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

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

Soon after he began working for German Army Ordnance at Kummersdorf in late 1932, Wernher von Braun began experiment­ing with rocket engines, which developed burnthroughs, “igni­tion explosions, frozen valves, fires in cable ducts and numerous other malfunctions." Learning “the hard way," von Braun called in “welding experts, valve manufacturers, instrument makers and pyrotechnicists. . . and with their assistance a regeneratively – cooled motor of 300 kilograms [about 660 pounds] thrust and pro­pelled by liquid oxygen and alcohol was static tested and ready for flight in the A-1 rocket which had been six months a-building." Von Braun’s boss, Walter Dornberger, added that the “650-pound-thrust

chamber . . . gave consistent performance" but yielded an exhaust velocity slower than needed even after the developers “measured the flame temperature, took samples of the gas jet, analyzed the gases, [and] changed the mixture ratio."1

As the staff at Kummersdorf grew, bringing in additional exper­tise, engine technology improved. But only with the hiring of Wal­ter Thiel did truly significant progress occur in the propulsion field. Thiel was “a pale-complexioned man of average height, with dark eyes behind spectacles with black horn rims." Fair-haired with “a strong chin," he joined the experimental station in the fall of 1936. Born in Breslau in 1910, the son of an assistant in the post office, he matriculated at the Technical Institute of Breslau as an under­graduate and graduate student in chemistry, earning his doctorate in 1935. He had served as a chemist at another army lab before com­ing to Kummersdorf.2

Dornberger said Thiel assumed “complete charge of propul­sion, with the aim of creating a 25-ton motor" (the one used for the A-4, providing 25 metric tons of thrust). Because Thiel remained at Kummersdorf until 1940 instead of moving to Peenemunde with the rest of the von Braun group in 1937, testing facilities limited him to engines of no more than 8,000 pounds of thrust from 1936 to 1940. Although Thiel was “extremely hard-working, conscientious, and systematic," Dornberger said he was difficult to work with. Ambitious and aware of his abilities, he “took a superior attitude and demanded. . . devotion to duty from his colleagues [equal to his own]." This caused friction that Dornberger claimed he had to mol­lify. Martin Schilling (chief of the testing laboratory at Peenemunde for Thiel’s propulsion development office and, later, head of the of­fice after Thiel died in a bombing raid in 1943) noted that Thiel was “high strung." He said, “Thiel was a good manager of such a great and risky development program. He was a competent and dynamic leader, and a pusher. At the same time, he was no match to von Braun’s or Steinhoff’s vision and optimism." (Ernst Steinhoff was chief of guidance and control.) 3

A memorandum Thiel wrote on March 13, 1937, after he had been on the job about six months, gives some idea of the state of development of a viable large engine at that time. It also suggests the approach he brought to his task. Although he certainly lacked optimism at some points in his career at Kummersdorf and Peene – munde, he did not betray that failing in his memo. He referred to “a certain completion of the development of the liquid rocket" that had been achieved “during the past years," surely an overstatement in view of the major development effort that remained. “Combustion

chambers, injection systems, valves, auxiliary pressure systems, pumps, tanks, guidance systems, etc. were completely developed from the point of view of design and manufacturing techniques, for various nominal sizes. Thus, the problem of an actually usable liq­uid rocket can be termed as having been solved."

Despite this assessment, he listed “important items" requiring further development. One was an increase in performance of the rocket engine, using alcohol as its fuel. He noted that the engines at Kummersdorf were producing a thermal efficiency of only 22 per­cent, and combustion-chamber losses were on the order of 50 per­cent. Thus, about half of the practically usable energy was lost to incomplete combustion. The use of gasoline, butane, and diesel oil theoretically yielded an exhaust velocity only some 10 percent higher, but measurements on these hydrocarbon fuels showed ac­tual exhaust velocities no higher than those with alcohol. Thiel felt that “for long range rockets, alcohol will always remain the best fuel," because hydrocarbons increased the danger of explosion, pro­duced coking in the injection system, and presented problems with cooling.

