to the new Advanced Research Projects Agency, created by the Department of Defense. For a time, ARPA exercised control over all military and civilian space projects before relinquishing the civilian responsibility to NASA in October 1958. Thereafter, for a year, ARPA remained responsible for all military space projects, including budgets. The new agency made Ehricke aware of Pratt & Whitney’s hydrogen pumps and encouraged Con – vair to submit a proposal using two 15,000-pound-thrust, pump-fed engines, which it did in August 1958. That same month, ARPA is­sued order number 19-59 for a high-energy, liquid-propellant upper stage to be developed by Convair-Astronautics Division of General Dynamics Corporation, with liquid-oxygen/liquid-hydrogen en­gines to be developed by Pratt & Whitney.7

In October and November 1958, at ARPA’s direction the air force followed up with contracts to Pratt & Whitney and Convair for the development of Centaur, but NASA’s first administrator, Keith Glen – nan, requested that the project transfer to his agency. Deputy Secre – 176 tary of Defense Donald Quarles agreed to this arrangement in prin – Chapter 5 ciple, but ARPA and the air force resisted the transfer until June 10,

1959, when NASA associate administrator Richard E. Horner pro­posed that the air force establish a Centaur project director, locate him at the Ballistic Missile Division in California, but have him report to a Centaur project manager at NASA Headquarters. NASA would furnish technical assistance, with the air force providing ad­ministrative services. The DoD agreed, and the project transferred to NASA on July 1, 1959. Lieutenant Colonel Seaberg from the Sun­tan project became the Air Research and Development Command project manager for Centaur in November 1958, located initially at command headquarters on the East Coast. Seaberg remained in that position with the transfer to NASA but moved his location to BMD. Milton Rosen became the NASA project manager. In November 1958, Ehricke became Convair’s project director for Centaur.8

Complicating Centaur’s development, in the fall of 1958 NASA engineers had conceived of using the first-stage engine of Vanguard as an upper stage for Atlas, known as Vega. NASA intended that it serve as an interim vehicle until Atlas-Centaur was developed. Under protest from Dempsey that Convair already had its hands full with Atlas and Centaur, on March 18, 1959, NASA contracted with General Dynamics to develop Atlas-Vega. With the first flight of the interim vehicle set for August 1960, Vega at first became a higher priority for NASA than Centaur. As such, it constituted an impediment to Centaur development until NASA canceled Vega on

December 11, 1959, in favor of the DoD-sponsored Agena B, which had a development schedule and payload capability similar to Ve­ga’s but a different manufacturer (Bell).9

Besides Vega’s competition for resources until this point, another hindrance to development of Centaur came from liquid hydrogen’s physical characteristics. Its very low density, extremely cold boiling point, low surface tension, and wide range of flammability made it extremely difficult to work with. Ehricke had some knowledge of this from working with Thiel, but the circumstances of the contract with the air force limited the amount of testing he could perform to overcome hydrogen’s peculiarities.10

Подпись:One limitation was funding. When ARPA accepted the initial proposal and assigned the air force to handle its direction, the stipu­lations were that there be no more than $36 million charged by Convair-Astronautics for its work, that a first launch attempt occur by January 1961, and that the project not interfere with Atlas devel­opment. At the same time, Convair was to use off-the-shelf equip­ment as well as Atlas tooling and technology as much as possible. Funding for the Pratt & Whitney contract was $23 million, bringing the total initial funding to $59 million for the first six launches the contract required, not including the costs of a guidance/control sys­tem, Atlas boosters, and a launch complex. Ehricke believed that, until it was too late, the limited funding restricted the necessary ground testing his project engineers could do. Also restrictive was the absence of the DoD’s highest priority (known as DX), which meant that subcontractors who were also working on projects with the DX priority could not give the same level of service to Centaur as they provided to higher-priority projects.11

Under these circumstances, Convair and Pratt & Whitney pro­ceeded with designs for the Centaur structure and engines. The Centaur stage used the steel-balloon structure of Atlas, with the same 10-foot diameter. The lightness of the resulting airframe seemed necessary for Centaur because of liquid hydrogen’s low den­sity, which made the hydrogen tank much larger than the oxygen tank. Conventional designs with longerons and ring frames would have created a less satisfactory mass fraction than did the pressur­ized tanks with thin skins (initially only 0.01 inch thick). The el­liptical liquid oxygen tank was on the bottom of the stage. To create the shortest possible length and the lowest weight, the engineers on Ehricke’s project team made the bottom of the liquid-hydrogen tank concave so that it fit over the convex top of the oxygen tank.

This arrangement solved space and weight problems (saving about 4 feet of length and roughly 1,000 pounds of weight) but created oth-

ers in the process. One resulted from the smallness of the hydrogen molecules and their extreme coldness. The skin of the oxygen tank had a temperature of about — 299°F, which was so much “warmer" than the liquid hydrogen at — 423°F that the hydrogen would gasify from the relative heat and boil off. To prevent that, the engineers de­vised a bulkhead between the two tanks that contained a fiberglass – covered Styrofoam material about 0.2 inch thick in a cavity between two walls. Technicians evacuated the air from the pores in the Styro­foam and refilled the spaces with gaseous nitrogen. They then welded the opening. When they filled the upper tank with liquid hydrogen, the upper surface of the bulkhead became so cold that it froze the nitrogen in the cavity, thus creating a vacuum as the nitrogen con­tracted into the denser solid state, a process called cryopumping.12

Because of the limited testing, it was not until the summer and early fall of 1961 that the Centaur engineers and managers learned of heat transfer across the bulkhead that was more than 50 times the amount expected. It turned out that there were very small cracks in the bulkhead through which the hydrogen was leaking and destroy – 178 ing the vacuum, causing the heat transfer and resultant boil-off of Chapter 5 the fuel. This necessitated venting to avoid excessive pressure and explosion in the hydrogen tank. But the venting depleted the fuel, leaving an insufficient quantity for the second engine burn required of Centaur for coasting in orbit and then propelling a satellite into a higher orbit.

