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

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.

Chapters 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.

Most 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.

Besides 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

evolved 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.

Soon after the engineers at Wallops began developing Scout, those at the Redstone Arsenal started work on the Saturn family of launch vehicles. Whereas the solid-propellant Scout was the runt among American launch vehicles, the liquid-fueled Saturn became the gi­ant. The standard Scout at the end of its career was a four-stage vehicle about 75 feet in height, but the Saturn V, with only three stages, stood almost 5 times as tall, about 363 feet. This made the Saturn taller than the Statue of Liberty—equivalent in height to a 36-story building and taller than a football field was long. Com­posed of some 5 million parts, it was a complex mass of propellant tanks, engines, plumbing fixtures, guidance/control devices, thrust structures, and other elements. Its electrical components, for exam­ple, included some 5,000 transistors and diodes, all of which had to be tested both individually and in conjunction with the rest of the vehicle to ensure that they would work properly when called upon. The Scout could launch only 454 pounds into a 300-mile orbit, but the Saturn V sent a roughly 95,000-pound payload with three astro­nauts on board toward six successful lunar orbits.64

Using clustering of rocket engines in its lower stages to achieve its massive initial thrust, the Saturn V was based on the earlier development of vehicles that ultimately came to be designated the Saturn I and the Saturn IB. These two interim space boosters were, in turn, based upon technologies developed for the Redstone, Ju­piter, Thor, Atlas, Centaur, and other vehicles and stages. Thus, although the Saturn family constituted the first group of rockets developed specifically for launching humans into space, it did not so much entail new technologies as it did a scaling up in size of 74 existing or already developing technology and an uprating of engine Chapter 2 performance. Even so, this posed significant technological hurdles and, often, a need to resort to empirical solutions to problems they raised. Existing theory and practice were inadequate for both the massive scale of the Saturns and the need to make them sufficiently reliable to carry human beings to an escape trajectory from Earth’s immediate vicinity out toward the Moon.65

The organization under Wernher von Braun at the Army Ballistic Missile Agency (ABMA)—which became NASA’s Marshall Space Flight Center (MSFC) in mid-1960—began the initial research and

development of the Saturn family of vehicles. In response to DoD projections of a need for a very large booster for communication and weather satellites and space probes, ABMA had begun in April 1957 to study a vehicle with 1.5 million pounds of thrust in its first stage. Then the formation of the Advanced Research Projects Agency (ARPA) on February 7, 1958, led ABMA to shift from initial plans to use engines still under development. Despite the crisis resulting from Sputnik, ARPA urged the use of existing and proven engines so that the new booster could be developed as quickly as possible at a minimal cost. ABMA then shifted to use of eight uprated Thor – Jupiter engines in a cluster to provide the 1.5 million pounds of thrust in the first stage, calling the new concept the Juno V. This led to an ARPA order on August 15, 1958, initiating what von Braun and others soon started calling the Saturn program, a name ARPA officially sanctioned on February 3, 1959.66

Under a contract signed on September 11, 1958, the Rocketdyne Division of North American Aviation in fact supplied an H-1 engine that was actually more than an uprated Thor-Jupiter powerplant. It resulted from research and development on an X-1 engine that the Experimental Engines Group at Rocketdyne began in 1957. Mean­while, the engineers at ABMA did some scrounging in their stock of leftover components to meet the demands of ARPA’s schedule within their limited budget. The schedule called for a full-scale, static firing of a 1.5-million-pound cluster by the end of 1959. In­stead of a new, single tank for the first stage, which would have required revised techniques and equipment, the ABMA engineers found rejected or incomplete 5.89-foot (in diameter) Redstone and 8.83-foot Jupiter propellant tanks. They combined one of the Jupi­ter tanks with eight from the Redstone to provide a cluster of pro­pellant reservoirs for the first-stage engines. In such a fashion, the Saturn program got started, with funding gradually increased even before the transfer of the von Braun group to NASA.67

Initially called the Saturn C-1, the Saturn I (properly so-called after February 1963) was at first going to have a third stage, but between January and March 1961, NASA decided to drop the third stage and to use Pratt & Whitney RL10 engines in the second stage. These were the same engines being developed for Centaur. Mean­while, on April 26, 1960, NASA had awarded a contract to the Doug­las Aircraft Company to develop the Saturn I second stage, which confusingly was called the S-IV. Unlike the Centaur, which used only two RL10s, the S-IV held six of the engines, requiring consider­able scaling up of the staging hardware. Using its own experience as well as cooperation from Centaur contractors Convair and Pratt &

Whitney, Douglas succeeded in providing an SL-IV stage in time for the January 29, 1964, first launch of a Saturn I featuring a live second stage. This was also the first launch with 188,000-pound – thrust H-1 engines in the first stage, and it succeeded in orbiting the second stage.68

Development of the Saturn I posed problems. Combustion insta­bility in the H-1 engines, stripped gears in an H-1 turbopump, slosh­ing in first-stage propellant tanks, and an explosion during static testing of the S-IV stage all required redesigns. But apart from the sloshing on the first Saturn I launch (SA-1), the 10 flights of Saturn I (from October 27, 1961, to July 30, 1965) revealed few problems. There were changes resulting from flight testing, but NASA counted all 10 flights successful, a tribute to the thoroughness and extensive ground testing of von Braun’s engineers and their contractors.69

At 191.5 feet tall (including the payload), Saturn I was still a far cry from Saturn V. The intermediate version, Saturn IB, consisted of a modified Saturn I with its two stages (S-IB and S-IVB) redesigned to reflect the increasing demands placed upon them, plus a further developed instrument unit with a new computer and additional flexibility and reliability. The S-IVB would serve as the second stage of the Saturn IB and (with further modifications) the third stage of the Saturn V, exemplifying the building-block nature of the devel­opment process. The first stage of the Saturn IB was also a modified S-I, built by the same contractor, Chrysler. The cluster pattern for the eight H-1 engines did not change, although uprating increased their thrust to 200,000 and then 205,000 pounds per engine. In its second stage, the Saturn IB had a new and much larger engine, the J-2, with thrust exceeding that of the six RL10s used on the Saturn I. Like the RL10s, it burned liquid hydrogen and liquid oxygen.70

Rocketdyne won a contract (signed on September 10, 1960) to develop the J-2, with specifications that the engine ensure safety for human flight yet have a conservative design to speed up avail­ability. By the end of 1961, it had become evident that the engine 76 would power not only stage two of Saturn IB but both the second Chapter 2 and third stages of the Saturn V. In the second stage of Saturn V, there would be a cluster of five J-2s; on the S-IVB second stage of Saturn IB and the S-IVB third stage of Saturn V, there would be a single J-2 a piece. Rocketdyne engineers had problems with injec­tors for the new engine until they borrowed technology from the RL10, a further example of shared information between competing contractors, facilitated by NASA.71

After completion of its development, the Saturn IB stood 224 feet high. Its initial launch on February 26, 1966, marked the first flight

tests of an S-IVB stage, a J-2 engine, and a powered Apollo space­craft. It tested two stages of the launch vehicle plus the reentry heat shield of the spacecraft. Except for minor glitches like failures of two parachutes for data cameras, it proved successful. Other flights also succeeded, but on January 27, 1967, a ground checkout of the vehicle for what would have been the fourth flight test led to the di­sastrous Apollo fire that killed three astronauts. The test flights of Saturn IB concluded with the successful launching and operation of the command and service modules (CSM), redesigned since the fire. Launched on October 11, 1968, the Saturn IB with a 225,000-pound – thrust J-2 in the second stage performed well in this first piloted Apollo flight. This ended the Saturn IB flights for Apollo, although the vehicle would go on to be used in the Skylab and Apollo-Soyuz Test Projects from 1973 to 1975.72

Development of some parts used exclusively for the Saturn V be­gan before design of other components common to both the Saturn IB and the Saturn V. For instance, on January 9, 1959, Rocketdyne won the contract for the huge F-1 engine used on the Saturn V (but not the IB); however, not until May 1960 did NASA select Rocket – dyne to negotiate a contract for the J-2 common to both launch ve­hicles. Configurations were in flux in the early years, and NASA did not officially announce the C-1B as a two-stage vehicle for Earth – orbital missions with astronauts aboard until July 11, 1962, renam­ing it the Saturn IB the next February. (NASA Headquarters had already formally approved the C-5 on January 25, 1962.) Thus, even though the Saturn IB served as an interim configuration between the Saturn I and the Saturn V, development of both vehicles over­lapped substantially, with planning for the ultimate moon rocket occurring even before designers got approval to develop the interim configuration.73

Burning RP-1 as its fuel with liquid oxygen as the oxidizer, the F-1 did not break new technological ground—in keeping with NASA guidelines to use proven propellants. But its thrust level required so much scaling up of the engine as to mark a major advance in the state of the art of rocket making. Perhaps the most intricate design feature of the F-1, and certainly one that caused great difficulty to engineers, was the injection system. As with many other engines, combustion instability was the problem. On June 28, 1962, during an F-1 hot-engine test at the rocket site on Edwards AFB, combus­tion instability led to the meltdown of the engine. Using essentially trial-and-error methods coupled with high-speed instrumentation and careful analysis, a team of engineers had to test perhaps 40 or 50 design modifications before they eventually found a combination of

baffles, enlarged fuel-injection orifices, and changed impingement angles that worked.74

Despite all the effort that went into the injector design, accord­ing to Roger Bilstein, it was the turbopump that “absorbed more design effort and time for fabrication than any other component of the engine." There were 11 failures of the system during the de­velopment period. All of them necessitated redesign or a change in manufacturing procedures. The final turbopump design provided the speed and high volumes needed for a 1.5-million-pound-thrust engine with a minimal number of parts and high ultimate reliabil­ity. However, these virtues came at the expense of much testing and frustration.75

There were many other engineering challenges during design and development of the Saturn V. This was especially true of the S-II second stage built by the Space and Information Systems Division of North American Aviation. Problems with that segment of the huge rocket delayed the first launch of the Saturn V from August until November 9, 1967. But on that day the launch of Apollo 4 (flight AS-501) without astronauts aboard occurred nearly without a flaw.76

After several problems on the second Saturn V launch (includ­ing the pogo effect on the F-1 engines) prevented the mission from being a complete technical success, engineers found solutions. As a result, the third Saturn V mission (AS-503, or Apollo 8) achieved a successful circumlunar flight with astronauts aboard in late 1968. Following two other successful missions, the Apollo 11 mission between July 16, 1969, and July 24, 1969, achieved the first of six successful lunar landings with the astronauts returning to Earth, fulfilling the goals of the Apollo program.77

With Saturns I and IB as interim steps, Saturn V was the culmi­nation of the rocket development work von Braun’s engineers had been carrying on since the early 1930s in Germany. In the interim, the specific engineers working under the German American baron 78 had included a great many Americans. There had been a continual Chapter 2 improvement of technologies from the V-2 through the Redstone, Jupiter, and Pershing missiles to the three Saturn launch vehicles.

Not all of the technologies used on Saturn V came from von Braun’s engineers, of course. Many technologies in the Saturn V resulted from those developed on other programs in which von Braun’s team had not participated or for which it was only partly responsible. This is notably true of much liquid-hydrogen technol­ogy, which stemmed from contributions by Lewis Research Cen­ter, Convair, Pratt & Whitney, Rocketdyne, and Douglas, among

others—showing the cumulative effects of much information shar­ing. But Marshall engineers worked closely with the contractors for the J-2, S-II, and S-IVB stages in overcoming difficulties and made real contributions of their own. This was also true in the develop­ment of the Saturn engines. Rocketdyne had started its illustrious career in engine development by examining a V-2 and had worked with von Braun and his engineers on the Redstone engine, a pro­cess that continued through Jupiter and the Saturn engines. But a great many of the innovations that led to the F-1 and J-2 had come more or less independently from Rocketdyne engineers, and even on the major combustion-instability and injector problems for the F-1, Rocketdyne’s contributions seem to have been at least as great as those from Marshall engineers. In other words, it took teamwork, not only among Americans and Germans at Marshall but among Marshall, other NASA centers, industry, universities, and the U. S. military to create the Apollo launch vehicles. Air force facilities and engineers at Edwards AFB, Holloman AFB, and the Arnold En­gineering Development Center also made key contributions to fac­ets of Saturn development.

Another key ingredient in the success of Saturn rocket develop­ment was the management system used at ABMA and the NASA Marshall Space Flight Center. As at Peenemunde, von Braun re­tained his role as an overall systems engineer despite other com­mitments on his time. At frequent meetings he chaired, he con­tinued to display his uncanny ability to grasp technical details and explain them in terms everyone could understand. Yet he avoided monopolizing the sessions, helping everyone to feel part of the team. Another technique to foster communication among key technical people was his use of weekly notes. Before each Monday, he required his project managers and laboratory chiefs to submit one-page summaries of the previous week’s developments and prob­lems. Each manager had to gather and condense the information. Then von Braun wrote marginal comments and circulated copies back to the managers. He might suggest a meeting between two in­dividuals to solve a problem or himself offer a solution. Reportedly, the roughly 35 managers were eager to read these notes, which kept them all informed about problems and issues and allowed them to stay abreast of overall developments, not just those in their separate areas. In this way, the notes integrated related development efforts and spurred efforts to solve problems across disciplinary and orga­nizational lines.78

Superficially, these “Monday notes" seemed quite different from Schriever’s “Black Saturdays" and Raborn’s PERT system. They

were certainly less formal and more focused on purely technical so­lutions than on cost control. But in the technical arena, von Braun’s system served the same systems-engineering function as the other systems.