He said the way to improved performance lay in exploiting the potential 50 percent energy gain available with alcohol and liquid oxygen. Fuller combustion could come from improving the injec­tion process, relocating the locus for mixing oxygen and fuel into a premixing chamber, increasing the speed of ignition and combus­tion, and increasing chamber pressure by the use of pumps, among other improvements. He knew about the tremendous increases in performance available through the use of liquid hydrogen, but he cited the low temperature of this propellant (-423°F), its high boil – 104 off rate, the danger of explosion, and huge tank volume resulting Chapter 3 from its low specific weight (as the lightest element of all), plus a requirement to insulate its tanks, as “strong obstacles" to its use (as indeed, later proved to be the case).

He made repeated reference to the rocket literature, including a mention of Goddard, but noted that “the development of practi­cally usable models in the field of liquid rockets. . . has far outdis­tanced research." Nevertheless, he stressed the need for coopera­tion between research and development, a process he would follow. He concluded by stating the need for further research in materials, injection, heat transfer, “the combustion process in the chamber," and “exhaust processes." 4

Despite Thiel’s optimism here, Martin Schilling referred in a postwar discussion of the development of the V-2 engine to the “mysteries of the combustion process." Thiel, indeed, said the

combustion process needed further research but did not discuss it in such an interesting way. Dornberger also failed to use such a term, but his account of the development of the 25-ton engine sug­gests that indeed there were mysteries to be dealt with. He pointed out that to achieve complete combustion of the alcohol before it got to the nozzle end of the combustion chamber, rocket research­ers before Thiel had elongated the chamber. This gave the alcohol droplets more time to burn than a shorter chamber would, they thought, and their analysis of engine-exhaust gases seemed to prove the idea correct. “Yet performance did not improve." They realized that combustion was not “homogeneous," and they experienced frequent burnthroughs of chamber walls.

Dornberger said he suggested finer atomization of both oxygen and alcohol by using centrifugal injection nozzles and igniting the propellants after mixing “to accelerate combustion, reduce length of the chamber, and improve performance." Thiel, he said, devel­oped this idea, then submitted it to engineering schools for research while he used the system for the 1.5-ton engine then under devel­opment. It took a year, but he shortened the chamber from almost 6 feet to about a foot. This increased exhaust speed to 6,600 and then 6,900 feet per second (from the roughly 5,300 to 5,600 feet per second in early 1937). This was a significant achievement, but with it came a rise in temperature and a decrease in the chamber’s cool­ing surface. Thiel “removed the injection head from the combus­tion chamber" by creating a “sort of mixing compartment," which removed the flames from the brass injection nozzles. This kept them, at least, from burning.5

In conjunction with the shortening of the combustion chamber, Thiel also converted the shape from cylindrical to spherical to en­compass the greatest volume in available space. This also served to reduce pressure fluctuations and increase the mixing of the propel­lants. Until he could use a larger test stand at Peenemunde, how­ever, Thiel was restricted in scaling up these innovations in the 1.5-ton engine to the full 25 tons. He thus went to an intermediate size of 4.2 tons that he could test at Kummersdorf, and he moved from one injector in the smaller engine to three in the larger one. Each had its own “mixing compartment" or “pot," and the clus­tering actually increased the efficiency of combustion further. But to go from that arrangement to one for the 25-ton engine created considerable problems of scaling up and of arranging the 18 “pots" that the researchers designed for the A-4 combustion chamber. At first, Thiel and his associates favored an arrangement of six or eight larger injectors around the sides of the chamber, but von Braun sug-

gested 18 pots of the size used for the 1.5-ton engine, arranged in concentric circles on the top of the chamber. Schilling said this was a “plumber’s nightmare" with the many oxygen and alcohol feed lines that it required, but it avoided the problems of combustion in – stability—as we now call it—that other arrangements had created.6

Cooling the engine remained a problem. Regenerative cooling used on earlier, less efficient engines did not suffice by itself for the larger engine. Oberth had already suggested the solution, film cool­ing—introducing an alcohol flow not only around the outside of the combustion chamber (regenerative cooling) but down the inside of the wall and the exhaust nozzle to insulate them from the heat of combustion by means of a “film" of fuel. Apparently, others in the propulsion group had forgotten this suggestion, and it is not clear that the idea as applied to the 25-ton engine came from Oberth. Several sources agree that diploma engineer Moritz Pohlmann, who headed the propulsion design office at Kummersdorf after August 1939, was responsible. Tested on smaller engines, the idea proved its validity, so on the 25-ton engine, there were four rings of small holes drilled into the chamber wall that seeped alcohol along the in­side of the motor and nozzle. This film cooling took care of 70 per­cent of the heat from the burning propellants, the remainder be­ing absorbed into the alcohol flowing in the regenerative cooling jacket on the outside of the chamber. Initially, 10 percent of the fuel flow was used for film cooling, but Pohlmann refined this by “oozing" rather than injecting the alcohol, without loss of cooling efficiency.7 Whether this procedure emanated from Oberth or was independently discovered by Pohlmann, it was an important inno­vation with at least the technical details worked out by Pohlmann.