General Dynamics had used Atlas manufacturing techniques for the materials on the bulkhead. Atlas’s quality-control procedures permitted detection of leaks in bulkheads down to about 1/10,000 inch. Inspections revealed no such leaks, but the engineers learned in the 1961 testing that hydrogen could escape through even finer openings. Very small cracks that would not be a problem in a liquid – oxygen tank caused major leakage in a liquid-hydrogen tank.13

By the time Convair-Astronautics had discovered this problem, NASA had assigned responsibility for the Centaur project to the Marshall Space Flight Center (on July 1, 1960), with Seaberg’s Cen­taur Project Office remaining at BMD in California. Hans Hueter became director of Marshall’s Light and Medium Vehicles Office in July, with responsibility for managing the Centaur and Agena upper stages. During the winter of 1959-60, NASA also established a Cen­taur technical team following the cancellation of the Vega project. This team consisted of experts at various NASA locations to rec­ommend ways the upper stage could be improved. In January 1960, navy commander W. Schubert became the Centaur project chief at NASA Headquarters.14

From December 11 to 14, 1961, John L. Sloop visited General Dy­namics/ Astronautics (GD/A) to look into Centaur problems, par­ticularly the one with heat transfer across the bulkhead. Sloop had been head of Lewis Laboratory’s rocket research program from 1949 until 1960, when he moved to NASA Headquarters. There in 1961 he became deputy director of the group managing NASA’s small and medium-sized launch vehicles. Following his visit, he wrote, “GD/A has studied the problem and concluded that it is not practi­cal to build bulkheads where such a vacuum [as the one Ehricke’s team had designed] could be maintained." The firm also believed “that the only safe way to meet all Centaur missions is to drop the integral tank design and go to separate fuel and oxidizer tanks." Sloop disagreed: “If a decision must be made now, I recommend we stick to the integral tank design, make insulation improvements, and lengthen the tanks to increase propellant capacity."15

Подпись:Sloop’s optimism was justified. After the Centaur team began “a program of designing and testing a number of alternate designs," tests revealed that adding nickel to the welding of the double bulk­head (and elsewhere), significantly increased the single-spot shear strength of the metal at —423°F.16

Centaur development experienced many other problems. Several of them involved the engines. After enduring “inadequate facilities, slick unpaved roads, mosquitoes, alligators, and 66 inches of rain in a single season" while developing the 304 engine for Suntan at West Palm Beach, Pratt & Whitney engineers also “discovered the slippery nature of hydrogen." The extreme cold of liquid hydrogen precluded using rubber gaskets to seal pipe joints, designers hav­ing to resort to aluminum coated with Teflon and then forced into flanges that mated with them. There had to be new techniques for seals on rotating surfaces, where carbon impregnated with silver found wide use. Another concern with the cryogenic hydrogen was that the liquid not turn to gas before reaching the turbopumps. The engineers initially solved that problem by flowing propellants to the pumps before engine start, precooling the system.17

The turbopump for the 304 engine used oil to lubricate its bear­ings. This had to be heated to keep it from freezing in proximity to the cold pump, creating a temperature gradient. To solve this prob­lem for the RL10, the Pratt & Whitney engineers coated the cages holding the bearings with fluorocarbons similar to Teflon and ar­ranged to keep the bearings cold with minute amounts of liquid hydrogen. This produced the same effect as lubrication, because it turned out that the main function of oil was to keep the bearings from overheating. The substance from which Pratt & Whitney nor-

mally made its gears, called Waspalloy, bonded in the hydrogen en­vironment. Engineers replaced it with carbonized steel coated with molybdenum disulfide for dry lubrication. This solved the bonding problem but subjected some unlucky engineers to observing tests of the new arrangement by using binoculars from an observation post with only a screen door. Late at night, alligator croakings and other noises created uneasiness for many young observers unused to swamp sounds.18

The first component tests of the combustion chamber for the RL10, including stainless-steel regenerative-cooling tubes brazed with silver, took place in May 1959. As with many other initial tests of combustion chambers, there were signs of burnthrough, so the engineers changed the angle at which the hydrogen entered the tubes and aligned the tubes more carefully so they did not pro­trude into the exhaust stream. Engine firings two months after this showed that the changes had solved the burnthrough problem, but the chamber’s conical shape produced inefficient burning. Engi­neers changed the design to a bell shape and conducted a successful 180 engine run in September 1959, less than a year from the date of the Chapter 5 initial contract.19

A major innovation in the design of the RL10 took advantage of the cold temperature of liquid hydrogen in order to dispense with a gas generator to drive the turbopump. The cryogenic fuel passed from the tank into the tubes of the combustion chamber for cooling. As it did so, it absorbed heat, which caused the fuel to vaporize and expand. This provided enough kinetic energy to drive the turbine that operated both the liquid-hydrogen and liquid-oxygen pumps. It also provided the power for the hydraulic actuators that gimballed the engine. This process, called the “bootstrap" cycle, still used hy­drogen-peroxide boost pumps to start the process. Hydrogen per­oxide also powered attitude-control rockets and ullage-control jets that propelled the Centaur forward in a parking orbit and thereby forced the liquid hydrogen to the rear of the tanks. There it could be pumped into the engines for ignition.20

Before the RL10 underwent its first test in an upright position on a test stand in its two-engine configuration for the Centaur, it under­went 230 successful horizontal firings. It produced 15,000 pounds of thrust and achieved a specific impulse of about 420 lbf-sec/lbm at an expansion ratio of 40:1 through its exhaust nozzle. As required by its missions in space, it reliably started, stopped, and restarted so that it could coast in a parking orbit until it reached the optimum point for injection into an intended orbit (or trajectory for interplan­etary voyages). On November 6, 1960, two RL10s, upright for the

first time on a test stand at the Pratt & Whitney facility in Florida, fired at the same time and did so successfully—for a short time un­til a problem occurred with the timer on the test stand. When en­gineers repeated the test the next day, only one engine fired. The other filled with hydrogen and oxygen until the flame from the first engine caused an explosion that damaged the entire propulsion sys­tem beyond repair.

Подпись:A tape recording of the countdown suggested that the problem had stemmed from the faulty operation of a test-stand sequencer, so engineers did not suspect difficulties with the engine itself. By Janu­ary 12, 1961, they repaired the test stand and tested another pair of engines. This time, they put a blast wall between the two engines and installed a shutoff valve on the hydrogen tank. They also sepa­rated the exhaust systems for the two engines by a greater distance. During this test, there was no problem with the sequencing, but the explosion recurred. In the vertical position, engineers learned, grav­ity was affecting the mixing of the oxygen with the hydrogen differ­ently than it had in the horizontal position. So in a further instance of cut-and-try engineering, designers had to adjust the method of hydrogen feed. They also designed a method of measuring the den­sity of the mixture to ensure the presence of enough oxygen for igni­tion. With these adjustments, the two engines fired simultaneously in the vertical test stand on April 24, 1961. Following this success, the engines completed 27 successful dual firings at Pratt & Whitney and 5 more at the rocket site on Edwards AFB in California. They then passed the flight-rating test from October 30 to November 4, 1961, in which they completed 20 firings equivalent in duration to six Centaur missions.21