While the Saturn I was undergoing development and flight test­ing, significant management changes occurred in NASA as a whole. From November 18, 1959 (when NASA assumed technical direc­tion of the Saturn effort), through March 16, 1960 (when the space agency took over administrative direction of the project and formal transfer took place), to July 1, 1960 (when both the Saturn program and the von Braun team of engineers transferred to the Marshall Space Flight Center), the administrator of NASA remained the capable and forceful but conservative T. Keith Glennan. Glennan had organized NASA, adding JPL and Marshall to the core centers inherited from the National Advisory Committee for Aeronautics, NASA’s predecessor.79

Once John F. Kennedy became president in early 1961 and ap­pointed the still more forceful and energetic but hardly conservative James E. Webb to succeed Glennan, there were bound to be man­agement changes. This was further ensured by Kennedy’s famous exhortation on May 25, 1961, “that this nation. . . commit itself to achieving the goal, before this decade is out, of landing a man on the Moon and returning him safely to earth." The commitment that followed gave an entirely new urgency to the Saturn program. To coordinate it and the other aspects of the Apollo program, NASA re­organized in November 1961. Even before the reorganization, Webb chose as head of a new Office of Manned Space Flight (OMSF) an engineer with Radio Corporation of America who had been project manager for the Ballistic Missile Early Warning System (BMEWS). D. Brainerd Holmes had finished the huge BMEWS project on time and within budget, so he seemed an ideal person to achieve a simi­lar miracle with Apollo.80

Holmes headed one of four new program offices in the reorga – 80 nized NASA Headquarters, with all program and center directors Chapter 2 now reporting to Associate Administrator Robert C. Seamans Jr., who also took over control of NASA’s budget. Webb apparently had not fully grasped Holmes’s character when he appointed him. The second NASA administrator had previously considered Abe Sil – verstein to head OMSF and rejected him because he wanted too much authority, especially vis-a-vis Seamans. Holmes, however, turned out to be “masterful, abrasive, and determined to get what he needed to carry out his assignment, even at the expense of other programs." Within two weeks of joining NASA, the confrontational

FIG. 2.10

Wernher von Braun (center) showing the Saturn launch system to Pres. John F. Kennedy and Deputy NASA Administrator Robert Seamans (left) at Cape Canaveral, November 16, 1963. (Photo courtesy of NASA)

as unlikely that a lunar landing could be achieved during Kenne­dy’s decade “with acceptable risk." They believed it would be late 1971 before a landing could occur. Mueller took the two men to Seamans’s office to repeat the findings. Seamans then told Mueller privately to find out how to get back on schedule, exactly the au­thority and leverage Mueller had evidently sought.83

On November 1, 1963, Mueller then instituted two major changes that offered a way to land on the Moon by the end of the decade and greatly strengthened his position. One change was all-up testing. In 1971 Mueller claimed he had been involved with the development of all-up testing at Space Technology Laboratories, although he may or may not have known that his organization had opposed the idea when Otto Glasser introduced it as the only way he could conceive to cut a year from development time for Minuteman at the insis­tence of the secretary of the air force. In any event, all-up testing had worked for Minuteman and obviously offered a way to speed up testing for the Saturn vehicles.84

NASA defined all-up testing to mean a vehicle was “as complete as practicable for each flight, so. . . a maximum amount of test information is obtained with a minimum number of flights." This conflicted with the step-by-step procedures preferred by the von Braun group, but on November 1, Mueller sent a priority message to Marshall, the Manned Spacecraft Center (MSC) in Houston, and the Launch Operations Center (LOC) in Florida. In it he announced a deletion of previously planned Saturn I launches with astronauts aboard and directed that the first Saturn IB launch, SA-201, and the first Saturn V flight, AS-501, should use “all live stages" and in­clude “complete spacecraft." In a memorandum dated October 26, 1963, Mueller wrote to Webb via Seamans, enclosing a proposed NASA press release about all-up testing: “We have discussed this course of action with MSFC, MSC, and LOC, and the Directors of these Centers concur with this recommendation," referring specifi­cally to eliminating “manned" Saturn I flights but by implication, 82 to the all-up testing. The press release stated that “experience in Chapter 2 other missile and space programs" had shown it to be “the quickest way of reaching final mission objectives" (a further example of how shared information was important to missile and launch-vehicle programs).85

If Mueller had really discussed all-up testing with the center di­rectors, this was not apparent at Marshall, where von Braun went over the message with his staff on November 4, creating a “furor." Staff members recalled numerous failed launches in the V-2, Red­stone, and Jupiter programs. William A. Mrazek believed the idea

of all-up testing was insane; other lab heads and project managers called it “impossible" and a “dangerous idea." Although von Braun and his deputy, Eberhard Rees, both had their doubts about the idea, in the end they had to agree with Mueller that the planned launches of individual stages would prevent landing on the Moon by the end of the decade. All-up testing prevailed despite von Braun’s doubts.86 And once again, it worked.

The second change introduced on November 1 entailed a reor­ganization of NASA, placing the field centers under the program offices once again, rather than under Seamans. Mueller obtained authority over Marshall, the MSC, and the LOC (renamed Kennedy Space Center in December).87

One aspect of the Marshall effort that did not fit with Mueller’s management concepts was the proclivity for technical decisions in Huntsville to be based on their merits instead of schedule or cost. This was true even though project managers were supposed to get jobs done “on time and within budget." A concern with time, bud­get, and what had come to be called configuration control, however, had become very important in the Minuteman program and quickly spread to NASA when Mueller arranged for Brig. Gen. Samuel Phil­lips to come to NASA as deputy director and then director (after Oc­tober 1964) of the Apollo program. The slender, handsome Phillips had moved from his post as director of the Minuteman program in August 1963 to become vice commander of the Ballistic Systems Division. He arrived at NASA Headquarters in early January 1964 and soon arranged for the issue in May of a NASA Apollo Configu­ration Management Manual, adapted from an air force manual.88

Phillips had expected resistance to configuration management from NASA. He was not disappointed. Mueller had formed an Apollo Executive Group consisting of the chief executives of firms working on Apollo plus directors of NASA field centers, and in June 1964, Phillips and a subordinate who managed configuration control for him presented the system to the assembled dignitaries. Von Braun objected to the premises of Phillips’s presentation on the ground that costs for development programs were “very much unknown, and configuration management does not help." He contended that it was impossible for the chairman of a configuration control board to know enough about all the disciplines involved to decide intel­ligently about a given issue. Phillips argued that if managers were doing their jobs, they were already making such decisions.

Von Braun retorted that the system tended to move decisions higher in the chain of management. William M. Allen of Boeing coun­tered that this was “a fundamental of good management." When von

Braun continued to argue the need for flexibility, Mueller explained that configuration management did not mean that engineers had “to define the final configuration in the first instance before [they knew] that the end item [was] going to work." It meant defining the ex­pected design “at each stage of the game" and then letting everyone know when it had to be changed. Center directors like von Braun did not prevail in this argument, but resistance continued in industry as well as NASA centers, with the system not firmly established until about the end of 1966.89

Mueller and Phillips introduced other management procedures and infrastructure to ensure control of costs and schedules. Phillips converted from a system in which data from the field centers came to Headquarters monthly to one with daily updates. He quickly con­tracted for a control room in NASA Headquarters similar to the one he had used for Minuteman, with data links to field centers. A com­puterized system stored and retrieved the data about parts, costs, and failures. Part of this system was a NASA version of the navy’s Program Evaluation and Review Technique (PERT), developed for Polaris, which most prime contractors had to use for reporting cost and scheduling data.90

Despite von Braun’s resistance to configuration management, Phillips recalled in 1988, “I never had a single moment of problem with the Marshall Space Flight Center. [Its] teamwork, cooperation, enthusiasm, and energy of participation were outstanding." Phillips added, “Wernher directed his organization very efficiently and par­ticipated in management decisions. When a decision had been made, he implemented it—complied, if you will, with directives." In part, no doubt, Phillips was seeing through the rose-colored glasses of memory. But in part, this reflected von Braun’s propensity to argue until he was either convinced that the contrary point of view was correct or until he saw that argument was futile. Then he became a team player.

Phillips had been at the receiving end of V-2s in England during 84 World War II, and he was prepared to dislike the German baron who Chapter 2 had directed their development. But the two became friends. He commented that von Braun “could make a person feel personally important to him and that [his or her] ideas were of great value." When asked about von Braun’s contributions to the space program, Phillips observed that a few years before, he probably would have said that American industry and engineering could have landed on the Moon without German input. But in 1988 he said, “When I think of the Saturn V, which was done so well under Wernher’s direction and which was obviously. . . essential to the lunar mis-

sion, . . . I’m not sure today that we could have built it without the ingenuity of Wernher and his team."91

The contributions of Mueller and Phillips were probably more critical to the ultimate success of Apollo than even von Braun’s. Phillips hesitated about characterizing Mueller but did say that “his perceptiveness and ability to make the right decision on important and far-reaching [as well as] complex technical matters" was “pretty unusual." On the other hand, Mueller’s biggest shortcoming, ac­cording to Phillips, was “dealing with people." John Disher, who admired Silverstein but characterized Mueller as the “only bona – fide genius I’ve ever worked with," had to observe that although his boss was “always the epitome of politeness, . . . deep down he [was] just as hard as steel." Also, human space program director of flight operations, Chris Kraft, who dealt with a great many people and frequently clashed with Mueller, had to say that “I’ve never dealt with a more capable man in terms of his technical ability." Difficult though Mueller was, without him and Phillips, American astronauts probably would not have gotten to the Moon before the end of the decade.92

Both men brought to NASA some of the culture and many of the management concepts used in the air force, plus the navy’s PERT system. These added to the amalgam of cultures already present in the space agency. Inherited from the National Advisory Com­mittee for Aeronautics (NACA) was a heavy research orientation coupled with a strong emphasis on testing. Von Braun’s Germans had brought a similar culture from their V-2 and army experience, together with a tendency to over-engineer rocket systems to ensure against failure. JPL had added its own blend of innovation plus a strong reliance on theory and use of mathematics that was also part of the V-2 experience. The Lewis Research Center brought to the mix its experience with liquid hydrogen, and the Langley Research Center brought the testing of rockets at Wallops Island and a wind – tunnel culture (present also at Ames Research Center). All of these elements (and the management styles that accompanied them) con­tributed to the Saturn-Apollo effort.

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

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

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.

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

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

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

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

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

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

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

The history of space-launch-vehicle technology in the United States dates back to the experimenting of U. S. physicist and rocket developer Robert H. Goddard (1882-1945). A fascinating character, Goddard was supremely inventive. He is credited with 214 patents, many of them submitted after his death by his wife, Esther. These led to a settlement in 1960 by the National Aeronautics and Space Administration (NASA) and the three armed services of $1 million for use of more than 200 patents covering innovations in the fields of rocketry, guided missiles, and space exploration. In the course of his rocket research, Goddard achieved many technological break­throughs. Among them were gyroscopic control of vanes in the ex­haust of the rocket engine, film cooling of the combustion chamber, parachutes for recovery of the rocket and any instruments on it, streamlined casing, clustered engines, a gimballed tail section for stabilization, lightweight centrifugal pumps to force propellants into the combustion chamber, a gas generator, igniters, injection heads, and launch controls, although he did not use them all on any one rocket.1

Despite these impressive achievements, Goddard had less de­monstrable influence on the development of subsequent missiles and space-launch vehicles than he could have had. One reason was that he epitomized the quintessential lone inventor. With excep – 8 tions, he pursued a pattern of secrecy throughout the course of his Chapter 1 career. This secretiveness hindered his country from developing missiles and rockets as rapidly as it might have done had he devoted his real abilities to the sort of cooperative development needed for the production of such complex devices.

Educated at Worcester Polytechnic Institute (B. S. in general science in 1908) and Clark University (Ph. D. in physics in 1911), Goddard seems to have begun serious work on the development of rockets February 9, 1909, when he performed his first experiment on the exhaust velocity of a rocket propellant. He continued experi­mentation and in 1916 applied to the Smithsonian Institution for $5,000 to launch a rocket within a short time to extreme altitudes (100-200 miles) for meteorological and other research. He received a grant for that amount in 1917. From then until 1941 he received a total of more than $200,000 for rocket research from a variety of civilian sources.2

In 1920 he published “A Method of Reaching Extreme Altitudes" in the Smithsonian Miscellaneous Collections. As Frank Winter has stated, this “publication established Goddard as the preemi­nent researcher in the field of rocketry" and “was unquestionably very influential in the space travel movement. . . ."3 However, im­portant and pathbreaking as the paper was, it remained largely theoretical, calling for “necessary preliminary experiments" still to be performed.4 Following the paper’s publication, with partial hiatuses occasioned by periods of limited funding, Goddard spent the rest of the interwar period performing these experiments and trying to construct a rocket that would achieve an altitude above that reached by sounding balloons.