106 Thiel’s group had to come up with a pumping mechanism to Chapter 3 transfer the propellants from their tanks to the injectors in the pots above the combustion chamber. The large quantities of propellant that the A-4 would use made it impractical to feed the propellants by nitrogen-gas pressure from a tank (as had been done on the earlier A-2, A-3, and A-5 engines). Such a tank would have had to be too large and heavy to provide sufficient pressure over the 65-second burning time of the engine, creating unnecessary weight for the A-4 to lift. This, in turn, would have reduced its effective performance. In 1937 Thiel had mentioned that there was a the need for pumps to increase the chamber pressure and that some pumps had already been developed. Indeed, von Braun had already begun working in the middle of 1935 with the firm of Klein, Schanzlin & Becker, with factories in southwestern and central Germany, on the develop­ment of turbopumps. In 1936 he began discussions with Hellmuth

Walter’s engineering office in Kiel about a “steam turbine" to drive the pumps.8

In the final design, a turbopump assembly contained separate cen­trifugal pumps for alcohol and oxygen on a common shaft, driven by the steam turbine. Hydrogen peroxide powered the pumps, con­verted to steam by a sodium permanganate catalyst. It operated at a rate of more than 3,000 revolutions per minute and delivered some 120 pounds of alcohol and 150 pounds of liquid oxygen per second, creating a combustion-chamber pressure of about 210 pounds per square inch. This placed extreme demands on the pump technology of the day, especially given a differential between the heat of the steam ( + 725°F) and the boiling point of the liquid oxygen (-297°F).9

Moreover, the pumps and turbine had to weigh as little as possible to reduce the load the engine had to lift. Consequently, there were problems with the development and manufacture of both devices. Krafft Ehricke, who worked under Thiel after 1942, said in 1950 that the first pumps “worked unsatisfactorily" so the development “transferred to Peenemunde." He claimed that Peenemunde also de­veloped the steam generator. Schilling suggested this as well, writing that for the steam turbine, “we borrowed heavily from" the Walter firm at Kiel. He said a “first attempt to adapt and improve a torpedo steam generator [from Walter’s works] failed because of numerous details (valves, combustion control)." Later, a successful version of the steam turbine emerged, and Heinkel in Bavaria handled the mass production. As for the pumps, there are references in Peenemunde documents as late as January 1943 to problems with them but also to orders for large quantities from Klein, Schanzlin & Becker.10

The problems with the pumps included warping of the pump housing because of the temperature difference between the steam and the liquid oxygen; cavitation because of bubbles in the propel­lants; difficulties with lubrication of the bearings; and problems with seals, gaskets, and choice of alloys (all problems that would recur in later U. S. missiles and rockets). The cavitation problem was especially severe since it could lead to vibrations in the com­bustion chamber, resulting in explosions. The solution came from redesigning the interior of the pumps and carefully regulating the internal pressure in the propellant tanks to preclude the formation of the bubbles.11

Ehricke also reported that development of “control devices for the propulsion system, i. e. valves, valve controls, gages, etc." pre­sented “especially thorny" problems. The items available from commercial firms either weighed too much or could not handle the propellants and pressure differentials. A special laboratory at Peene-

munde had to develop them during the period 1937 to 1941, with a pressure-reducing valve having its development period extended until 1942 before it worked satisfactorily.12

Technical institutes contributed a small but significant share of the development effort for both the engine and the pumps. A profes­sor named Wewerka of the Technical Institute in Stuttgart provided valuable suggestions for solving design problems in the turbopump. He had written at least two reports on the centrifugal turbopumps in July 1941 and February 1942. In the first, he investigated dis­charge capacity, cavitation, speed relationships, and discharge and inlet pressures on the alcohol pump, using water instead of alcohol as a liquid to pass experimentally through the pump. Because the oxygen pump had almost identical dimensions to those of the alco­hol pump, he merely calculated corrections to give values for the oxygen pump with liquid oxygen flowing through it instead of wa­ter through the alcohol pump. In the second report, he studied both units’ efficiencies, effects of variations in the pump inlet heads upon pump performance, turbine steam rates, discharge capacities of the pump, and the pumps’ impeller design. He performed these tests with water at pump speeds up to 12,000 revolutions per minute.13