To protect the liquid hydrogen in its tank from boiling off while the vehicle was on the launching pad and during ascent through the atmosphere, engineers had designed four jettisonable insulation panels made of foam-filled fiberglass. These were about a centime­ter (0.39 inch) thick, held on the tank by circumferential straps. To keep air from freezing between the tank and the insulating foam, thereby bonding the panels to the tank, engineers designed a helium system to purge the air. To limit the weight penalty imposed by the panels (1,350 pounds), they had to be jettisoned as soon after launch as the atmosphere thinned and the ambient temperature dropped.22

Because of delays resulting from the engine ignition problem, dif­ficulties with elaborate test instrumentation (such as a television camera and sensors inside the liquid-hydrogen tank), and other is­sues, an Atlas LV-3 with a Centaur upper stage did not launch for the first time until May 8, 1962, 15 months later than planned. The

Centaur PropulsionTHRUST- 15,000 LB (ALTITUDE) THRUST DURATION-470SEC SPECIFIC IMPULSE-433SEC ENGINE WT DRY-298 LB EXIT-TO THROAT AREA RATIO – 40 ТОЇ

Подпись:Подпись:PROPELLANTS-LOX & LH2 PROPELLANT FLOW RATE – 35 LB/SEC

CONTRACTOR-

PRATT & WHITNEY SYSTEM-SAT I/S-1V (6 ENGINES) CENTAUR (2 ENGINES)

l-RM-D IND 8І4ЮВ

goals of the test flight were to proceed through the boost phase with jettison of the insulation and a nose fairing, followed by Centaur’s separation from the Atlas. With only a partial load of fuel, the Cen­taur was to coast for 8 minutes, reignite, and burn for 25 seconds.23

On the launch, the two stages rose normally until they ap­proached maximum dynamic pressure (with resultant aerodynamic buffeting) as the vehicle got close to the speed of sound 54.7 sec­onds into the launch. Then, an explosion occurred as the liquid – hydrogen tank split open. Initially, engineers decided that aerody­namic forces had destroyed the insulation and ruptured the tank. About five years later, tests suggested that the real culprit was dif­ferential thermal expansion between a fiberglass nose fairing and the steel tank, causing a forward ring to peel off the tank.24

Even before this launch, the difficulties with engine develop­ment, resultant schedule delays, and problems such as the one with the bulkhead between the hydrogen and oxygen tanks had led to close scrutiny of the Centaur project and danger of its cancellation. Following John Sloop’s visit to General Dynamics to look into such problems, he had expressed concerns about the firm’s organization. Krafft Ehricke, the program director, had only five men reporting directly to him, and Deane Davis, the project engineer, had direct charge of only two people. Many other people worked on Centaur (27 of them full-time), but most of them were assigned to six oper-

ating divisions not directly under project control. Sloop wrote, “As far as I could tell in three days of discussion, the only people who have direct and up-to-date knowledge of all Centaur systems are Mr. Ehricke and Mr. Davis." Marshall Space Flight Center had “a very competent team of four men stationed at GD/A," and they were well aware of the “management deficiencies" emphasized in Sloop’s comments.25

Hans Hueter wrote on January 4, 1962, to GD/A president James Dempsey stating his concern about the way the Centaur Program Office was organized in “relation to the line divisions." He men­tioned that the two of them had discussed this issue “several times" and reiterated his and other NASA employees’ “impression that the systems engineering is carried on singlehandedly by your ex­cellent associates, Krafft Ehricke and Dean [sic] Davis." He added, “The individual fields such as propulsion, thermal and liquid be­havior, guidance and control, and structures are covered in depth in the various engineering departments but coordination is sorely lacking."26

Подпись:In response to NASA’s concerns about this matrix organization, Dempsey shifted to a “projectized" arrangement in which roughly 1,100 employees at Astronautics were placed under the direct au­thority of the Centaur program director. Ehricke was reassigned as the director of advanced systems and Grant L. Hansen became Cen­taur program director and Astronautics vice president on February 1, 1962. Trained as an electrical engineer at Illinois Institute of Tech­nology, Hansen had worked for Douglas Aircraft from 1948 to 1960 on missile and space systems, including the Thor, with experience in analysis, research and development, design, and testing. He came to GD/A in 1960 to direct the work of more than 2,000 people on Atlas and Centaur. After February 1962, Ehricke continued to offer Hansen advice. Although he was imaginative and creative, the com­pany had decided Ehricke “wasn’t enough of a[n] S. O.B. to manage a program like this." Hansen proved to be effective, although it is only fair to note that he was given authority and an organization Ehricke had lacked.27 S. O.B. or not, had Ehricke started with Hansen’s or­ganization and adequate funding, Centaur development could have been smoother from the beginning. In any event, this sequence of events showed how management arrangements and technical prob­lems interacted.

Several other programmatic changes occurred around this time. On January 1, 1962, for example, NASA (in agreement with the DoD) transferred the Centaur Project Office from Los Angeles to Huntsville, Alabama, and converted existing air force contracts to

NASA covenants. Lieutenant Colonel Seaberg ceased being proj­ect manager, and Francis Evans at Marshall Space Flight Center as­sumed those duties under Hueter’s direction. By this time, funding had grown from the original $59 million to $269 million, and the number of Centaur vehicles to be delivered had risen from 6 to 10.28

Meanwhile, following the May 8, 1962, explosion, a congressional Subcommittee on Space Sciences, chaired by Rep. Joseph E. Karth (D-Minnesota), began hearings on the mishap. In a report issued on July 2, 1962, the parent Committee on Science and Astronautics in the U. S. House of Representatives stated that “management of the Centaur development program has been weak and ineffective both at NASA headquarters and in the field."29 NASA did not immediately make further changes, but Marshall management of Centaur posed problems. These came out in the hearings, prompting unfavorable comment in the committee report. Von Braun had remarked about GD/A’s “somewhat bold approach. In order to save a few pounds, they have elected to use some rather, shall we say, marginal solu­tions where you are bound to buy a few headaches before you get 184 it over with." Hansen agreed that his firm was inclined “to take a Chapter 5 little bit more of a design gamble to achieve a significant improve­ment, whereas I think they [Marshall engineers] build somewhat more conservatively." The congressional report noted, “Such a dif­ference in design philosophy can have serious consequences."30