After experiencing frustrating problems using solid propellants, Goddard switched to liquid propellants in 1921. But it was not until March 26, 1926—nine years after his initial proposal to the Smith­sonian—that he was able to achieve the world’s first known flight of a liquid-propulsion rocket at the farm of Effie Ward, a distant relative, in Auburn, Massachusetts. Goddard continued his rocket research in the desert of New Mexico after 1930 for greater isola­tion from human beings, who could reveal his secrets as well as be injured by his rockets. But when he finally turned from develop­ment of high-altitude rockets to wartime work in 1941, the highest altitude one of his rockets had reached (on March 26, 1937) was estimated at between 8,000 and 9,000 feet—still a long way from his stated goals.5

One reason he had not achieved the altitudes he originally sought was that he worked with a small number of technicians instead of cooperating with other qualified rocket engineers. He achieved sig­nificant individual innovations, but he never succeeded in design­ing and testing all of them together in a systematic way so that the entire rocket achieved the altitudes he sought. Trained as a scien­tist, Goddard failed to follow standard engineering practices.6

More important than this shortcoming was his unwillingness to publish technical details of his rocket development and testing. At the urging of sponsors, he did publish a second paper, titled “Liq­uid-Propellant Rocket Development," in 1936 in the Smithsonian Miscellaneous Collections. There, Goddard addressed, much more explicitly than in his longer and more theoretical paper of 1920, the case for liquid-propellant rockets, stating their advantages over powder rockets—specifically their higher energy. Although he did discuss some details of the rockets he had developed and even in­cluded many pictures, in general the rather low level of detail and the failure to discuss many of the problems he encountered at every step of his work made this paper, like the earlier one, of limited usefulness for others trying to develop rockets.7

In 1948, Esther Goddard and G. Edward Pen dray did publish his notes on rocket development. These contained many specifics missing from his earlier publications, but by that time the Germans under Wernher von Braun and his boss, Walter Dornberger, had developed the A-4 (V-2) missile, and a group at the Jet Propulsion Laboratory (JPL) in Pasadena, California, had also advanced well be-

yond Goddard in developing rockets and missiles. He patented and developed a remarkable number of key innovations, and the two pa­pers he did publish in his lifetime significantly influenced others to pursue rocket development. But both the Germans under von Braun and Dornberger and the U. S. effort at JPL demonstrated in varying degrees that it took a much larger effort than Goddard’s to achieve the ambitious goals he had set for himself.

Because of Goddard’s comparative secrecy, Romanian-German rocket theoretician Hermann Oberth (1894-1989), oddly, may have contributed more to U. S. launch-vehicle technology than his American counterpart. Unlike Goddard, Oberth openly published 12 the details of his more theoretical findings and contributed to their Chapter 1 popularization in Germany. Because of these efforts, he was signifi­cantly responsible for the launching of a spaceflight movement that directly influenced the V-2 missile. Then, through the immigration of Wernher von Braun and his rocket team to the United States af-

ter World War II, Oberth contributed indirectly to U. S. missile and spaceflight development.

Born almost 12 years after Goddard on June 25, 1894, in the partly Saxon German town of Hermannstadt, Transylvania, Oberth attended a number of German universities but never earned a Ph. D. because none of his professors would accept his dissertation on rocketry. Undaunted by this rejection, Oberth nevertheless “re­frained from writing another" dissertation on a more acceptable and conventional topic.8

He succeeded in publishing Die Rakete zu den Planetenraumen (The Rocket into Interplanetary Space) in 1923. Although Goddard always suspected that Oberth had borrowed heavily from his 1920 paper,9 in fact Oberth’s book bears little resemblance to Goddard’s paper. Not only is Die Rakete much more filled with equations but it is also considerably longer than the paper—some 85 pages of smaller print than the 69 pages in Goddard’s paper as reprinted by the Amer­ican Rocket Society in 1946. Oberth devoted much more attention than Goddard to such matters as liquid propellants and multiple – stage rockets, whereas the American dealt mostly with solid pro­pellants and atmospheric studies but did mention the efficiency of hydrogen and oxygen as propellants. Oberth also set forth the basic principles of spaceflight to a greater extent than Goddard had done in a work much more oriented to reporting on his experimental re­sults than to theoretical elaboration. Oberth discussed such matters as liquid-propellant rocket construction for both alcohol and hydro­gen as fuels; the use of staging to escape Earth’s atmosphere; the use of pumps to inject propellants into the rocket’s combustion cham­ber; employment of gyroscopes for control of the rocket’s direction; chemical purification of the air in the rocket’s cabin; space walks; microgravity experiments; the ideas of a lunar orbit, space stations, reconnaissance satellites; and many other topics.10

The book itself was influential. Besides writing it, Oberth collab­orated with Max Valier, an Austrian who wrote for a popular audi­ence, to produce less technical writings that inspired a great deal of interest in spaceflight.11 According to several sources, Oberth’s first book directly inspired Wernher von Braun (the later technical direc­tor of the German Army Ordnance facilities at Peenemunde where the V-2 was developed, subsequently director of NASA’s Marshall Space Flight Center) to study mathematics and physics, so necessary for his later work. Von Braun had already been interested in rocketry but was a poor student, especially in math and physics, in which he had gotten failing grades. However, in 1925 he had seen an ad for Oberth’s book and ordered a copy. Confronting its mathematics,

he took it to his secondary school math teacher, who told him the only way he could understand Oberth was to study his two worst subjects. He did and ultimately earned a Ph. D. in physics.12 Without Oberth’s stimulation, who knows whether von Braun would have become a leader in the German and U. S. rocket programs?

Similarly, von Braun’s boss at Peenemunde, Walter Dornberger, wrote to Oberth in 1964 that reading his book in 1929 had opened up a new world to him. And according to Konrad Dannenberg, who had worked at Peenemunde and come to the United States in 1945 with the rest of the von Braun team, many members of the group in Germany had become interested in space through Oberth’s books. Also in response to Oberth’s first book, in 1927 the German Society for Space Travel (Verein fur Raumschiffart) was founded to raise money for him to perform rocket experiments. He served as presi­dent in 1929-30, and the organization provided considerable practi­cal experience in rocketry to several of its members (including von Braun). Some of them later served under von Braun at Peenemunde, although they constituted a very small fraction of the huge staff there (some 6,000 by mid-1943).13

Both Goddard and Oberth exemplified the pronouncement of Goddard at his high school graduation speech “that the dream of yesterday is the hope of today and the reality of tomorrow."14 But ironically it appears to have been Oberth who made the more im­portant contribution to the realization of both men’s dreams.15 In any event, both men made extraordinary, pioneering contributions that were different but complementary.

While NASA was just getting started with the massive development effort for the Saturn launch vehicles, the air force began work on what became the Titan family of launch vehicles, beginning with the Titan IIIs and ending with Titan IVBs. Essentially, most of these vehicles consisted of upgraded Titan II cores with a series of upper

stages plus a pair of huge segmented, solid-propellant, strap-on mo­tors to supplement the thrust of the Titan II core vehicle. And after September 1988, a limited number of actual Titan IIs, refurbished and equipped with technology and hardware from the Titan III program, joined the other members of the Titan family of launch vehicles. Be­ginning in June 1989, the Titan IV with a stretched core and seven (instead of Titan III’s five or five and a half) segments in its solid – rocket motors became the newest member of the Titan family.93

By September 1961, the DoD had agreed to the concept of com­bining a suitably modified Titan II with strap-on solid motors to sat­isfy military requirements; and the following month, a DoD-NASA Large Launch Vehicle Planning Group recommended the Titan III, as the vehicle had come to be designated. It would feature 120-inch – diameter solid motors and would serve both DoD and NASA needs “in the payload range of 5,000 to 30,000 pounds, low-Earth orbit equivalent."94

Although the air force’s Space Systems Division, which oversaw development of the Titan III, was later to complain about “daily redirection" of the program from the office of the director of de­fense, research and engineering, initially the launch vehicle got off to a quick start. Titan II contractor Martin Marietta Company (so – named since October 10, 1961, as a result of Martin’s merger with the American Marietta Company) won a contract on February 19, 1962. A second contract, highly significant in its requirements for development of new technology, covered the large solid-propellant rocket motors to boost the Titan III. On May 9, 1962, the air force selected a new firm, named United Technology Corporation (UTC), to develop the solid-rocket motors.95

Not long after the founding of UTC in 1958 (under the name United Research Corporation), United Aircraft Corporation pur­chased a one-third interest in the rocket firm, later becoming its sole owner. When United Aircraft changed its name to United Technologies Corporation in 1975, its solid-propellant division be – 86 came Chemical Systems Division (CSD). Formerly a contributor Chapter 2 to Minuteman, UTC’s second president, Barnet R. Adelman, had been an early proponent of segmentation for large solid-rocket mo­tors to permit easier transportation. Other firms, including Aerojet, Lockheed, and Thiokol, participated in the early development of the technology, but UTC developed its own clevis joint design to connect the segments of such boosters and its own variant on the propellant used for Minuteman to provide the propulsion.96

Because there was a Titan IIIA that did not include the solid- rocket motors, some of the Titan III first-stage engines would fire

at ground level, whereas those used on the Titan IIIC would start at altitude after the solid-rocket motors lifted the vehicle to about 100,000 feet. Titan III also featured a new third stage known as Transtage.97 This featured a pressure-fed engine using the same pro­pellants as stages one and two. Aerojet won this contract in addi­tion to those for the first two stages, with a two-phase agreement signed in 1962 and 1963. Aerojet designed the Transtage engine to feature two ablatively cooled thrust chambers and a radiatively cooled nozzle assembly.98

The Transtage engine could start and stop in space, allowing it to place multiple satellites into different orbits on a single launch or to position a single satellite in a final orbit without a need for a sepa­rate kick motor. In August 1963, tests at the simulated-altitude test chamber of the air force’s Arnold Engineering Development Center (AEDC) in Tullahoma, Tennessee, confirmed earlier suspicions that the combustion chamber would burn through before completing a full-duration firing. How Aerojet solved this and other problems is not explained in the sources for this book, only that it required “ex­tensive redesign and testing."99 Obviously, Aerojet engineers had not anticipated these problems in their initial design. Clearly, this was another example of the roles of testing and problem solving in rocket development as well as the involvement of multiple organi­zations in the process.

In any event, engine deliveries did not occur in mid-December, as initially planned, but in April 1964. Additionally, Aerojet had to test the engine at sea level and extrapolate the data to conditions at altitude. When the data from the simulated-altitude tests at AEDC came back, the extrapolated data were 2.5 percent higher than the Arnold figures. This might seem a small discrepancy to the casual reader. But since the program needed exact performance data to project orbital injection accurately, Aerojet had to investigate the discrepancy. The explanation proved to be simple, but it illustrates the difficulty of pulling together all relevant data for development of something as complex as a rocket engine, even within the same firm. It also meant that engineers did not have their procedures “down to a science" but sometimes operated with an incomplete understand­ing of the phenomena they were testing in programs where fund­ing and schedules precluded thorough and meticulous research. It turned out that other engineers working on a solid rocket had al­ready learned to decrease the calculations by 2.5 percent to extrapo­late for conditions at altitude. Once aware of this, Transtage engi­neers found several references to this correction in the literature. But obviously, they initially had failed to find those references.100

There were several problems with the Titan IIIC, resulting in 4 failures in 18 launches from September 1, 1964, to April 8, 1970.101 In ensuing years, there were many versions of the Titan III. Besides the Titan IIIA, there was a Titan 23C with uprated thrust for the core liq­uid stages and a simplified and lightened thrust-vector-control sys­tem for the solid-rocket motors. The 23C flew 22 times by March 6, 1982, with 19 successful missions and 3 failures. Overall, between the original Titan IIIC and the Titan 23C versions, Titan IIIC had 33 successful launches and 7 failures for a success rate of 82.5 per­cent. Four of the 7 failures were due to Transtage problems, without which the overall vehicle would have had a much more successful career.102

Another version of the Titan III was the Titan IIIB with an Agena D replacing the Transtage in the core stack of three stages. The Ti­tan IIIB did not use solid-rocket boosters. With the Agena D’s 5,800 pounds of thrust compared with Transtage’s roughly 1,600, the Ti­tan IIIB could launch a 7,920-pound payload to a 115-mile Earth orbit compared with 7,260 pounds for the Titan IIIA. At some point, certainly by 1976, a stretched version of the first stage converted the vehicle to a 24B configuration. And in 1971 a Titan 33B ver­sion first operated, featuring an “Ascent Agena"—so-called because it became purely a launch stage instead of staying attached to the payload to provide power and attitude control while it was in orbit. Between June 29, 1966, and February 12, 1987, various versions of the Titan IIIB (including 23B and 34B) with Agena D third stages launched some 68 times with only 4 known failures—a 94 percent success rate.103