Schilling pointed to important work that Wewerka and the Tech­nical Institute in Stuttgart had done in the separate area of nozzle design, critical to achieving the highest possible performance from the engine by establishing as optimal an expansion ratio as possible. This issue was complicated by the fact that an ideal expansion ra­tio at sea level, where the missile was launched, quickly became less than ideal as atmospheric pressure decreased with altitude. Wewerka wrote at least four reports during 1940 studying such 108 things as the divergence of a Laval nozzle and the thrust of the jet Chapter 3 discharged by the nozzle. In one report in February, he found that a nozzle divergence of 15 degrees produced maximum thrust. Ger­hard Reisig, as well as Schilling, agreed that this was the optimal exit-cone half angle for the A-4. In his account of engine develop­ment, Reisig, chief of the measurement group under Steinhoff until 1943, also gives Wewerka, as well as Thiel, credit for shortening the nozzle substantially. In another report, Wewerka found that the nozzle should be designed for a discharge pressure of 0.7 to 0.75 at­mosphere, and Reisig says the final A-4 nozzle was designed for 0.8 atmosphere.14

Schilling also pointed to other professors, Hase of the Technical Institute of Hannover and Richard Vieweg of the Technical Insti­tute of Darmstadt, for their contributions to the “field of power – plant instrumentation." Other “essential contributions" that Schil-

ling listed included those of Schiller of the University of Leipzig for his investigations of regenerative cooling, and Pauer and Beck of the Technical Institute of Dresden “for clarification of atomiza­tion processes and the experimental investigation of exhaust gases and combustion efficiency, respectively." In an immediate postwar interview at Garmisch-Partenkirchen, an engineer named Hans Lindenberg even claimed that the design of the A-4’s fuel-injection nozzles “was settled at Dresden." Lindenberg had been doing re­search on fuel injectors for diesel engines at the Technical Institute of Dresden from 1930 to 1940. Since 1940, partly at Dresden and partly at Peenemunde, he had worked on the combustion chamber of the A-4. His claim may have constituted an exaggeration, but he added that Dresden had a laboratory for “measuring the output and photographing the spray of alcohol jets." Surely it and other techni­cal institutes contributed ideas and technical data important in the design of the propulsion system.15

Along similar lines, Konrad Dannenberg, who worked on the combustion chamber and ignition systems at Peenemunde from mid-1940 on, described their development in general terms and then added, “Not only Army employees of many departments par­ticipated, but much of the work was supported by universities and contractors, who all participated in the tests and their evaluation. They were always given a strong voice in final decisions."16

One final innovation, of undetermined origin, involved the igni­tion process, which used a pyrotechnic igniter. In the first step of the process, the oxygen valve opened by means of an electrically activated servo system, followed by the alcohol valve. Both opened to about 20 percent of capacity, but since the propellants flowed only as a result of gravity (and slight pressure in the oxygen tank), the flow was only about 10 percent of normal. When lit by the ig­niter, the burning propellants produced a thrust of about 2.5 tons. When this stage of ignition occurred, the launch team started the turbopump by opening a valve permitting air pressure to flow to the hydrogen peroxide and sodium permanganate tanks. The perman­ganate solution flowed into a mixing chamber, and as soon as pres­sure was sufficient, a switch opened the peroxide valve, allowing peroxide to enter the mixing chamber. When pressure was up to 33 atmospheres as a result of decomposing the hydrogen peroxide, the oxygen and alcohol valves opened fully, and the pressure on the tur­bines in the pumps caused them to operate, feeding the propellants into the combustion chamber. It required only about three-quarters of a second from the time the valve in the peroxide system was elec­trically triggered until the missile left the ground.17