Ehricke characterized the design approach of the von Braun team as “Brooklyn Bridge" construction. The contrast between that and the approach of General Dynamics appears in an account of a Mar­shall visit to GD/A that Deane Davis wrote at an unspecified date soon after Marshall took over responsibility for Centaur in July 1960. A group led by von Braun and including Hueter and structures chief William Mrazek had come to GD/A for a tour and briefings on Atlas and Centaur. Mrazek and Bossart had gotten into a discussion of the structure of the steel-balloon tanks, with Mrazek (according to Davis’s account) unwilling to admit that they could have any structural strength without ribs. Bossart took him out to a tank and handed him a fiberglass mallet containing lead to give it a weight of 7 pounds. It had a rubber cover and a 2-foot handle. Bossart invited Mrazek to hit the tank with it. After a tap and then a harder whack, he could not find a dent. Bossart urged him to “stop fiddling around. Hit the damned thing!" When Mrazek gave it a “smart crack," the mallet bounced back so hard it flew about 15 feet, knocking off the German’s glasses on the way and leaving only a black smear (no dent) on the tank. Davis wrote that Hueter was as amazed as Mrazek by the strength of the tank.31

This account is difficult to accept entirely at face value because Mrazek had already designed the Redstone with an integral-tank structure that was hardly as light as Bossart’s steel balloon but was also not quite bridgelike. Nevertheless, even in 1962 von Braun was clearly uncomfortable with Bossart’s “pressure-stabilized tanks," which he called “a great weightsaver, but. . . also a continuous pain in the neck" that “other contractors, for example the Martin Co., for this very reason have elected not to use." No doubt because of such concerns, von Braun sought quietly to have the Centaur can­celed in favor of a Saturn-Agena combination.32

Подпись:Faced with this situation, on October 8, 1962, NASA Headquar­ters transferred management of the Centaur program to the Lewis Research Center, to which Silverstein had returned as center di­rector in 1961 from his position at NASA Headquarters. A “sharp, aggressive, imaginative, and decisive leader," Silverstein could be “charming or abrasive," in the words of John Sloop. Deane Davis, who worked with him on Centaur, called him a “giant among gi­ants" and a man he “admired, adored, hated, wondered about—and mostly always agreed with even when I fought him. Which was of­ten." Under Silverstein’s direction, the Lewis center required much more testing than even the Marshall group had done. Lewis tested everything that could “possibly be proven by ground test." Yet de­spite such aggressive oversight, Grant Hansen expressed admiration for Lewis and its relationship with his own engineers.33

Because the RL10 had been planned for use on Saturn as well as Centaur, its management remained at Marshall. The reason given for Centaur’s transfer was that it would allow the Huntsville engi­neers to concentrate on the Saturn program. A NASA news release quoted NASA administrator James Webb, “This, I feel, is necessary to achieve our objectives in the time frame that we have planned. It will permit the Lewis Center to use its experience in liquid hydro­gen to further the work already done on one of the most promising high energy rocket fuels and its application to Centaur. . . ."34

Long before this transfer, engineers from the Cleveland facility had been actively involved in helping solve both engine and struc­tural problems with the vehicle. Their involvement included use of an altitude chamber at their center. Other facilities, including a rocket sled track at Holloman AFB, New Mexico, had also been involved in Centaur development. For example, in 1959 GD/A had done some zero-G testing in an air force C-131D aircraft at Wright- Patterson AFB (and also, at some point, in a KC-135). The same year, the firm had acquired a vacuum chamber for testing gas ex­pansion and components. With additional funding (to a total of

about $63 million) in 1960, GD/A extended testing to include use of the vacuum test facility at the air force’s Arnold Engineering Development Center in Tullahoma, Tennessee, zero-G test flights using Aerobee rockets, and additional static ground testing, includ­ing modifying test stand 1-1 at the rocket site on Edwards AFB for Centaur’s static tests. In 1961, when GD/A’s funding rose to $100 million, there were wind-tunnel tests of the Centaur’s insulation panels at NASA’s Langley Research Center, additional zero-G test­ing, and construction of a coast-phase test stand to evaluate the attitude-control system.35

At Lewis, Silverstein decided to direct the Centaur project him­self, assisted by two managers under his personal direction and some 41 people involved with technical direction. Some 40 Mar­shall engineers helped briefly with the program’s transition. By Janu­ary 1963, the changeover was mostly complete and Centaur had acquired a DX priority. Then, costs for Centaur were estimated at $350 million, and containing them became an issue. Despite this, Silverstein decided that the first eight Centaurs after the transfer 186 would constitute test vehicles. By this time, Surveyor spacecraft Chapter 5 had been assigned as Centaur payloads, and Silverstein determined that none of them would be launched until the test vehicles had demonstrated Centaur’s reliability.36

By February 1963, Silverstein had appointed David Gabriel as Centaur manager but placed the project office in the basement of his own administrative building so he could continue to keep tabs on the project. Some continuity with the period of Marshall man­agement came in the retention of Ronald Rovenger as chief of the NASA field office at GD/A. Instead of 4, his office rose to a comple­ment of 40 NASA engineers. It took until April 1964, but Lewis renegotiated the existing contracts with GD/A into a single cost- plus-fixed-fee document for 14 Centaur upper stages plus 21 test articles. The estimated cost of the agreement was roughly $321 mil­lion plus a fixed fee of $31 million, very close to the estimate of $350 million at the beginning of 1963. However, Silverstein felt the need for a second contract to cover further modifications resulting from Lewis’s technical direction. Soon the Lewis staff working on Centaur grew to 150 people. Silverstein continued to give the proj­ect his personal attention and made a major decision to abandon temporarily the use of a parking orbit and restart for Surveyor. This required a direct ascent to the Moon, considerably narrowing the “window" for each launch.37

These and other changes under Lewis direction did not imme­diately solve all of Centaur’s problems. Test flights and resultant

Date

Mission

Objective

Outcome

Nov. 27,

R&D,

Achieve separation

Successful,

1963

single-burn

of Centaur, Earth

achieved

orbit, data on nose

orbit close to

cone, insulation

that planned,

panels

gathered data

June 30,

R&D,

Test jettison

Jettison

1964

single-burn,

of redesigned

successful

restart

insulation panels

but failure of

boost

and nose cone,

driveshaft in

pumps

gather data from

hydraulic pump

restart

prevented

gimballing

Dec. 11,

R&D,

Restart engines,

Partial success;

1964

two-burn

carry Surveyor

first burn

model

successful but ullage motors not powerful enough to keep LH2 at bottom of tank;a weak

restart

Mar. 2,

R&D,

Simulate Surveyor

Failed; Atlas fuel

1965

single-burn,

launch

valve closed,

separable

causing an

Surveyor

model

explosion

Aug. 11,

R&D,

Demonstrate

Successful in

1965

single-burn,

capability of

separating

separable

launching Surveyor

model and

Surveyor

model similar to

sending on

model

actual spacecraft

planned course

Apr. 7,

R&D,

Perform 25-minute

Partial failure;