On November 15, 1967, the Titan III Systems Program Office be­gan designing, developing, and ultimately producing the Titan IIID, which essentially added Titan IIIC’s solid-rocket motors to the Ti­tan IIIB. Perhaps more accurately, it can be considered a Titan IIIC without the Transtage. By this time, Air Force Systems Command had inactivated Ballistic and Space Systems Divisions (BSD and 88 SSD) and reunited the two organizations into the Space and Mis – Chapter 2 sile Systems Organization (SAMSO), headquartered in the former SSD location at Los Angeles Air Force Station. The D models car­ried many photo-reconnaissance payloads that were too heavy for the B models. The Titan IIID could carry a reported 24,200 pounds of payload to a 115-mile orbit, compared with only 7,920 for the B model.104 The D model appears to have launched 22 heavy-imaging satellites from June 15, 1971, to November 17, 1982. All 22 launches seem to have been successful, giving the Titan IIID a perfect launch

record.105

On June 26, 1967, NASA contracted with Martin Marietta to study the possibility of using a Titan-Centaur combination for mis­sions such as those to Mars and the outer planets in the solar sys­tem. When this possibility began to look promising, in March 1969, NASA Headquarters assigned management of the vehicle to the Lewis Research Center, with follow-on contracts going to Martin Marietta (via the air force) and General Dynamics/Convair (directly) to study and then develop what became the Titan IIIE and to adapt the Centaur D-1 for use therewith.106 The Titan IIIE and Centaur D-1T were ready for a proof flight on February 11, 1974. Unfortu­nately, the upper stage failed to start. But from December 10, 1974, to September 5, 1977, Titan IIIE-Centaurs launched two Helios so­lar probes, two Viking missions to Mars, and two Voyager missions to Jupiter and Saturn, all successful.107

By the mid – to late 1970s, air force planners perceived a need for still another Titan configuration to carry increasingly large payloads such as the Defense Satellite Communication System III (DSCS III) satellites into orbit before the Space Shuttle was ready to assume such responsibilities. (The first DSCS III weighed 1,795 pounds, a significant jump from the DSCS II weight of 1,150 pounds.) Even after the shuttle became fully operational, the Titan 34D, as the new vehicle came to be called, would continue in a backup role in case the shuttle was unavailable for any reason. The air force con­tracted with Martin Marietta in July 1977 for preliminary design, with a production contract for five Titan 34D airframes following in January 1978. SAMSO retained program responsibility for the Titan family of vehicles, and it contracted separately with suppli­ers of the component elements. It appears that the long-tank first stage was the driving element in the new vehicle. This seems to be the premise of a 1978 article in Aviation Week & Space Technol­ogy stating that CSD’s solid-rocket motors (SRMs) would add half a segment “to make them compatible with the long-tank first stage." Thus, the SRMs contained five and a half segments in place of the five used on previous Titans.108

Equipped with these longer solid-rocket motors and an uprated Transtage, the Titan 34D could carry 32,824 pounds to a 115-mile orbit, as compared with 28,600 pounds for the Titan IIIC. The 34D could lift 4,081 pounds to geosynchronous orbit, which compared favorably with 3,080 pounds for the IIIC but not with the 7,480 pounds the Titan IIIE-Centaur could carry to the same orbit.109

A quite different but important upper stage had its maiden launch on the first Titan 34D and later launched on several Titan IVs. This was the Inertial Upper Stage (IUS) that sat atop stage two on the

first Titan 34D launch. Unlike the rest of the booster, this stage was anything but easy to develop. In August 1976, the air force selected Boeing Aerospace Company as the prime IUS contractor. Soon after­ward, Boeing subcontracted with CSD to design and test the solid mo­tors to be used in the IUS. CSD chose to use a hydroxy-terminated poly­butadiene propellant (also being used by Thiokol in the Antares IIIA motor for Scout, developed between 1977 and 1979). Problems with the propellant, case, and nozzles delayed development of IUS. Vari­ous technical and managerial problems led to more than two years of delay in the schedule and cost overruns that basically doubled the originally projected cost of the IUS. These problems showed that despite more than two and a half decades of rocket development, rocket engineering still often required constant attention to small details and, where new technology was involved, a certain amount of trial and error. Including its first (and only) IUS mission, the Titan 34D had a total of 15 launches from both the Eastern and Western Test Ranges between October 1982 and September 4, 1989. There were 3 failures for an 80 percent success rate.110

By the mid-1980s, the air force had become increasingly uncom­fortable with its dependence on the Space Shuttle for delivery of military satellites to orbit. Eventually, this discomfort would lead to the procurement of a variety of Titan IV, Delta II, and Atlas II ex­pendable launch vehicles, but the air service also had at its disposal 56 deactivated Titan II missiles in storage at Norton AFB. Conse­quently, in January 1986 Space Division contracted with Martin Marietta to refurbish a number of the Titan IIs for use as launch vehicles. Designated as Space Launch Vehicle 23G, the Titan II had only two launches during the period covered by this book, on Sep­tember 5, 1988, and the same date in 1989, both carrying classified payloads from Vandenberg AFB. For a polar orbit from Vandenberg, the Titan II could carry only about 4,190 pounds into a 115-mile orbit, but this compared favorably with the Atlas E. Although the Atlas vehicle could launch about 4,600 pounds into the same orbit, 90 it required two Thiokol TE-M-364-4 solid-rocket motors in addition Chapter 2 to its own thrust to do so.111

The Titan IV grew out of the same concern about the availability of the Space Shuttle that had led to the conversion of Titan II mis­siles to space-launch vehicles. In 1984 the air force decided that it needed to ensure access to space in case no Space Shuttle was available when a critical DoD payload needed to be launched. Con­sequently, Space Division requested bids for a contract to develop a new vehicle. Martin Marietta proposed a modified Titan 34D and won a development contract on February 28, 1985, for 10 of the

vehicles that became Titan IVs. Following the Challenger disaster, the air force amended the contract to add 13 more vehicles.112

The initial version of the new booster (later called Titan IVA) had twin, 7-segment solid-rocket motors produced by CSD as a subcon­tractor to Martin Marietta. These contained substantially the same propellant and grain configuration as the Titan 34D but with an ad­ditional 1.5 segments, bringing the length to about 122 feet and the motor thrust to 1.394 million pounds per motor at the peak (vacuum) performance. The Aerojet stages one and two retained the same con­figurations as for the Titan 34D except that stage one was stretched about 7.9 feet to allow for more propellant and thus longer burning times. Stage two, similarly, added 1.4 feet of propellant tankage.113

The first launch of a Titan IV took place at Cape Canaveral on June 14, 1989, using an IUS as the upper stage. There were four more Titan IV launches during the period covered by this book, but the vehicle went on to place many more satellites into orbit into the first years of the 21st century.114 Including the 14 Titan II missiles reconfigured into launch vehicles after the missiles them­selves were retired, 12 of which had been launched by early 2003, there had been 214 Titan space-launch vehicles used by that point in time. Of them, 195 had succeeded in their missions and 19 had failed, for a 91.1 percent success rate.115 This is hardly a brilliant record, but with such a variety of types and a huge number of com­ponents that could (and sometimes did) fail, it is a creditable one. It shows a large number of missions that needed the capabilities of the Titan family members for their launch requirements.

However, if the handwriting was not yet quite on the wall by 1991, it had become clear by 1995 that even in its Titan IVB con­figuration, the Titan family of launch vehicles was simply too ex­pensive to continue very far into the 21st century as a viable launch vehicle. Based on studies from the late 1980s and early 1990s, the air force had come up with what it called the Evolved Expendable Launch Vehicle (EELV) program to replace the then-existing Delta II, Atlas II, Titan II, and Titan IV programs with a family of boosters that would cost 25 to 50 percent less than their predecessors but could launch 2,500 to 45,000 pounds into low-Earth orbit with a 98 percent reliability rate, well above that achieved historically by the Titan family.116

The initial decision to use liquid-hydrogen technology in the upper stages of the Saturn launch vehicles came from a Saturn Vehicle Team, chaired by Abe Silverstein and including other representa­tives from NASA Headquarters, the air force, the Office of Defense Research and Engineering, and the Army Ballistic Missile Agency 190 (von Braun, himself). Meeting in December 1959, this group, in­Chapter 5 fluenced by Silverstein’s convictions about the performance capa­bilities of liquid hydrogen, agreed to employ it in the Saturn upper stages. Silverstein managed to convince even von Braun, despite reservations, to take this step. But von Braun later told William Mrazek he was not greatly concerned about the difficulties of the new fuel because many Centaur launches were scheduled before the first Saturn launch with upper stages. His group could profit from what these launches revealed to solve any problems with the Saturn I upper stages.44

THE S-IV STAGE

On April 26, 1960, NASA awarded a contract to the Douglas Aircraft Company to develop the Saturn I second stage, the S-IV. Between January and March 1961, NASA decided to use Pratt & Whitney RL10 engines in this stage. But instead of the two RL10s in Centaur, the S-IV held six such engines. Benefiting from consultations NASA arranged with Convair and Pratt & Whitney, Douglas did use a tank design similar to Convair’s, with a common bulkhead between the liquid oxygen and the liquid hydrogen. But Douglas also relied on its own experience in its use of materials and methods of manufac­ture. So the honeycomb material in the common bulkhead of the propellant tank was different from Convair’s design, drawing upon Douglas’s work with panels in aircraft wings and some earlier mis­sile designs. Douglas succeeded in making the larger tanks and S-IV

stage in time for the first launch (SA-5) of a Saturn I featuring a live second stage on January 29, 1964.45

Remarkably, this launch was successful despite a major accident only five days earlier. Douglas engineers and technicians knew that they had to take special precautions with liquid oxygen and liquid hydrogen. The latter was especially insidious because if it leaked and caught fire in the daylight, the flames were virtually invisible. Infrared TV cameras did not totally solve the problem because of the difficulty of positioning enough of them to cover every cranny where hydrogen gas might hide. So crews with protective clothes carried brooms in front of them. If a broom caught fire, hydrogen was leaking and burning.

Despite such precautions, on January 24, 1964, at a countdown to a static test of the S-IV, the stage exploded. Fortunately, the re­sultant hydrogen fire was short-lived, and a NASA committee with Douglas Aircraft membership determined that the cause was a rup­ture of a liquid-oxygen tank resulting from the failure of two vent valves to relieve pressure that built up. The relief valves were in­capacitated by solid oxygen, which had frozen because helium gas to pressurize the oxygen tank had come from a sphere submerged in the liquid hydrogen portion of the tank. This helium was colder than the freezing point of oxygen. The pressure got so high because the primary shutoff valve for the helium failed to close when nor­mal operating pressure had developed in the oxygen tank. Testing of the shutoff valve showed that it did not work satisfactorily in cold conditions. Because this valve had previously malfunctioned, it should have been replaced by this time. In any event, Saturn proj­ect personnel did apparently change it to another design before the launch five days later. The committee “found that no single person, judgment, malfunction or event could be directly blamed for this incident," but if “test operations personnel had the proper sensitiv­ity to the situation the operation could have been safely secured" before the accident got out of hand.46

On the six test flights with the S-IV stage (SA-5 through SA-10, the last occurring July 30, 1965), it and the already tested RL10 engines worked satisfactorily. They provided 90,000 pounds of thrust and demonstrated, among other things, that liquid-hydrogen technol­ogy had matured significantly, at least when using RL10 engines.47

THE S-IVB STAGE

For the intermediate version of the Saturn launch vehicle, the Sat­urn IB, engineers for the S-IVB second stage further added to the payload capacity of the overall vehicle through reducing the weight

of the stage by some 19,800 pounds. Part of the reduction came from redesigned and smaller aerodynamic fins. Flight experience with the Saturn I also revealed that the initial design of the stage had been excessively conservative, and engineers were able to trim propellant tanks, a “spider [structural] beam," and other compo­nents as well as to remove “various tubes and brackets no longer required." But production techniques and most tooling did not change significantly.48

The S-IVB featured a totally new and much larger engine, the J-2, with more thrust than the six RL10s used on the Saturn I. This was the liquid-hydrogen/liquid-oxygen engine the Silverstein commit­tee had recommended for the Saturn upper stages on December 15,

1959, following which NASA requested proposals from industry to design and build it. There were five companies competing for the contract, with the three top candidates being North Ameri­can Aviation’s Rocketdyne Division, Aerojet, and Pratt & Whit­ney. Having built the RL10, Pratt & Whitney might seem to have been the logical choice, but even though NASA’s source evaluation

192 board had judged all three firms as capable of providing a satisfac- Chapter 5 tory engine, Pratt & Whitney’s proposal cost more than twice those of Aerojet and Rocketdyne. Rocketdyne’s bid was lower than Aero­jet’s, based on an assumption of less testing time, but even if the testing times were equalized, it appeared that Rocketdyne’s cost was still lower. Thus, on May 31, 1960, Glennan decided to negoti­ate with Rocketdyne for a contract to design and build the engine. The von Braun group and Rocketdyne then worked together on the design of the engine. A final contract signed on September 10,

1960, stated that the engine would ensure “maximum safety for manned flight" while using a conservative design to speed up development.49

Rocketdyne began the development of the J-2 on September 1, 1960, with a computer simulation to assist with the configuration. Most of the work took place at the division’s main facility at Ca – noga Park in northwestern Los Angeles, with firing and other tests at the Santa Susana Field Laboratory in the nearby mountains. By early November, the Rocketdyne engineers had designed a full – scale injector and by November 11 had conducted static tests of it in an experimental engine. Rocketdyne also built a large vacuum chamber to simulate engine firings in space. By the end of 1961, it was evident that the J-2 would provide power for not only the sec­ond stage of Saturn IB but the second and third stages of the Saturn V (then known as the Saturn C-5). In the second stage of Saturn V, there would be a cluster of five J-2s; on the S-IVB second stage of