Even after the propulsion system was operational, the propulsion group had by no means solved Schilling’s “mysteries of the combus­tion process." The engine ultimately developed an exhaust velocity of 6,725 feet/second, which translated into a specific impulse of 210 pounds of thrust per pound of propellant burned per second (lbf-sec/lbm), the more usual measure of performance today. Quite low by later standards, this was sufficient to meet the requirements set for the A-4 and constituted a remarkable achievement for the time. As von Braun said after the war, however, “the injector for the A-4 [wa]s unnecessarily complicated and difficult to manufac­ture." Certainly the 18-pot design of the combustion chamber was inelegant. And despite all the help from an excellent staff at Peene – munde and the technical institutes, Thiel relied on a vast amount of testing. Von Braun said, “Thiel’s investigations showed that it required hundreds of test runs to tune a rocket motor to maximize performance," and Dannenberg reported “many burn-throughs and chamber failures," presumably even after he arrived in 1940.18

But through a process of trial and error, use of theory where it was available, further research, and testing, the team under von Braun and Thiel had achieved a workable engine that was sufficient to do the job. As late as 1958 in the United States, “The development of almost every liquid-propellant rocket ha[d] been plagued at one time or another by the occurrence of unpredictable high-frequency pres­sure oscillations in the combustion chamber"—Schilling’s “mys­teries" still at work. “Today [1958], after some fifteen years of con­centrated effort in the United States on liquid-propellant rocket development, there is still no adequate theoretical explanation for combustion instability in liquid-propellant rockets," wrote a no – 110 table practitioner in the field of rocketry.19

Chapter 3 That the propulsion team at Kummersdorf and Peenemunde was able to design a viable rocket engine despite the team’s own and later researchers’ lack of fundamental understanding of the com­bustion processes at work shows their skill and perseverance. It also suggests the fundamental engineering nature of their endeavor. Their task was not necessarily to understand all the “mysteries" (although they tried) but to make the rocket work. Their work con­stituted rocket engineering, not rocket science, because they still did not fully understand why what they had done was effective, only that it worked.

Even without a full understanding of the combustion process, the propulsion group went on to design engines with better injec­tors. They did so for both the Wasserfall antiaircraft missile and the A-4, although neither engine went into full production. Both fea-

tured an injector plate with orifices so arranged that small streams of propellants impinged upon one another. The streams produced oscillations in the engine (combustion instability), but the develop­ers found the correct angle of impingement that reduced (but never completely eliminated) the oscillations (characterized by chug­ging and screeching). They also designed a cylindrical rather than a spherical combustion chamber for the A-4, but it had a slightly lower exhaust velocity than the spherical engine.20

Under difficult, wartime conditions, in-house contributions and those from technical institutes and industry came together through discussions among the contributors at Kummersdorf and Peene – munde. The pooling of their expertise probably contributed in innu­merable ways to the progress of technological development, but the

process can only be partially documented. Certainly, technical re­ports written by both staff at Peenemunde and people at the techni­cal institutes contributed to the fund of engineering knowledge that Peenemunde passed on to the United States. Germans from Peene­munde immigrated to the United States after the war, carrying their knowledge and expertise with them; but in addition, much of the documentation of the engineering work done in Germany was cap­tured by U. S. forces at the end of the war, moved to Fort Eustis, Vir­ginia, and even translated. The full extent of what these documents contributed to postwar rocketry is impossible to know, but the infor­mation was available to those engineers who wanted to avail them­selves of it. Finally, many actual V-2 missiles, captured and taken to the United States, also provided a basis for postwar developments that went beyond the V-2 but started with its technology.

Despite the experience with Centaur and the Saturn upper-stage en­gines, the main engines for the Space Shuttle presented a formidable challenge, mainly because of the extreme demands placed upon the engines in a system that also used solid-propellant rocket boosters

but still required a great deal of thrust from the main engines. In a partly reusable system, NASA’s requirements for staged combus­tion and extremely high chamber pressure made development of the space shuttle main engines (SSMEs) extraordinarily difficult.

The story of this development began in one sense on June 10, 1971, when—with the general configuration of the Space Shuttle still in flux—Dale D. Myers, NASA’s associate administrator for Manned Space Flight, communicated to the directors of the Manned Spacecraft Center (MSC), the Marshall Space Flight Center (MSFC), and the Kennedy Space Center (KSC) the management plan for the Space Shuttle. This gave lead-center responsibilities to MSC but retained general direction of the program at NASA Headquarters in Washington, D. C. MSC would have responsibility for system engineering and integrating the components, with selected person­nel from MSFC and KSC collocated in Houston to support this ef­fort. Marshall would have responsibility for the main propulsion elements, while Kennedy would manage the design of launch and recovery infrastructure and launch operations.77

Myers had managed the Navaho missile effort for North Ameri­can and had become vice president of the Space Division, where he had been the general manager for the Apollo spacecraft. He had also overseen North American Rockwell’s studies for the Space Shuttle. In addition, he had experience with aircraft projects. Thus, he came to his new job with a strong background in all aspects of the shuttle (as launch vehicle, spacecraft, and airplane). At Marshall, von Braun had moved on in 1970 to become deputy associate administrator for planning at NASA Headquarters.