1966

two-burn,

coast in parking

in parking

separable

orbit, re-ignite

orbit there was

Surveyor

Centaur engine,

a hydrogen

model

and send Surveyor

peroxide leak

model to a target

and too little

location simulating

remained to

the Moon

power tank boost pumps

Flight

AC-2

AC-3

AC-4

AC-5

AC-6

AC-8

(continued)

Flight

Date

Mission

Objective Outcome

AC-9

Oct. 26,

R&D,

Demonstrate Successful

1966

two-burn,

restart capability,

separable

send Surveyor

Surveyor

model on

model

simulated

trajectory to Moon

aLH2 is liquid hydrogen.

difficulties are summarized in table 5.1, beginning with Atlas – Centaur 2 (AC-2).38

Data from instrumentation on the insulation panels over the liq­uid-hydrogen tank on AC-2 showed conclusively that the design for the panels used on AC-1 was not adequate. Engineers designed thicker panels with heavier reinforcement, increasing their weight 188 by almost 800 pounds. This made it all the more important to jet – Chapter 5 tison them at about 180 seconds after launch to get rid of the un­wanted weight. A minor redesign fixed the problem with the drive – shaft that failed on AC-3. To fix the problem on AC-4 with liquid hydrogen moving away from the bottom of the tank where the fuel had to exit, however, required investigation and multiple modifica­tions. A slosh baffle in the liquid-hydrogen tank helped limit move­ment of the fuel away from the tank bottom. Screens in the ducts bringing bleed-off hydrogen gas back to the tank reduced energy that could disturb the liquid. On the coasting portion of AC-4’s orbit, liquid hydrogen had gotten into a vent intended to exhaust gaseous hydrogen, thereby releasing pressure from boil-off. The liq­uid exiting into the vacuum of space created a sideward thrust that tumbled the Centaur and Surveyor models. Fixing this problem re­quired a complete redesign of the venting system.

A further change increased thrust in both the yaw – and pitch – control engines as well as those that settled liquid hydrogen in the bottom of the tank during coast. The added thrust in both types of engines helped keep the Centaur on course and hold the easily dis­placed liquid hydrogen in the bottom of its tank. Fortunately, these changes were unnecessary before the launch of AC-5 but were im­plemented for AC-8, which also incorporated the uprated RL10A – 3-3 engine with slightly greater specific impulse from a larger expansion ratio for the exhaust nozzle and an increased chamber pressure.39

Meanwhile, in response to the explosion on AC-5, engineers locked the Atlas valves in the open position. AC-6 amounted to a semioperational flight. The Surveyor model went to the coordi­nates in space it was intended to reach (simulating travel to the Moon) even without a trajectory correction in midcourse. With AC-7 shifted to a later launch and AC-8 having problems with hy­drogen peroxide rather than liquid hydrogen, the Atlas-Centaur combination was ready for operational use, although there would be one more research-and-development flight sandwiched between launches of operational spacecraft (AC-9; see table 5.1). Atlas – Centaur performed satisfactorily on all of the Surveyor launches, although two of the spacecraft had problems. But five of the seven missions were successful, providing more than 87,000 photographs and much scientific information for Apollo landings and lunar stud­ies. Surveyors 1, 2, and 4 all used single-burn operations by Centaur, but Surveyors 3 and 5-7 employed dual-burn trajectories. On Sur­veyors 5-7 the Atlases were all SLV-3Cs with longer tanks, hence greater propellant volumes. The SLV-3C flew only 17 missions but was successful on all of them before being replaced by the SLV-3D, used with the advanced Centaur D-1A.40

Подпись:The D-1A resulted from a NASA decision to upgrade the Cen­taur, with the Lewis Research Center responsible for overseeing the $40 million improvement program, the central feature of which was a new guidance/control computer, developed at a cost of about $8 million. Among payloads for the Centaur D-1A were Intelsat com­munications satellites. With the first launch of Intelsat V, having more relay capacity (and weight), on December 6, 1980, the Centaur began to use engines that were adjusted to increase their thrust (per engine) from the original 15,000 to about 16,500 pounds. The 93.75 percent success rate for the 32 SLV-3D/D-1A (and D-1AR) launches showed that Silverstein’s insistence on extended testing and detailed oversight had paid off.41

During the early 1980s, General Dynamics converted to new ver­sions of Atlas and Centaur. The Atlas G added 81 inches to the length of the propellant tanks, and Pratt & Whitney made several changes to the Centaur engines, including removal of the boost pump, for a significant weight savings. There was no change in the RL10’s thrust, but further modification shifted from hydrogen per­oxide to the more stable hydrazine for the attitude-control and pro­pellant-settling engines. This made the RL10A-3-3A a substantially different machine than its predecessor, the RL10A-3-3.42

As of early 1991, the Centaur had had a 95 percent success rate on 76 flights. This included 42 successes in a row for Centaur D-1

and D-1A between 1971 and 1984. The vehicle, as well as its Atlas booster, would continue to evolve into the 21st century, with the successful launch of an Atlas V featuring a Russian RD-180 engine and a Centaur with a single RL10 engine, signifying both the end of the cold war and the continuing evolution of the technology. Meanwhile, development of the Centaur had led to the use of liq­uid-hydrogen technology both on upper stages of the Saturn launch vehicle and on the Space Shuttle. Despite a difficult start and con­tinuing challenges, the Centaur had made major contributions to U. S. launch-vehicle technology.43

Analysis and Conclusions

The development of missiles and rockets for DoD needs arguably contributed to national defense and, through deterrence, kept the

cold war from becoming hotter than it actually got in Korea, Viet­nam, and Afghanistan, among other places. For the purposes of this book, however, the importance of the missiles and rockets discussed in this chapter lay in the technology that could transfer to launch – vehicle uses. In many cases, actual missiles, with some adaptations, became either launch vehicles or stages in larger combinations of rockets used to place satellites or spacecraft on their trajectories. Without the perceived urgency created by cold-war concerns and without the heterogeneous engineering of missile proponents, it conceivably would have taken much longer for launch vehicles to develop, although many satellites themselves were high on the DoD’s priority lists.