Saturn IB and the S-IVB third stage of Saturn V, there would be a single J-2.50

Rocketdyne’s engineers borrowed technology from Pratt & Whitney’s RL10, but since the J-2 (with its initial design goal of 200,000 pounds of thrust at altitude) was so much larger than the 15,000-pound RL10, designers first tried flat-faced copper injectors similar to designs Rocketdyne was used to in its liquid-oxygen/ RP-1 engines. Heating patterns for liquid hydrogen turned out to be quite different from those for RP-1, and injectors got so hot the copper burned out. The RL10 had used a porous, concave injector of a mesh design, cooled by a flow of gaseous hydrogen, but Rock – etdyne would not adopt that approach until 1962, when Marshall engineers insisted designers visit Lewis Research Center to look at examples. Under pressure, the California engineers adopted the RL10 injector design, and problems with burnout ceased. In this instance, a contractor benefited from an established design from another firm, even if only under pressure from the customer, il­lustrating the sometimes difficult process of technology transfer. Thus, Rocketdyne avoided further need for injector design, which, in NASA’s assistant director for propulsion A. O. Tischler’s words, was still “more a black art than a science."51

Rocketdyne expertise seems to have been more effective in de­signing the combustion chamber, consisting of intricately fashioned stainless-steel cooling tubes with a chamber jacket made of Inco­nel, a nickel-chromium alloy capable of withstanding high levels of heating. Using a computer to solve a variety of equations having to do with energy, momentum, heat balance, and other factors, de­signers used liquid hydrogen to absorb the heat from combustion before it entered the injector, “heating" the fuel in the process from —423°F to a gaseous temperature of -260°F. The speed of passage through the cooling tubes varied, with adjustments to match com­puter calculations of the needs of different locations for cooling.52

Because of the low density of hydrogen and the consequent need for a higher-volume flow rate for it vis-a-vis the liquid oxygen (al­though by weight, the oxygen flowed more quickly), Rocketdyne decided to use two different types of turbopumps, each mounted on opposite sides of the thrust chamber. For the liquid oxygen, the firm used a conventional centrifugal pump of the type used for both fuel and oxidizer in the RL10. This featured a blade that forced the propellant in a direction perpendicular to the shaft of the pump. It operated at a speed of 7,902 revolutions per minute and achieved a flow rate of 2,969 gallons per minute. For the liquid hydrogen, an axial-type pump used blades operating like airplane propellers to

force the propellant in the direction of the pump’s shaft. Operating in seven stages (to one for the liquid-oxygen pump), the fuel pump ran at 26,032 revolutions per minute and sent 8,070 gallons of liq­uid hydrogen per minute to the combustion chamber. (By contrast, in terms of weight, 468 pounds of liquid oxygen to 79 pounds of liquid hydrogen per second flowed from the pumps.) A gas genera­tor provided fuel-rich gas to drive the separate turbines for the two pumps, with the flow first to the hydrogen and then to the oxygen pump. The turbine exhaust gas flowed into the main rocket nozzle for disposal and a slight addition to thrust.53

In testing the J-2, engineers experienced problems with such is­sues as insulation of the cryogenic liquid hydrogen, sealing it to avoid leaks that could produce explosions, and a phenomenon known as hydrogen embrittlement in which the hydrogen in gas­eous form caused metals to become brittle and break. To prevent this, technicians had to coat high-strength super alloys with copper or gold. Solving problems that occurred in testing often involved trial-and-error methods. Engineers and technicians never knew, 194 until after further testing, whether a given “fix" actually solved Chapter 5 a problem (or instead created a new one). Even exhaustive testing did not always discover potential problems before flights, but engi­neers always hoped to find problems in ground testing rather than flight.54

Rocketdyne completed the preliminary design for the 200,000- pound-thrust J-2 in April 1961, with the preflight readiness testing finished in 1964 and engine qualification, in 1965. The engine was gimballed for steering, and it had a restart capability, using helium stored in a separate tank within the liquid-hydrogen tank to oper­ate the pneumatic system. Soon after the 200,000-pound J-2 was qualified, Rocketdyne uprated the engine successively to 205,000, 225,000, and then 230,000 pounds of thrust at altitude. Engineers did this partly by increasing the chamber pressure. They also ad­justed the ratio of oxidizer to fuel. The 200,000-pound-thrust engine used a mixture ratio of 5:1, but the more powerful versions could adjust the mixture ratio in flight up to 5.5:1 for maximum thrust and as low as 4.5:1 for a lower thrust level. During the last portion of a flight, the valve position shifted to ensure the simultaneous emptying of the liquid oxygen and the liquid hydrogen from the propellant tanks (technically, a single tank with a common bulk­head, but referred to in the plural as if there were separate tanks). The 225,000-pound-thrust engine had replaced the 200,000-pound version on the production line by October 1966, with the 230,000- pound engine available by about September 1967. As the uprated

versions became available, Rocketdyne gradually ceased producing the lower-rated ones.55

Even with six RL10s, the S-IV stage had been only about 39.7 feet tall by 18.5 feet in diameter. To contain the single J-2 and its propellant tank, the S-IVB had to be 58.4 feet tall by 21.7 feet in di­ameter. NASA selected Douglas to modify its S-IV to accommodate the J-2 on December 21, 1961. Douglas had already designed the S-IV to have a different structure from that of the Centaur, with the latter’s steel-balloon design (to provide structural support) be­ing replaced by a self-supporting structure more in keeping with the “man-rating" that had initially been planned for Saturn I and transferred to Saturn IB, which actually would launch astronauts into orbit. This structure was made of aluminum and consisted of “skin-and-stringer" type construction.

The propellant tank borrowed a wafflelike structure with ribs from the Thor tanks Douglas had designed. The common bulkhead between the liquid hydrogen and the liquid oxygen required only minor changes from the smaller one in the S-IV. After conferring with Convair about the external insulation used to keep the liquid hydrogen from boiling away rapidly in the Centaur, Douglas en­gineers had decided on internal insulation for the fuel tank in the S-IV. They chose woven fiberglass threads cured with polyurethane foam to form a tile that technicians shaped and installed inside the tank. This became the insulation for the S-IVB as well.56 Thus, in this case technology did not transfer between firms, but shared in­formation helped with a technical decision.

For steering the S-IVB during the firing of the J-2, Douglas had initially designed a slender actuator unit to gimbal the engine, simi­lar to devices on the firm’s aircraft landing gear. Marshall engineers said the mission required stubbier actuators. This proved to be true, leading Douglas to subcontract the work to Moog Servo Controls, Inc., of Aurora, New York, which used Marshall specifications to build the actuators. The gimballed engine could adjust the stage’s direction in pitch and yaw. For roll control during the firing of the J-2, and for attitude control in all three axes during orbital coast, an auxiliary propulsion system provided the necessary thrust.57

Although they had the same designation, the S-IVB used on the Saturn V was heavier and different in several respects from the one on the Saturn IB. As the third stage on the Saturn V, the S-IVB profited greatly from the development and testing for the Saturn IB second stage. But unlike the latter, it required an aft interstage that flared out to the greater diameter of the Saturn V plus control mechanisms to restart the engine in orbit for the burn that would

send the Apollo spacecraft on its trajectory to lunar orbit. To match with the greater girth of the S-II, the aft skirt for the third stage was heavier than the one for the S-IVB second stage. The forward skirt was heavier as well to permit a heavier payload. The auxiliary propulsion and ullage system weighed more for the third stage of the Saturn V than the comparable second stage on the IB because of increased attitude control and venting needed for the lunar mis­sions. Finally, the propulsion system was heavier for the Saturn V third stage because of the need to restart. The total additions came to some 11,000 pounds of dry weight. Whereas the first burn of the single J-2 engine would last only about 2.75 minutes to get the third stage and payload to orbital speed at about 17,500 miles per hour, the second burn would last about 5.2 minutes and would accelerate the stage and spacecraft to 24,500 miles per hour, the typical escape velocity for a lunar mission.58

On the aft skirt assembly, mounted 180 degrees apart, were two auxiliary propulsion modules. Each contained three 150-pound – thrust attitude-control engines and one 70-pound-thrust ullage – 196 control engine. Built by TRW, the attitude-control engines burned Chapter 5 a hypergolic combination of nitrogen tetroxide and monomethyl hydrazine. They used ablative cooling and provided roll control dur­ing J-2 firing and control in pitch, yaw, and roll during coast periods. The ullage-control engines, similar to those for attitude control, fired before the coast phase to ensure propellants concentrated near the aft end of their tanks. They fired again before engine restart to position propellants next to feed lines. There were also two ullage – control motors 180 degrees apart between the auxiliary propulsion modules. These motors fired after separation from the S-II stage to ensure that the propellants in the engine’s tanks were forced to the rear of the tanks before ignition of the third-stage J-2. The two motors were Thiokol TX-280s burning solid propellants to deliver about 3,390 pounds of thrust.59

Despite the relatively modest changes in the S-IVB for Saturn V, development was not problem-free. In acceptance testing of the third stage at Douglas’s Sacramento test area on January 20, 1967, the entire stage exploded. Investigation finally revealed that a he­lium storage sphere had been welded with pure titanium rather than an alloy. When it exploded, it cut propellant lines and allowed the propellants to mix, ignite, and explode, destroying the stage and adjacent structures. The human error led to revised welding specifications and procedures. Despite the late date of this mishap, the S-IVB was ready for the first Saturn V mission on November 9,

1967, when it performed its demanding mission, including restart, without notable problems.60

THE S-II STAGE

The S-II second stage for the Saturn V proved to be far more problem­atic than the S-IVB third stage. On September 11, 1961, NASA had selected North American Aviation to build the S-II. The division of North American that won the S-II contract was the Space and Infor­mation Systems Division (previously the Missile Division), headed by Harrison A. Storms Jr., who had managed the X-15 project. An able, articulate engineer, Storms was charismatic but mercurial. His nickname, “Stormy," reflected his personality as well as his last name. (People said that “while other men fiddle, Harrison storms.") His subordinates proudly assumed the title of Storm Troopers, but he could be abrasive, embodying what X-15 test pilot and engineer Scott Crossfield called “the wire brush school of management."61

When Storms’s division began bidding on the S-II contract, the configuration of the stage was in flux. Early in 1961 when NASA administrator James Webb authorized Marshall to initiate contrac­tor selection, 30 aerospace firms attended a preproposal conference. There, NASA announced that the stage would contain only four J-2 engines (instead of the later five), and it would be only about 74 feet tall (compared with the later figure of 81 feet, 7 inches for the actual S-II). The projected width was 21 feet, 6 inches (rather than the later 33 feet). It still seemed imposingly large, but it was “the precision it would require [that] gave everybody the jitters—like building a locomotive to the tolerance of a Swiss watch," as Storms’s biog­rapher put it. This sort of concern whittled the number of inter­ested firms down to seven. A source evaluation board eliminated three, leaving Aerojet, Convair, Douglas, and North American to learn that they were now bidding on a stage enlarged to at least a diameter of 26 feet, 9 inches—still well short of the final diameter. Also still missing was precise information about configuration of the stages above the S-II. The Marshall procurement officer did em­phasize that an important ingredient in NASA’s selection would be “efficient management."62

Once Storms’s division won the contract for the stage, it did not take long for NASA to arrive at the decision, announced Janu­ary 10, 1962, that the S-II would hold five J-2 engines. Designers decided to go with a single tank for the liquid hydrogen and liquid oxygen with a common bulkhead between them, like the design for Douglas’s much smaller common tank for the S-IVB. (The S-II

contained 260,000 gallons of liquid hydrogen and 83,000 gallons of liquid oxygen to 63,000 and 20,000 gallons, respectively, in the S-IVB.) As with the Douglas stage, common parlance referred to each segment as if it were a separate tank. Obviously, the common bulkhead was much larger in the second than the third stage (with a diameter of 33 rather than 21.75 feet), requiring unusual precision in the welding to preclude leakage. The bulkhead consisted of the top of the liquid-oxygen tank, a sheet of honeycombed phenolic insula­tion bonded to the metal beneath it, and the bottom of the liquid – hydrogen tank. Careful fitting, verified by ultrasonography, ensured complete bonding and the absence of gaps. Not only did fit have to be perfect but there were complex curvatures and a change in thick­ness from a maximum of about 5 inches in the center to somewhat less at the periphery.63

Unlike Douglas but like Convair (in the Centaur), North Ameri­can decided to use external insulation, which (it argued) increased the strength of the tank because of the extreme cold inside the tank, which was imparted to the tank walls. Initially, Storms’s en – 198 gineers tried insulation panels, but the bonding failed repeatedly Chapter 5 during testing. Using trial-and-error engineering, designers turned to spraying insulation directly onto the tank, allowing it to cure, and then adjusting it to the proper dimensions. Once the tanks were formed and cleaned, North American installed slosh baffles inside the tanks.64

The reason that insulation on the outside of the liquid-hydrogen tank increased its strength was the use of an aluminum alloy desig­nated 2014 T6 as the material for the S-II tanks. Employed long be­fore on the Ford Trimotor, it had the unusual characteristic of get­ting stronger as it got colder. At -400°F, it was 50 percent stronger than at room temperature. With the insulation on the outside, this material provided a real advantage with the -423°F liquid hydrogen inside. Both the oxidizer and fuel tank walls could be 30 percent thinner than with another material.