His deputy director for scientific and technical matters, Eberhard Rees, had succeeded him as Marshall center director until Rees re­tired in 1973, to be succeeded by Rocco A. Petrone, who had earned a doctorate in mechanical engineering from MIT. Petrone had come from NASA Headquarters and returned there in 1974. He was suc­ceeded by William R. Lucas, a chemist and metallurgist with a doctorate from Vanderbilt University who had worked at Redstone Arsenal and then Marshall since 1952 and become deputy direc­tor in 1971. Petrone reorganized Marshall, deemphasizing in-house capabilities to oversee and test large project components and giving more authority to project officers, less to lab directors, a change Myers approved. As Rees put it, Myers was “somewhat allergic to ‘too much’ government interference" with contractors, preferring less stringent oversight than Marshall had provided in the past.78

In February 1970, Marshall had released a request for proposals for the Phase B (project definition) study of the space shuttle main

engine. Contracts went to Rocketdyne, Pratt & Whitney, and Aero­jet General. The engine was to burn liquid hydrogen and liquid oxy­gen in a 6:1 ratio at a combustion-chamber pressure of 3,000 pounds per square inch, well above that of any production engine, including the Saturn J-2, which had featured a pressure of about 787 pounds at the injector end of the 230,000-pound-thrust version. The shut­tle engine was to produce a thrust of 415,000 pounds of force at sea level or 477,000 pounds at altitude. Although Rocketdyne had built the J-2 and a development version, the J-2S, with a thrust of 265,000 pounds and chamber pressure of 1,246 pounds per square inch, Pratt & Whitney had been developing an XLR129 engine for the Air Force Rocket Propulsion Laboratory. The engine actually delivered 350,000 pounds of thrust and operated at a chamber pres­sure of 3,000 pounds per square inch during 1970.79 Pratt & Whitney thus seemed to have an advantage in the competition.

At Rocketdyne, seasoned rocket engineer Paul Castenholz, who had helped troubleshoot the F-1 combustion-instability and injector problems and had been project manager for the J-2, headed the SSME 206 effort as its first project manager, even though he was a corporate Chapter 5 vice president. He saw that there was not time to build sophisti­cated turbopumps, so he decided to build a complete combustion chamber fed by high-pressure tanks. The NASA study contract did not provide funds for such an effort, so Castenholz convinced North American Rockwell to approve up to $3 million in company funds for the effort. By 1971, testing the engine at Nevada Field Laboratory near Reno, Rocketdyne had a cooled thrust chamber that achieved full thrust for 0.45 second. The thrust was 505,700 pounds at a cham­ber pressure of 3,172 pounds per square inch, exceeding the perfor­mance of Pratt & Whitney’s XLR129 by a considerable margin.80

Funding constraints led to combining Phase C and D contracts (to include actual vehicle design, production, and operations), so on March 1, 1971, Marshall released to the three contractors a request for proposals to design, develop, and deliver 36 engines. In July NASA selected Rocketdyne as winner of the competition, but Pratt & Whitney protested the choice to the General Accounting Office (GAO) as “manifestly illegal, arbitrary and capricious, and based upon unsound, imprudent procurement decisions." On March 31, 1972, the GAO finally decided the case in favor of Rocketdyne, with the contract signed August 14, 1972. This protest delayed de­velopment, although Rocketdyne worked under interim and letter contracts until the final contract signature.81