Подпись: 47 German and U.S. Missiles and Rockets, 1926-66

Quite apart from their contributions to launch-vehicle technol­ogy, the missiles and rockets discussed in this chapter also illustrate many of the themes that will be further explored in subsequent chapters. Missiles such as the Titan II and Minuteman showed the ways in which technology for earlier missiles contributed to their successors. Although this chapter provides only an overview of mis­sile development, it shows several examples of trial-and-error engi­neering that was necessary to overcome often unforeseen problems. Clearly, the missiles discussed here required a wide range of talents and a huge number of different organizations to design and develop them. Also important was a considerable sharing of information, even between competing organizations and firms. Finally, manage­ment systems such as the one Schriever adopted at WDD (and a similar system called Program Evaluation and Review Technique [PERT] adopted by Raborn for the Polaris program) enabled very complicated missiles and launch vehicles to be developed reason­ably on time and in such a way that all component systems (such as propulsion, structures, guidance and control) worked together effectively.

Подпись: U.S. Space- Launch Vehicles, 1958-91 LAUNCH VEHICLES FREQUENTLY USED MIS­siles as first stages, but these required many modi­fications, particularly when they had to boost hu­mans into space. Even for satellite and spacecraft launches, technology for the booster stages fre­quently represented modification of technologies missiles needed for their ballistic paths from one part of Earth to another. Thus, the history of the Thor-Delta, Atlas, Scout, Saturn, Titan, and Space Shuttle launch vehicles differed from, but remained

dependent on, the earlier development of the missiles discussed in chapter 1. Missiles and launch vehicles represented a continuum, with many of the same people contributing to both. But they re­mained different enough from one another to require separate treat­ment in this chapter.

Despite the differences, launch-vehicle development exhibited many of the same themes that characterized missiles. It featured the same engineering culture that relied heavily on extensive test­ing on the ground. But this did not always succeed in revealing all problems that occurred in flight. When unexpected problems oc­curred, it was not always possible for engineers to understand the exact causes. But they were able to arrive at fixes that worked. There continued to be a wide range of organizations and disciplines that contributed to launch-vehicle development, including the solution of unanticipated problems. Also characteristic of launch vehicles was a competitive environment that nevertheless featured sharing of information among organizations involved in development. In part, this sharing occurred through the movement of knowledge­able engineers from one organization to another. More often, the information sharing (plus its recording and validation) occurred through professional societies, papers delivered at their meetings, and publication of reports in professional journals.1 Finally, mis­siles and launch vehicles shared the use of management systems that tracked development of components to ensure that all of them occurred on schedule and that they all worked together effectively.

Polaris Propulsion

Meanwhile, the navy’s Polaris missile had made more far-reaching contributions. Until Polaris A1 became operational in 1960, all U. S.

long-range missiles had used liquid propellants. These had obvious advantages in their performance, but their extensive plumbing and large propellant tanks made protecting them in silos difficult and costly. Such factors also made them impractical for use on ships. After the operational date of Minuteman I in 1962, the Department of Defense began phasing out liquid-propellant strategic missiles.34

Подпись:Meanwhile, given the advantages that liquid propellants en­joyed in terms of performance, their head start within the defense establishment, and the disinclination of most defenders of liquids to entertain the possibility that solid propellants could satisfy the demanding requirements of the strategic mission, how did this solid-propellant breakthrough occur? The answer is complicated and technical. But fundamentally, it happened because a number of heterogeneous engineers promoted solids; a variety of partners in their development brought about significant technical innovations; and although interservice rivalries encouraged the three services to development separate missiles, interservice cooperation ironically helped them do so. Despite such cooperation and the accumulat­ing knowledge about rocket technology, however, missile designers still could not foresee all the problems that their vehicles would develop during ground and flight testing. Thus, when problems did occur, rocket engineers still had to gather information about what had caused problems and exercise their ingenuity to develop solu­tions that would cope with the unexpected.

By the time that Polaris got under way in 1956 and Minuteman in 1958, solid-propellant rocketry had already made the tremen­dous strides forward discussed previously. But there were still enor­mous technical hurdles to overcome. The problems remaining to be solved included higher performance; unstable combustion; the inadequate durability of existing nozzle materials under conditions of heat and exposure to corrosive chemicals from the exhaust of the burning propellants; a lack of materials and technology to provide light but large combustion chambers so the burning propellants had to overcome less mass during launch; and ways to terminate com­bustion of the propellant immediately after the desired velocity had been achieved (for purposes of accuracy) and to control the direction of the thrust (for steering).35

Once the navy had overcome the bureaucratic obstacles to devel­oping its own, solid-propellant missile, the Special Project Office (SPO) under Adm. William F. Raborn and Capt. Levering Smith achieved breakthroughs in a number of these technical areas. In early January 1956, the navy had sought the assistance of the Lock­heed Missile and Space Division and the Aerojet General Corpora-

tion in developing a solid-propellant ballistic missile. The initial missile the two contractors and the SPO conceived was the Jupiter S (for “solid"). It had enough thrust to carry an atomic warhead the re­quired distance, a feat it would achieve by clustering six solid rock­ets in a first stage and adding one for the second stage. The problem was that Jupiter S would be about 44 feet long and 10 feet in diam­eter. An 8,500-ton vessel could carry only 4 of them but could carry 16 of the later Polaris missiles. With Polaris not yet developed, the navy and contractors still were dissatisfied with Jupiter S and con­tinued to seek an alternative.36

One contribution to a better solution came from Atlantic Re­search Corporation (ARC). Keith Rumbel and Charles B. Hender­son, chemical engineers with degrees from MIT who were working 240 for ARC, had begun theoretical studies in 1954 of how to increase Chapter 6 solid-propellant performance. They learned that other engineers, in­cluding some from Aerojet, had calculated an increase in specific impulse from adding aluminum powder to existing ingredients. But these calculations had indicated that once aluminum exceeded 5 per­cent of propellant mass, performance would again decline. Hence, basing their calculations on contemporary theory and doing the cumbersome mathematics without the aid of computers, the other researchers abandoned aluminum as an additive except for damping combustion instability. Refusing to be deterred by theory, Rumbel and Henderson tested polyvinyl chloride with much more alumi­num in it. They found that with additional oxygen in the propellant and a flame temperature of at least 2,310 kelvin, a large percent­age of aluminum by weight yielded a specific impulse significantly higher than that of previous composite propellants.37

ARC’s polyvinyl chloride, however, did not serve as the binder for Polaris. Instead, the binder used was a polyurethane material developed by Aerojet in conjunction with a small nitropolymer pro­gram funded by the Office of Naval Research about 1947 to seek high-energy binders for solid propellants. A few Aerojet chemists synthesized a number of high-energy compounds, but the process required levels of heating that were unsafe with potentially explo­sive compounds. Then one of the chemists, Rodney Fischer, found “an obscure reference in a German patent" suggesting “that iron chelate compounds would catalyze the reaction of alcohols and iso­cyanates to make urethanes at essentially room temperature." This discovery started the development of polyurethane propellants in many places besides Aerojet.