Unfortunately, aluminum 2014 T6 was difficult to weld with al­most 104 feet of circumference. On the first try at attaching two cylinders to one another, welders got about four-fifths around the circle when the remaining portion of the metal “ballooned out of shape from the heat buildup." The Storm Troopers had to resort to powerful automated welding equipment to do the job. Each ring to be welded had to be held in place by a huge precision jig with about 15,000 adjustment screws around the circumference, each less than an inch from the next. A mammoth turntable rotated the seam through fixed weld heads with microscopic precision. A huge

clean room allowed the humidity to be kept at 30 percent. In all of this, Marshall’s experience with welding, including that for the S-IC stage, helped Storms’s people solve their problems.65

Despite such help, there was considerable friction between Storms’s division, on the one hand, and Marshall on the other, espe­cially with Eberhard Rees, von Braun’s deputy director for technical matters. North American fell behind schedule and had increasing technical and other problems. Marshall officials began to complain about management problems with the contractor, including a fail­ure to integrate engineering, budgeting, manufacturing, testing, and quality control. At the same time, Storms’s division was the victim of its own delays on the Apollo spacecraft it was also building. The weight of Apollo payloads kept increasing. This required lightening the launch-vehicle stages to compensate. The logical place to do so was the S-IVB stage, because a pound reduced there had the same effect as 4 or 5 pounds taken off the S-II (or 14 pounds from the S-IB). This resulted from the lower stages having to lift the upper ones plus themselves. But the S-IVB, used on the Saturn IB, was already in production, so designers had to make reductions in the thickness and strength of the structural members in the S-II.66

By mid-1964, the S-II insulation was still a problem. Then in Oc­tober 1964, burst tests showed that weld strength was lower than expected. On October 28, a rupture of the aft bulkhead for an S-II occurred during hydrostatic testing. As the date for launch of the first Saturn V (1967) approached, von Braun proposed eliminating a test vehicle to get the program back on schedule. Sam Phillips agreed. Instead of a dynamic as well as a structural test vehicle, the structural stage would do double duty.

But on September 29, 1965, the combined structural and dynamic test vehicle underwent hydraulic testing at Seal Beach. While the tanks filled with water, the vehicle was simultaneously subjected to vibration, twisting, and bending to simulate flight loads. Even though the thinned structure was substantially less strong than it would have been at the colder temperatures that would have pre­vailed with liquid hydrogen in the tanks, Marshall had insisted on testing to 1.5 times the expected flight loads. At what was subse­quently determined to be 1.44 times the load limit, the welds failed and the stage broke apart with a thunderous roar as 50 tons of water cascaded through the test site. The program was short another test vehicle. Storms’s people looked at the effect on the cost of the pro­gram and concluded that to complete the program after the failure would raise the cost of the contract from the initial $581 million to roughly $1 billion.67

When Lee Atwood, president of North American, flew to Hunts­ville on October 14, Brig. Gen. Edmund O’Connor of the air force, director of Marshall’s Industrial Operations, told von Braun, “The S-II program is out of control. . . . [Management of the project at both the program level and the division level. . . has not been ef­fective." Von Braun told Atwood the S-II needed a more forceful manager than William F. Parker, quiet but technically knowledge­able, whom Storms had appointed to head the program in 1961. Von Braun apparently got Atwood’s agreement to replace Parker and put a senior manager in charge of monitoring the program.68

The day after Atwood’s visit to Huntsville, Rees flew to Hous­ton, where he met with other Apollo managers, including Phillips. The Manned Spacecraft Center was managing Storms’s programs for the Apollo spacecraft, and Houston manager Joseph Shea had complaints similar to those of Rees about Storms’s control of costs and schedules. Phillips decided to head an ad hoc fact-finding (“ti­ger") team with people from Marshall and Houston to visit North American and investigate.69

200 The team descended upon North American on November 22, Chapter 5 and on December 19, 1965, Phillips presented the findings. George Mueller had already expressed concerns to Lee Atwood about the S-II and spacecraft programs at Storms’s Space and Information Sys­tems Division. In a letter to Atwood dated December 19 he reiter­ated, “Phillips’ report has not only corroborated my concern, but has convinced me beyond doubt that the situation at S&ID requires positive and substantive actions immediately in order to meet the national objectives of the Apollo Program." After pointing to nu­merous delays and cost overruns on both the S-II and the spacecraft, Mueller wrote, “It is hard for me to understand how a company with the background and demonstrated competence of NAA could have spent 4 1/2 years and more than half a billion dollars on the S-II project and not yet have fired a stage with flight systems in op­eration." He said Sam Phillips was convinced the division could do a better job with fewer people and suggested transferring to another division groups like Information Systems that did not contribute directly to the spacecraft and S-II projects.70

A memorandum from Phillips to Mueller the day before had been even more scathing: “My people and I have completely lost confidence in NAA’s competence as an organization to do the job we have given them." He made specific recommendations for man­agement changes, including “that Harrison Storms be removed as President of S&ID. . . . [H]is leadership has failed to produce re­sults which could have and should have been produced." After as-

suring Phillips and Mueller he would do what he could to correct problems, Atwood visited Downey and was reportedly impressed by the design work. He did not replace Storms, but Stormy him­self had already placed retired air force Maj. Gen. Robert E. Greer in a position to oversee the S-II. In January 1966, Greer added the titles of vice president and program manager for the program, keep­ing Bill Parker as his deputy. Greer agreed in a later interview that there were serious problems with S-II management. He revamped the management control center to ensure more oversight and in­corporated additional meetings the Storm Troopers called “Black Saturdays," implicitly comparing them with Schriever’s meetings at the Western Development Division. However, Greer, who had served at the (renamed) Ballistic Missile Division, held them daily at first, then several times a week, not monthly. With Greer’s sys­tems management and Parker’s knowledge of the S-II, there seemed to be hope for success.71

But setbacks continued. On May 28, 1966, in a pressure test at the Mississippi Test Facility, another S-II stage exploded. Human error was to blame for a failure to reconnect pressure-relief switches af­ter previous tests, but inspection revealed tiny cracks in the liquid – hydrogen cylinders that also turned up on other cylinders already fabricated or in production. Modification and repair occasioned more delays. But it took the Apollo fire in the command module during January 1967 and extreme pressure from Webb to cause At­wood to separate Information Systems from the Space Division (as it became), to move Storms to a staff position, and to appoint recent president of Martin Marietta William B. Bergen as head of Space Division, actually a demotion for which he volunteered from a posi­tion in which he had been Storms’s boss. Bergen’s appointment may have been more important for the redesign of the command module than for the S-II, and certainly Storms and North American were not solely to blame for the problems with either the stage or the spacecraft. But by late 1967, engineers had largely solved problems with both or had them on the way to solution.72

Because the German V-2 missile’s technology became available to U. S. missile and rocket programs after the end of World War II, it helped stimulate further development of American rocket technol­ogy. The V-2 was by no means the only contributor to that technol­ogy. More or less purely American rocket efforts also occurred be­tween the beginnings of the rocket development work by Germans working under von Braun and 1945 when some of those Germans and V-2s began to arrive in the United States. But in view of the im­portance of the V-2 to the development of American missiles and 14 launch vehicles after World War II, this section considers the work of Chapter 1 the Germans. A later section will trace the separate American efforts leading to U. S. ballistic missiles and, ultimately, launch vehicles.

Research leading to the V-2 began in 1932 when von Braun started working under Dornberger at the German army proving grounds in

Kummersdorf. The young man and his assistants experienced nu­merous failures, including burnthroughs of combustion chambers. They proceeded through test rockets labeled A-1, A-2, A-3, and A-5—the A standing for Aggregat (German for “assembly"). But as the size of their rockets (and the workforce) increased, they moved their operations to a much larger facility at Peenemunde on the German Baltic coast. There, they could launch their test rockets eastward along the Pomeranian coast.16

All of the test rockets contributed in various ways to the A-4, as did considerable collaboration with German universities, technical institutes, and industrial firms, showing that, as later in the United States, multiple organizations and skills were needed to develop missiles and rockets. Despite a truly massive amount of research – and-development work both at Peenemunde and at such associated entities, the A-4 still required a lot of modifications after its initial launch on October 3, 1942, with many failed launches after that. Even when actually used in the German war effort, the V-2 was nei­ther accurate nor reliable. Nevertheless, at about 46 feet long, 5 feet 5 inches in diameter, an empty weight of 8,818 pounds, and a range of close to 200 miles, it was an impressive technological achieve­ment whose development contributed much data and experience to later American missile and rocket development.17

Von Braun himself was a key factor in the relative success of the V-2. Born in the east German town of Wirsitz (later, Wyrzysk, Po­land) to noble parents on March 23, 1912, Freiherr (Baron) Wernher Magnus Maximilian von Braun earned a prediploma (Vordiplom) in mechanical engineering at the Berlin-Charlottenburg Institute of Technology in 1932, followed by a Ph. D. in physics from the University of Berlin in 1934.18 Both his boss, Walter Dornberger, and von Braun played the role of heterogeneous engineers, meeting with key figures in the government and Nazi Party, from successive Armaments Ministers Fritz Todt and Albert Speer, on up to Adolf Hitler himself, to maintain support for the missile.19

Von Braun also excelled as a technical manager after overcom­ing some initial lapses attributable to his youth and inexperience. He played a key role in integrating the various systems for the V-2 so that they worked effectively together. He did this by fostering communication between different departments as well as within individual elements of the Peenemunde organization. He met indi­vidually with engineers and perceptively led meetings of technical personnel to resolve particular issues. According to Dieter Huzel, who held a variety of positions at Peenemunde in the last two years of the war, von Braun “knew most problems at first hand. . . . He

repeatedly demonstrated his ability to go coherently and directly to the core of a problem or situation, and usually when he got there and it was clarified to all present, he had the solution already in mind—a solution that almost invariably received the wholehearted support of those present."20 This described technical management of the first order and also a different kind of heterogeneous engi­neering from that discussed previously, the ability not only to envi­sion a solution but to get it willingly accepted.

As another Peenemunder, Ernst Stuhlinger, and several col­leagues wrote in 1962, “Predecessors and contemporaries of Dr. von Braun may have had a visionary genius equal or superior to his, but none of them had his gift of awakening in others such strong en­thusiasm, faith and devotion, those indispensable ingredients of a successful project team." They added, “It is his innate capability, as a great engineer, to make the transition from an idea, a dream, a dar­ing thought to a sound engineering plan and to carry this plan most forcefully through to its final accomplishment." Finally, Stuhlinger and Frederick Ordway, who knew von Braun in the United States, wrote in a memoir about him, “Regardless of what the subject was— combustion instability, pump failures, design problems, control theory, supersonic aerodynamics, gyroscopes, accelerometers, bal­listic trajectories, thermal problems—von Braun was always fully knowledgeable of the basic subject and of the status of the work. He quickly grasped the problem and he formulated it so that everyone understood it clearly."21 These qualities plus the hiring of a number of able managers of key departments contributed greatly to the de­velopment of the V-2.

Meanwhile, the Space Shuttle marked a radical departure from the pattern of previous launch vehicles. Not only was it (mostly) reus-

able, unlike its predecessors, but it was also part spacecraft, part airplane. In contradistinction to the Mercury, Gemini, and Apollo launch vehicles, in which astronauts had occupied the payload over the rocket, on the shuttle the astronauts rode in and even piloted from a crew compartment of the orbiter itself. The mission com­mander also landed the occupied portion of the Space Shuttle and did so horizontally on a runway. The orbiter had wings like an air­plane and set down on landing gear, as airplanes did. Indeed, the very concept of the Space Shuttle came from airliners, which were not discarded after each mission the way expendable launch vehi­cles had been but were refurbished, refueled, and used over and over again, greatly reducing the cost of operations.

Because of the complex character of the Space Shuttles, their ante­cedents are much more diverse than those of the expendable launch vehicles and missiles discussed previously. Given the scope and length of this book, it will not be possible to cover all of the various aspects of the orbiters in the same way as other launch vehicles.117

Studies of a reusable launch vehicle like the shuttle—as distin­guished from a winged rocket or orbital reconnaissance aircraft/ bomber—date back to at least 1957 and continued through the 1960s. But it was not until the early 1970s that budgetary realism forced planners to accept a compromise of early schemes. Grim fiscal real­ity led to NASA’s decision in the course of 1971-72 to change from a fully reusable vehicle to an only partly reusable stage-and-a-half shuttle concept. Gradually, NASA and its contractors shifted their focus to designs featuring an orbiter with a nonrecoverable external propellant tank. This permitted a smaller, lighter orbiter, reducing the costs of development but imposing a penalty in the form of ad­ditional costs per launch. McDonnell Douglas and Grumman sepa­rately urged combining the external tank with strap-on solid-rocket boosters that would add their thrust to that of the orbiter’s engines. Despite opposition to the use of solids by Marshall Space Flight Center (responsible for main propulsion elements) and in spite of 92 their higher overall cost, solid-rocket boosters with a 156-inch di­Chapter 2 ameter offered lower developmental costs than other options, hence lower expenditures in the next few years, the critical ones from the budgetary perspective.

On January 5, 1972, Pres. Richard M. Nixon had announced his support for development of a Space Shuttle that would give the country “routine access to space by sharply reducing costs in dol­lars and preparation time." By mid-March 1972, the basic configura­tion had emerged for the shuttle that would actually be developed. It included a delta-winged orbiter attached to an external tank with

two solid-rocket boosters on either side of the tank.118 Meanwhile, in February 1970, Marshall released a request for proposals for the study of the space shuttle main engine. Study contracts went to Rocketdyne, Pratt & Whitney, and Aerojet General. The engine was to burn liquid hydrogen and liquid oxygen at a combustion-chamber pressure well above that of any other production engine, includ­ing the Saturn J-2. In July 1971, NASA announced the selection of Rocketdyne as the winner of the competition.119

The SSME featured “staged combustion." This meant that un­like the Saturn engines, whose turbine exhaust contributed little to thrust, in the shuttle the turbine exhaust—having burned with a small amount of oxygen and thus still being rich in hydrogen— flowed back into the combustion chamber where the remaining hydrogen burned under high pressure and contributed to thrust. This was necessary in the shuttle because the turbines had to burn so much fuel to produce the high chamber pressure critical to performance.120

Timing for such an engine was delicate and difficult. As a result, there were many problems during testing—with turbopumps as well as timing. Disastrous fires and other setbacks delayed develop­ment, requiring much analysis and adjustment to designs. In 1972, the shuttle program had expected to launch a flight to orbit by the beginning of March 1978. By then, the expected first-flight date had slipped to March 1979, but various problems caused even a Septem­ber 1979 launch to be postponed. Not until early 1981 was the space shuttle main engine fully qualified for flight. Finally on April 12, 1981, the first Space Shuttle launched, and the main engines per­formed with only a minor anomaly, a small change in mixture ratio caused by radiant heating in the vacuum of space. Some insulation and a radiation shield fixed the problem on subsequent flights. It had taken much problem solving and redesign, but the main en­gines had finally become operational.121

The sophistication of the SSME explained all its problems. “In assessing the technical difficulties that have been causing delays in the development and flight certification of the SSME at full power, it is important to understand that the engine is the most advanced liquid rocket motor ever attempted," wrote an ad hoc committee of the Aeronautics and Space Engineering Board in 1981. “Chamber pressures of more than 3,000 psi, pump pressures of 7,000-8,000 psi, and an operating life of 7.5 hours have not been approached in previous designs of large liquid rocket motors."122

Although more advanced, the SSMEs (producing 375,000 pounds of thrust at sea level and 470,000 pounds at altitude) were consid-

erably less powerful than the Saturn V’s F-1s (with 1.522 million pounds of thrust). At a length of 13.9 feet and a diameter of 8.75 feet, the SSMEs were also smaller than the F-1s, with a length of 19.67 feet and diameter of 12.25 feet. Nevertheless, they were im­pressively large, standing twice as tall as most centers in the Na­tional Basketball Association.123

Because they ignited before launch, the SSMEs did perform some of the same functions for the shuttle that the F-1s did for the Saturn V, but in most respects the twin solid-rocket boosters served as the principal initial sources of thrust. They provided 71.4 percent of the shuttle’s thrust at liftoff and during the initial stage of ascent until about 75 seconds into the mission, when they separated from the orbiter to be later recovered and reused.124

Even before the decision in March 1972 to use solid-rocket boost­ers, Marshall had provided contracts of $150,000 each to the Lock­heed Propulsion Company, Thiokol, United Technology Center, and Aerojet General to study configurations of such motors. Thiokol emerged as winner of the competition, based on its cost and mana­gerial strengths. NASA announced the selection on November 20, 1973.125 The design for the solid-rocket boosters (SRBs) was inten­tionally conservative, using a steel case of the same type employed on Minuteman and the Titan IIIC. The Ladish Company of Cudahy, Wisconsin, made the cases for each segment without welding. Each booster consisted of four segments plus fore and aft sections. The propellant used the same three principal ingredients employed in the first stage of the Minuteman missile. One place shuttle design­ers departed from the Marshall mantra to avoid too much innova­tion lay in the tang-and-clevis joints linking the segments of the SRBs. Although superficially the shuttle joints resembled those for Titan IIIC, they were different in orientation and the use of two O-rings instead of just one.126

In part because of its simplicity compared with the space shuttle main engine, the solid-rocket booster required far less testing than 94 the liquid-propellant engine. Testing nevertheless occasioned sev – Chapter 2 eral adjustments in the design. The SRBs completed their qualifica­tion testing by late May 1980, well before the first shuttle flight.127 Of course, this was well after the first planned flight, so if the main – engine development had not delayed the flights, presumably the booster development would have done so on its own.

The third part of the main shuttle propulsion system was the ex­ternal tank (ET), the only major nonreusable part of the launch ve­hicle. It was also the largest component at about 154 feet in length and 27.5 feet in diameter. On August 16, 1973, NASA selected Mar-

tin Marietta (Denver Division) to negotiate a contract to design, develop, and test the ET. Larry Mulloy, who was Marshall’s project manager for the solid-rocket booster but also worked on the tank, said that the ET posed no technological challenge, although it did have to face aerodynamic heating and heavy loads on ascent. But it had to do so within a weight limit of about 75,000 pounds. As it turned out, this was in fact a major challenge. It came to be fully ap­preciated only after loss of Space Shuttle Columbia on February 1, 2003, to a “breach in the Thermal Protection System on the leading edge of the left wing" resulting from its being struck by “a piece of insulating foam" from the ET. During reentry into the atmosphere, this breach caused aerodynamic superheating of the wing’s alumi­num structure, its melting, and the subsequent breakup of the or – biter under increasing aerodynamic forces.128

The air force had a great deal of influence on the requirements for the shuttle because its support had been needed to get the program approved and make it viable economically. NASA needed a com­mitment from the military that all of its launch needs would be carried on the shuttle. To satisfy DoD requirements, the shuttle had to handle payloads 60 feet long with weights of 40,000 pounds for polar orbits or 65,000 pounds for orbits at the latitude of Kennedy Space Center. On July 26, 1972, NASA announced that the Space Transportation Systems Division of North American Rockwell had won the contract for the orbiters.129

That firm subcontracted much of the work. The design, called a double-delta planform, derived from a Lockheed proposal. The term referred to a wing in which the forward portion was swept more heavily than the rear part. Throughout the development of the shuttle, wind-tunnel testing at a variety of facilities, including those at NASA Langley and NASA Ames Research Centers plus the air force’s Arnold Engineering Development Center, provided data, showing the continuing role of multiple organizations in launch – vehicle design. Before the first shuttle flight in 1981, there was a total of 46,000 hours of testing in various wind tunnels.130

An elaborate thermal protection system (designed primarily for reentry and passage through the atmosphere at very high speeds) and the guidance, navigation, and control system presented many design problems of their own. The launch vehicle that emerged from the involved and cost-constrained development of its many com­ponents was, as the Columbia Accident Investigation Board noted, “one of the most complex machines ever devised." It included “2.5 million parts, 230 miles of wire, 1,060 valves, and 1,440 cir­cuit breakers." Although it weighed 4.5 million pounds at launch, its solid-rocket boosters and main engines accelerated it to 17,500 miles per hour (Mach 25) in slightly more than eight minutes. The three main engines burned propellants fast enough to drain an aver­age swimming pool in some 20 seconds.131

From the first orbital test flight on April 12, 1981, to the end of 1991, there were 44 shuttles launched with 1 failure, an almost 98 percent success rate. On these missions, the shuttles had launched many communications satellites; several tracking and data relay satellites to furnish better tracking of and provision of data to (and from) spacecraft flying in low-Earth orbits; a number of DoD payloads; many scientific and technological experiments; and several key NASA spacecraft.132

Before launching some of these spacecraft, such as Magellan, Ulysses, and the Hubble Space Telescope, however, NASA had en – 96 dured the tragedy of losing the Space Shuttle Challenger and all Chapter 2 of its seven-person crew to an explosion. Since this is not an op­erational history, it is not the place for a detailed analysis, but be­cause the accident reflected upon the technology of the solid-rocket boosters and resulted in a partial redesign, it requires some discus­sion. On the 25th shuttle launch, Challenger lifted off at 11:38 a. m. on January 28, 1986. Even that late in the day, the temperature had risen to only 36°F, 15 ° below the temperature on any previous shuttle launch. Engineers at Morton Thiokol (the name of the firm after 1982 when the Morton Salt Company took over Thiokol Cor-

poration) had voiced reservations about launching in cold tempera­tures, but under pressure to launch in a year scheduled for 15 flights (6 more than ever before), NASA and Morton Thiokol agreed to go ahead. Almost immediately after launch, smoke began escaping from the bottommost field joint of one solid booster, although this was not noticed until postflight analysis. By 64 seconds into the launch, flames from the joint began to encounter leaking hydrogen from the ET, and soon after 73 seconds from launch, the vehicle exploded and broke apart.133

On February 3, 1986, Pres. Ronald Reagan appointed a commission to investigate the accident, headed by former Nixon-administration secretary of state William P. Rogers. The commission determined that the cause of the accident was “the destruction of the seals [O-rings] that are intended to prevent hot gases from leaking through the joint during the propellant burn of the rocket motor." It is pos­sible to argue that the cause of the Challenger accident was faulty assembly of the particular field joint that failed rather than faulty design of the joint. But it seems clear that neither NASA nor Morton Thiokol believed the launch would lead to disaster. The fact that they went ahead with it shows (in one more instance) that rocket engineers still did not have launching such complex vehicles com­pletely “down to a science." Some engineers had concerns, but they were not convinced enough of their validity to insist that the launch be postponed.134

Following the accident there was an extensive redesign of many aspects of the shuttle, notably the field joints. This new design al­legedly ensured that the seals would not leak under twice the an­ticipated structural deflection. Following Challenger, both U. S. policy and law changed, essentially forbidding the shuttle to carry commercial satellites and largely restricting the vehicle to missions both using the shuttle’s unique capabilities and requiring people to be onboard. A concomitant result was the rejuvenation of the air force’s expendable launch-vehicle program. Although the Delta II was the only launcher resulting directly from the 32-month hia­tus in shuttle launches following the accident, the air force also ordered more Titan IVs and later, other expendable launch vehicles. The shuttle became a very expensive launch option because its eco­nomic viability had assumed rapid turnaround and large numbers of launches every year. Yet in 1989 it flew only five missions, in­creased to six in 1990 and 1991.135

As further demonstrated by the Columbia accident, the shuttle clearly was a flawed launch vehicle but not a failed experiment. Its flaws stemmed largely from its nature as an outgrowth of

heterogeneous engineering, involving negotiations of NASA man­agers with the air force, the Office of Management and Budget, and the White House, among other entities. Funding restrictions dur­ing development and other compromises led to higher operational costs. For example, compromises on reusability (the external tank) and employment of solid-rocket motors plus unrealistic projections of many more flights per year than the shuttles ever achieved vir­tually ensured failure in this area from the beginning. Also, as the Columbia Accident Investigation Board pointed out, “Launching rockets is still a very dangerous business, and will continue to be so for the foreseeable future as we gain experience at it. It is unlikely that launching a space vehicle will ever be as routine an undertak­ing as commercial air travel."136

Yet for all its flaws, the shuttle represents a notable engineering achievement. It can perform significant feats that expendable launch vehicles could not. These have ranged from rescue and relaunch of satellites in unsatisfactory orbits to the repair of the Hubble Space Telescope and the construction of the International Space Station. These are remarkable accomplishments that yield a vote for the overall success of the shuttle, despite its flaws and tragedies.