In the meantime, in 1949 Karl Klager, then working for the Of­fice of Naval Research in Pasadena, suggested to Aerojet’s parent

firm, General Tire, that it begin work on foamed polyurethane, leading to two patents held by Klager, along with Dick Geckler and R. Parette of Aerojet. In 1950, Klager began working for Aerojet. By 1954, he headed the rocket firm’s solid-propellant development group. Once the Polaris program began in December 1956, Klager’s group decided to reduce the percentage of solid oxidizer as a compo­nent of the propellant by including oxidizing capacity in the binder, using a nitromonomer as a reagent to produce the polyurethane plus some inert polynitro compounds as softening agents. In April 1955, the Aerojet group found out about the work of Rumbel and Hender­son. Overcoming explosions due to cracks in the grain and profiting from other developments from multiple contributors, they discov­ered successful propellants for both stages of Polaris A1.

Подпись:These consisted of a cast, case-bonded polyurethane composition including different percentages of ammonium perchlorate and alu­minum for stages one and two, both of them featuring a six-point, internal-burning, star configuration. With four nozzles for each stage, this propellant yielded a specific impulse of almost 230 lbf – sec/lbm for stage one and nearly 255 lbf-sec/lbm for stage two. The latter specific impulse was higher in part because of the reduced atmospheric pressure at the high altitudes where it was fired, com­pared with stage one, which was fired at sea level.38

The addition of aluminum to Aerojet’s binder essentially solved the problem of performance for Polaris. Other innovations in the areas of warhead size plus guidance and control were necessary to make Polaris possible, but taken together with those for the propel­lants, they enabled Polaris A1 to be only 28.6 feet long and 4.5 feet in diameter (as compared with Jupiter S’s 44 feet and 10 feet, respec­tively). The weight reduction was from 162,000 pounds for Jupiter S to less than 29,000 pounds for Polaris. The cases for both stages of Polaris A1 consisted of rolled and welded steel. This had been modified according to Aerojet specifications resulting from exten­sive metallurgical investigations.

Each of the four nozzles for stage one (and evidently, stage two as well) consisted of a steel shell, a single-piece throat of molybdenum, and an exit-cone liner made of “molded silica phenolic between steel and molybdenum." A zirconium oxide coating protected the steel portion. The missile’s steering came from jetavators designed by Willy Fiedler of Lockheed, a German who had worked on the V-1 program during World War II and had developed the concept for the device while employed by the U. S. Navy at Point Mugu Naval Air Missile Test Center, California. He had patented the idea and then adapted it for Polaris. Jetavators for stage one were molybdenum

FIG. 6.4

Polaris PropulsionAn unidentified solid-rocket motor being tested in an altitude wind tunnel at NASA’s Lewis Research Center (later Glenn Research Center) in 1959, one kind of test done for Polaris. (Photo courtesy of NASA)

rings with spherical inside surfaces that rotated into the exhaust stream of the four nozzles and deflected the flow to provide pitch, yaw, and roll control. The jetavators for stage two were similar.39

Besides requiring steering, the missile needed precise thrust ter­mination when it reached the correct point in its trajectory. This could be achieved on liquid-propellant missiles simply by stopping the flow of propellants. For solids, the task was more difficult. The Polaris team used pyrotechnics to blow out plugs in six ports in front of the second stage at the proper moment in the trajectory. This per­mitted exhaust gases to escape and halt the acceleration so that the warhead would travel on a ballistic path to the target area.40

Flight testing of Polaris revealed, among other problems, a loss of control due to electrical-wiring failure at the base of stage one. This resulted from aerodynamic heating and a backflow of hot exhaust gases. To diagnose and solve the problem, engineers in the program obtained the help of “every laboratory and expert," using data from four flights, wind tunnels, sled tests, static firings, “and a tremen­dous analytic effort by numerous laboratories." The solution placed fiberglass flame shielding supplemented by silicone rubber over the affected area to shield it from hot gases and flame.41

Another problem for Polaris to overcome was combustion in­stability. Although this phenomenon is still not fully understood,
gradually it has yielded to research in a huge number of institutions, including universities and government labs, supported by funding by the three services, the Advanced Research Projects Agency, and NASA. Levering Smith credited Edward W. Price in the Research Department at NOTS with helping to understand the phenomenon. By this time, Price had earned a B. S. in physics and math at UCLA. In February 1960, he completed a major (then-classified) paper on combustion instability, which stated, “This phenomenon results from a self-amplifying oscillatory interaction between combustion of the propellant and disturbances of the gas flow in the combustion chamber." It could cause erratic performance, even destruction of motor components. Short of this, it could produce vibrations that would interfere with the guidance/control system. To date, “only marginal success" had been achieved in understanding the phenom­enon, and “trial-and-error development continues to be necessary." But empirical methods gradually were yielding information, for example, that energy fluxes could amplify pressure disturbances, which had caused them in the first place.42 The subsequent success of Polaris showed that enough progress had been made by this time that unstable combustion would not be a major problem for the missile.

Подпись:Long before Polaris A1 was operational, in April 1958 the DoD had begun efforts to expand the missile’s range from the 1,200- nautical-mile reach of the actual A1 to the 1,500 nautical miles origi­nally planned for it. The longer-range missile, called Polaris A2, was originally slated to achieve the goal through higher-performance propellants and lighter cases and nozzles in both stages. But the navy Special Projects Office decided to confine these improvements to the second stage, where they would have greater effect. (With the second stage already at a high speed and altitude when it began fir­ing, it did not have to overcome the weight of the entire missile and the full effects of the Earth’s gravity at sea level.) Also, in the sec­ond stage, risk of detonation of a high-energy propellant after igni­tion would not endanger the submarine. Hence, the SPO invited the Hercules Powder Company to propose a higher-performance second stage.43

As a result, Aerojet provided the first stage for Polaris A2, and Her­cules, the second. Aerojet’s motor was 157 inches long (compared with 126 inches for Polaris A1; the additional length could be ac­commodated by the submarines’ launch tubes because the navy had them designed with room to spare). It contained basically the same propellant used in both stages of Polaris A1 with the same grain con­figuration. Hercules’ second stage had a filament-wound case and a

cast, double-base grain that contained ammonium perchlorate, ni­trocellulose, nitroglycerin, and aluminum, among other ingredients. The grain configuration consisted of a 12-point, internal-burning star. It yielded a specific impulse of more than 260 lbf-sec/lbm under firing conditions. The motor was 84 inches long and 54 inches in diameter, featuring four swiveling nozzles with exit cones made of steel, asbestos phenolic, and Teflon plus a graphite insert.44