Although flight testing the Saturn launch vehicles went remark­ably well, there were problems, some of which involved the upper stages. For example, on April 4, 1968, during the launch of AS-502 (Apollo 6), there was “an all-important dress rehearsal for the first manned flight" planned for AS-503. Stage-two separation occurred, and all five J-2 engines ignited. Then, at 319 seconds after launch, there was a sudden 5,000-pound decrease in thrust, followed by a

FIG.5.2

The second (S-II) stage of the Saturn V launch vehicle being lifted onto the A-2 test stand at the Mississippi Test Facility (later the Stennis Space Center) in 1967, showing the five J-2 engines that powered this stage. (Photo courtesy of NASA)

cutoff signal to the number two J-2 engine. This signal shut down not only engine number two but number three as well (about a sec­ond apart). It turned out that signal wires to the two engines had been interchanged. This loss of the power from two engines was a severe and unexpected test for the instrument unit (IU), but it ad­justed the trajectory and the time of firing (by about a minute) for the remaining three engines to achieve (in fact, exceed) the planned altitude for separation of the third stage.73

When the IU shut down the three functioning engines in the S-II and separated it from the S-IVB, that stage’s lone J-2 ignited and placed itself, the instrument unit, and the payload in an elongated parking orbit. To do this, the IU directed it to burn 29.2 seconds longer than planned to further compensate for the two J-2s that had
cut off in stage two. The achievement of this orbit demonstrated “the unusual flexibility designed into the Saturn V." However, al­though the vehicle performed adequately during orbital coast, the J-2 failed to restart and propel the spacecraft into a simulated trans­lunar trajectory. After repeated failures to get the J-2 to restart, mis­sion controllers separated the command and service modules from the S-IVB, used burns of the service module’s propulsion system to position the command module for reentry tests, and performed these tests to verify the design of the heat shield, with reentry oc­curring “a little short of lunar space velocity," followed by recovery. Although this is sometimes counted a successful mission (in which Phillips and von Braun both said a crew could have returned safely), von Braun also said, “With three engines out, we just cannot go to the Moon." And in fact, restart of the S-IVB’s J-2 was a primary ob­jective of the mission, making it technically a failure.74

A team of engineers from Marshall and Rocketdyne attacked the unknown problem that had caused the J-2 engine failures. (It turned out to be a single problem for two engines that had failed, one in stage two and the one in stage three that would not restart.) The team, which included Jerry Thomson from the F-1 combustion – instability effort, examined the telemetry data from the flight and concluded that the problem had to be a rupture in a fuel line. But why had it broken?

Increasing pressures, vibrations, and flow rates on test stands, computer analyses, and other tests led engineers to suspect a bellows section in the fuel line. To allow the line to bend around various ob­structions, this area had a wire-braid shielding. On the test stand it did not break from the abnormal strains to which it was subjected. (Artificially severing the line did produce measurements that dupli­cated those from the flight, however.) Finally, Rocketdyne test per­sonnel tried it in a vacuum chamber simulating actual conditions in space. Eight lines tested there at rates of flow and pressures no greater than during normal operations led to failures in the bellows section of all eight lines within 100 seconds. Motion pictures of the tests quickly revealed that in the absence of atmospheric moisture in the vacuum chamber (and in space), frost did not form inside the wire braiding as it had in regular ground tests during cryogenic liquid-hydrogen flow. The frost had kept the bellows from vibrating to the point of failure, but in its absence, a destructive resonance occurred. Engineers eliminated the bellows and replaced them with a stronger design that still allowed the necessary bends. Testing of the fuel-line redesign on the J-2 at the Mississippi Test Facility in August 1968 showed that this change had solved the problem.75

The successful Apollo 8 mission around the Moon verified the success of all the modifications to the launch vehicle since AS-502, with all launch-vehicle objectives for the mission achieved. AS-504 for Apollo 9 was the first Saturn V to use five 1.522-pound-thrust engines in stage one and six 230,000-pound-thrust J-2 engines in the upper stages. It had minor problems with rough combustion but was successful. The Saturn V for AS-505 (Apollo 10) and all subse­quent Apollo missions through Apollo 17 (the final lunar landing) used F-1 and J-2 engines with the same thrust ratings as AS-504. There were comparatively minor adjustments in the launch vehi­cles that followed AS-505—“in timing, sequences, propellant flow rates, mission parameters, trajectories." On all missions there were malfunctions and anomalies that required fine-tuning. For example, evaluations of the nearly catastrophic Apollo 13 flight showed that oscillations in the S-II’s feed system for liquid oxygen had resulted in a drop in pressure in the center engine’s plumbing to below what was necessary to prevent cavitation in the liquid-oxygen pump. Bub­bles formed in the liquid oxygen, reducing pump efficiency, hence 204 thrust from the engine. This led to automatic engine shutdown.

Chapter 5 Although the oscillations remained local, and even engine shut­down did not hamper the mission, engineers at the Space Division of North American Rockwell (as the firm had become following a merger with Rockwell Standard) nevertheless developed two modi­fications to correct the problem. One was an accumulator. It served as a shock absorber, consisting of a “compartment or cavity located in the liquid oxygen line feeding the center engine." Filled with gaseous helium, it served to dampen or cushion the pressures in the liquid-oxygen line. This changed the frequency of any oscillation in the line so that it differed from that of the engines as a whole and the thrust structure, thus prevented coupling, which had caused the problem in Apollo 13. As a backup to the accumulator, engineers installed a “G" switch on the center engine’s mounting beam con­sisting of three acceleration switches that tripped in the presence of excessive low-frequency vibration and shut off the center engine. With these modifications, the J-2 and Saturn V were remarkably successful on Apollo 14 through 17.76

Meanwhile, a much smaller American effort at rocket develop­ment began at the California Institute of Technology (Caltech) in 1936. A graduate student of aerodynamicist Theodore von Karman, Frank J. Malina, together with Edward S. Forman and John W. Par­sons—described respectively by Malina as “a skilled mechanic" and “a self-trained chemist" without formal schooling but with “an un­inhibited fruitful imagination"—began to do research for Malina’s doctoral dissertation on rocket propulsion and flight.22 Gradually, 16 the research of these three men expanded into a multifaceted, pro­Chapter 1 fessional rocket development effort. As with the work under von Braun in Germany, there were many problems to be overcome. The difficulty of both endeavors lay partly in the lack of previous, de­tailed research-and-development reports. It also resulted from the

many disciplines involved. In May 1945, Homer E. Newell, then a theoretical physicist and mathematician at the U. S. Naval Re­search Laboratory, wrote that the “design, construction, and opera­tional use of guided missiles requires intimate knowledge of a vast number of subjects. Among these. . . are aerodynamics, kinemat­ics, mechanics, elasticity, radio, electronics, jet propulsion, and the chemistry of fuels."23 He could easily have added other topics such as thermodynamics, combustion processes, and materials science.

Malina and his associates consulted existing literature. Malina paid a visit to Goddard in 1936 in a fruitless attempt to gather un­published information and cooperation from the secretive New Englander.24 Initially as part of the Guggenheim Aeronautical Lab­oratory at Caltech (GALCIT), directed by von Karman (and after 1943-44 as the Jet Propulsion Laboratory [JPL]), Malina and his staff used available data, mathematics, experimentation, innovations by other U. S. rocketeers, and imagination to develop solid – and liquid-propellant JATOs (jet-assisted takeoff devices), a Private A solid-propellant test rocket, and a WAC Corporal liquid-propellant sounding rocket before Malina left JPL in 1946 and went to Europe. He ultimately became an artist and a promoter of international cooperation in astronautics.25 (Incidentally, in 1942 several of the people at GALCIT founded the Aerojet Engineering Corporation, later known as Aerojet General Corporation, to produce the rocket engines they developed. It became one of the major rocket firms in the country.)26

Under the successive leadership of Louis Dunn and William Pick­ering, JPL proceeded to oversee and participate in the development of the liquid-propellant Corporal and the solid-propellant Sergeant missiles for the U. S. Army. Their development encountered many problems, and they borrowed some engine-cooling technology from the V-2 to solve one problem with the Corporal, illustrating one case where the V-2 influenced U. S. missile development. The Cor­poral became operational in 1954 and deployed to Europe beginning in 1955. Although never as accurate as the army had hoped, it was far superior in this respect to the V-2. At 45.4 feet long, the Corporal was less than a foot shorter than the V-2, but its diameter (2.5 feet) was slightly less than half that of the German missile. However, even with a slightly higher performance than the V-2, its range (99 statute miles) was only about half that of the earlier missile, mak­ing it a short-lived and not very effective weapon.27

In 1953, JPL began working on a solid-propellant replacement for the Corporal, known as the Sergeant. In February 1956, a Sergeant contractor-selection committee unanimously chose the Sperry Gyro-

scope Company (a division of Sperry Rand Corp.) as a co-contractor for the development and ultimate manufacture of the missile. Mean­while, on April 1, 1954, the Redstone Arsenal, which controlled de­velopment of the missile for the army, had entered into a supple­mental agreement with the Redstone Division of Thiokol Chemical Corporation to work on the Sergeant’s solid-propellant motor. The program to develop Sergeant began officially in January 1955.28

The Sergeant missile took longer to develop than originally planned and did not become operational until 1962, by which time the U. S. Navy had completed the much more capable and important Polaris A1 and the U. S. Air Force was close to fielding the significant and successful Minuteman I. The technology of the Sergeant paled by comparison. JPL director Louis Dunn had warned in 1954 that if the army did not provide for an orderly research and development program for the Sergeant, “ill-chosen designs. . . [would] plague the system for many years." In the event, the army did fail to pro­vide consistent funding and then insisted on a compressed sched­ule. This problem was complicated by differences between JPL and Sperry and by JPL’s becoming a NASA instead of an army contractor in December 1958. The result was a missile that failed to meet its in-flight reliability of 95 percent. It met a slipped ordnance sup­port readiness date of June 1962 but remained a limited-production weapons system until June 1968. However, it was equal to its pre­decessor, Corporal, in range and firepower while being only half as large and requiring less than a third as much ground support equip­ment. Its solid-propellant motor could be ready for firing in a matter of minutes instead of the hours required for the liquid-propellant Corporal. An all-inertial guidance system on Sergeant made it vir­tually immune to enemy countermeasures, whereas Corporal de­pended on a vulnerable electronic link to guidance equipment on the ground.29

Thus, Sergeant was far from a total failure. In fact, although not at the forefront of solid-propellant technology by the time of its com­pletion, the army missile made some contributions to the develop­ment of launch-vehicle technology—primarily through a smaller, test version of the rocket. JPL had scaled down Sergeant motors from 31 to 6 inches in diameter for performing tests on various solid propellants and their designs. By 1958, the Lab had performed static 18 tests on more than 300 of the scaled-down motors and had flight – Chapter 1 tested 50 of them—all without failures. Performance had accorded well with predictions. These reliable motors became the basis for upper stages in reentry test vehicles for the Jupiter missile (called Jupiter C) and in the launch vehicles for Explorer and Pioneer satel-

lites, which used modified Redstone and Jupiter C missiles as first stages.30

Because von Braun’s group of engineers developed the Redstone and Jupiter C, this was an instance where purely American and German-American technology blended. It is instructive to compare management at JPL with that in von Braun’s operation in Germany. At JPL, the dynamic von Karman served as director of the project until the end of 1944, when he left to establish the Scientific Ad­visory Board for the U. S. Army Air Forces. Malina held the title of chief engineer of the project until he succeeded von Karman as (acting) director. But according to Martin Summerfield, head of a liquid-propellant section, there was no counterpart at GALCIT/JPL to von Braun at Peenemunde. Instead, Summerfield said, the way the professionals in the project integrated the various components of the rockets and the various developments in fields as disparate as aerodynamics and metallurgy was simply by discussing them as colleagues. He seemed to suggest that much of this was done infor­mally, but like Peenemunde, JPL also had many formal meetings where such issues were discussed. In addition, a research analysis section did a good deal of what later was called systems engineering for JPL.31

Dunn succeeded Malina as acting director of JPL on May 20, 1946, becoming the director (no longer acting) on January 1, 1947. Whereas Malina operated in an informal and relaxed way, Dunn brought more structure and discipline to JPL than had prevailed pre­viously. He was also cautious, hence concerned about the growth of the Lab during his tenure. From 385 employees in June 1946, the number grew to 785 in 1950 and 1,061 in 1953, causing Dunn to create division heads above the section heads who had reported to him directly. There were four such division heads by Septem­ber 1950, with William Pickering heading one on guided-missile electronics.

In August 1954, Dunn resigned from JPL to take a leading role in developing the Atlas missile for the recently established Ramo – Wooldridge Corporation. At Dunn’s suggestion, Caltech appointed Pickering as his successor. A New Zealander by birth, Pickering continued the tradition of having foreign-born directors at JPL (von Karman coming from Hungary and Malina, Czechoslovakia). Easier to know than the formal Dunn, Pickering was also less stringent as a manager. Whereas Dunn had favored a project form of organization, Pickering returned to one organized by disciplines. He remained as director until 1976. Howard Seifert, who had come to GALCIT in 1942 and worked with Summerfield on liquid-propellant develop-

ments, characterized the three JPL directors in terms of an incident when some mechanics cut off the relaxed Malina’s necktie because he was too formal. Seifert said they would never have cut Dunn’s necktie off without losing their jobs, and they would not have cut Pickering’s necktie either, but he would not have fired them for that offense alone. He added that Dunn had a rigid quality but undoubt­edly was extremely capable.32

Despite all the changes in personnel and management from Ma – lina and von Karman, through Dunn to Pickering, and despite the differences in personalities and values, one constant seems to have been a not-very-structured organization, not well suited for dealing with outside industry and the design and fielding of a weapon sys­tem, as distinguished from a research vehicle. Even Dunn’s project organization seems not to have been compatible with the kind of systems engineering soon common in missile development.33

It may be, however, that JPL’s rather loose organization in this period was conducive to innovations that it achieved in both liquid – and solid-propellant rocketry (to be discussed in ensuing chapters). In addition to the direct influence they had upon rocketry, many people from JPL besides Louis Dunn later served in positions of im­portance on other missile and rocket projects, carrying with them, no doubt, much that they had learned in their work at JPL, as well as their talents. Thus, in a variety of ways—some of them incal­culable—the early work at JPL contributed to U. S. rocketry, even though the Lab itself got out of the rocket propulsion business in the late 1950s.34