This second-stage motor resulted from an innovation that in­creased performance by adding ammonium perchlorate to the cast, double-base process used in Hercules’ third stage for the Vanguard launch vehicle. Hercules’ ABL developed this new kind of propel­lant, known as composite-modified double base (CMDB), by 1958, evidently with the involvement of John Kincaid and Henry Shuey, 244 developers of the earlier cast, double-base process.45

Chapter 6 Even before 1958, however, Atlantic Research Corporation had developed a laboratory process for preparing CMDB. In its manu­factured state, nitrocellulose is fibrous and unsuitable for use as an unmodified additive to other ingredients being mixed to create a propellant. Arthur Sloan and D. Mann of ARC, however, developed a process that dissolved the nitrocellulose in nitrobenzene and then separated out the nitrocellulose by mixing it with water under high shear (a process known as elutriation). The result was a series of compact, spherical particles of nitrocellulose with small diameters (about 1 to 20 microns). Such particles combined readily with liq­uid nitroplasticizers and crystalline additives in propellant mixers. The result could be cast into cartridge-loaded grains or case-bonded rocket cases and then converted to a solid with the application of moderate heat. Sloan and Mann patented the process and assigned it to ARC. Then, in 1955, Keith Rumbel and Charles Henderson at ARC began scaling the process up to larger grain sizes and devel­oping propellants. They developed two CMDB formulations begin­ning in 1956. When ARC’s pilot plant became too small to support the firm’s needs, production shifted to Indian Head, Maryland. Be­cause the plastisol process they had developed was simpler, safer, and cheaper than other processes then in existence, Henderson said that Hercules and other producers of double-base propellants even­tually adopted his firm’s basic method of production.

Engineers did not use it for upper stages of missiles and launch vehicles until quite a bit later, however, and then only after chem­ists at several different laboratories had learned to make the pro­pellant more rubberlike by extending the chains and cross-linking the molecules to increase the elasticity. Hercules’ John Allabashi at ABL began in the early 1960s to work on chain extenders and

cross-linking, with Ronald L. Simmons at Hercules’ Kenvil, New Jersey, plant continuing this work. By about the late 1960s, chem­ists had mostly abandoned use of plastisol nitrocellulose in favor of dissolving nitrocellulose and polyglycol adipate together, followed by a cross-linking agent such as isocyanate. The result was the type of highly flexible CMDB propellant used on the Trident submarine- launched ballistics missiles beginning in the late 1970s.46

Подпись:Meanwhile, the rotatable nozzles on the second stage of Polaris A2, which were hydraulically operated, were similar in design to those already being used on the air force’s Minuteman I, and award of the second-stage contract to Hercules reportedly resulted from the performance of the third-stage motor Hercules was developing for Minuteman I, once again illustrating technology transfer be­tween services. (Stage one of the A2 retained the jetavators from A1.) The A2 kept the same basic shape and guidance/control system as the A1, the principal change being more reliable electronics for the guidance/control system. By the time Polaris A2 became opera­tional in June 1962, now-Vice Admiral Raborn had become deputy chief of naval operations for research and development. In February

1962, Rear Adm. I. J. “Pete" Gallantin became director of the Spe­cial Projects Office with Rear Adm. Levering Smith remaining as technical director.47

As a follow-on to Polaris A2, in September 1960, Secretary of De­fense Robert McNamara approved development of a 2,500-nautical – mile version of Polaris that became the A3. To create a missile that would travel an additional 1,000 nautical miles while being launched from the same tubes on the submarine as the A1 required new propellants and a higher mass fraction. The new requirement also resulted in a change from the “bottle shape" of the A1 and A2 to a shape resembling a bullet. In the attempt to help increase the mass fraction, Aerojet, the first-stage manufacturer, acquired the Houze Glass Corporation at Point Marion, Pennsylvania, and moved that firm’s furnaces, patents, technical data, and personnel to the Aerojet Structural Materials Division in Azusa, California, in early

1963. This acquisition gave Aerojet the capability to make filament – wound cases like those used by Hercules on stage two of Polaris A2. The new propellant Aerojet used was a nitroplasticized polyurethane containing ammonium perchlorate and aluminum, configured with an internal-burning, six-point star. This combination raised the spe­cific impulse less than 10 lbf-sec/lbm.48

Unfortunately, the flame temperature of the new propellant was so high that it destroyed the nozzles on the first stage. Aerojet had to reduce the flame temperature from a reported 6,300°F to slightly

less than 6,000°F and to make the nozzles more substantial, using silver-infiltrated tungsten throat inserts to withstand the high tem­perature and chamber pressure. As a result, the weight saving from using the filament-wound case was lost in the additional weight of the nozzle. Hence, the mass fraction for the A3’s first stage was actually slightly lower than that for the A2, meaning that the same quantity of propellant in stage one of the A3 had to lift slightly more weight than did the lower-performing propellant for stage one of the A2.49

For stage two of the A3, Hercules used a propellant containing an energetic high explosive named HMX, a smaller amount of am­monium perchlorate, nitrocellulose, nitroglycerin, and aluminum, among other ingredients. It configured this propellant into an in – 246 ternal-burning cylindrical configuration with many major and mi – Chapter 6 nor slots in the aft end of the stage, creating a cross section that resembled a Christmas-tree ornament. This propellant offered a significantly higher specific impulse than stage two of Polaris A2. Also, the new stage two used a different method of achieving thrust vector control (steering). It injected Freon into the exhaust, creating a shock pattern to deflect the stream and achieve the same results as movable nozzles at a much smaller weight penalty.

A further advantage of this system was its lack of sensitivity to the temperature of the propellant flame. The Naval Ordnance Test Station performed the early experimental work on this use of a liq­uid for thrust vector control. Aerojet, Allegany Ballistics Laboratory, and Lockheed did analytical work, determined the ideal locations for the injectors, selected the fluid to be used, and developed the injectors as well as a system for expelling excess fluid. The Polaris A3 team first successfully tested the new technology on the sec­ond stage of flight A1X-50 on September 29, 1961. This and other changes increased the mass fraction for stage two of Polaris A3 to 0.935 (from 0.871 for stage two of the A2), a major improvement that together with the increased performance of the new propellant, contributed substantially to the greater range of the new missile.50