It was not until May 1972 that Rocketdyne could begin signifi­cant work on the space shuttle main engine in something close to its

final configuration, although some design parameters would change even after that. By then, however, NASA had decided on a “parallel burn" concept in which the main engines and the solid-rocket boosters would both ignite at ground level. The space agency had already determined in 1969 that the engine would employ staged combustion, in which the hydrogen-rich turbine exhaust contrib­uted to combustion in the thrust chamber. It was the combination of high chamber pressure and staged combustion that made the SSMEs a huge step forward in combustion technology. In the mean­time, they created great problems for the shuttle, but one of them was not combustion instability, the usual plague for engine devel­opment. Castenholz and his engineers had started development of the engine with an injector based on the J-2, which had shown good stability. For the shuttle, according to Robert E. Biggs, a member of the SSME management team at Rocketdyne since 1970, the firm had added “two big preventors [of instability] on an injector that was basically stable to begin with." He evidently referred to coaxial baffles, and they seem to have worked.82

The XLR129 had been a staged-combustion engine, and its success had given NASA and industry the confidence to use the same concept on the shuttle. But timing for such an engine’s igni­tion was both intricate and sensitive, as Rocketdyne and Marshall would learn. Rocketdyne’s design used two preburners with low – and high-pressure turbopumps to feed each of the propellants to the combustion chamber and provide the required high pressure. The XLR129 had used only a single preburner, but two of them provided finer control for the shuttle in conjunction with an engine-mounted computer, subcontracted to Honeywell for development. This computer monitored and regulated the propulsion system during start, automatically shut it down if it sensed a problem, throttled the thrust during operation, and turned off the engine at mission completion.83

By the winter and spring of 1974, development of the Honeywell controller had experienced difficulties relating to its power supply and interconnect circuits. These problems attracted the attention of NASA administrator James C. Fletcher and his deputy, George M. Low. The latter commented that Rocketdyne had done a “poor job" of controlling Honeywell, which itself had done a “lousy job" and was in “major cost, schedule, and weight difficulty." Rocketdyne had fallen behind in converting test stands at Santa Susana for test­ing components of the engine, including turbopumps. A cost over­run of about $4 million required congressional reprogramming. In a program that was underfunded to begin with, this was intolerable,

so pressured by Fletcher and Low, Rockwell International, as the firm became in 1973, shifted Castenholz to another position, replac­ing him ultimately with Dominick Sanchini, a tough veteran who had led development of the main-engine proposal in 1971. Despite 27 successful years devoted to the rocket business, with important achievements to his credit, Castenholz would no longer contribute directly to launch-vehicle development.

Meanwhile, about the same time, Marshall made J. R. Thompson its project manager for the space shuttle main engine. Trained as an aeronautical engineer at Georgia Institute of Technology, where he graduated in 1958, Thompson had worked for Pratt & Whitney before becoming a liquid-propulsion engineer at Marshall on the Saturn project in 1963, the year he earned his master’s degree in mechanical engineering at the University of Florida. He became the space engine section chief in 1966, chief of the man/systems inte­gration branch in 1969, and main-engine project manager in 1974.84

In May 1975, both component testing (at Santa Susana) and pro­totype engine testing began, the latter at NASA’s National Space 208 Technology Laboratories (the former Mississippi Test Facility). Typ- Chapter 5 ically, there was about a month between testing of a component at Santa Susana and a whole engine in Mississippi. But test personnel soon learned that the highly complicated test hardware at Santa Susana was inadequate. As Robert Frosch, NASA administrator, said in 1978, “We have found that the best and truest test bed for all major components, and especially turbopumps, is the engine it­self." Consequently, because of insufficient equipment to test com­ponents as well as engines, the program gradually ceased testing at Santa Susana between November 1976 and September 1977.85

There were many problems during testing, especially with turbo­pumps and timing. The timing problems involved “how to safely start and shut down the engine." After five years of analysis, as Biggs explained, Rocketdyne engineers had “sophisticated com­puter models that attempted to predict the transient behavior of the propellants and engine hardware during start and shutdown." Test personnel expected that the engine would be highly sensitive to minute shifts in propellant amounts, with the opening of valves be­ing time-critical. Proceeding very cautiously, testers took 23 weeks and 19 tests, with replacement of eight turbopumps, to reach two seconds into a five-second start process. It took another 12 weeks, 18 tests, and eight more turbopump changes to momentarily reach the minimum power level, which at that time was 50 percent of rated thrust. Eventually Biggs’s people developed a “safe and repeat­able start sequence" by using the engine-mounted computer, also

called the main-engine controller. “Without the precise timing and positioning" it afforded, probably they could not have developed even a satisfactory start process for the engine, so sensitive was it.

Following purging of the propulsion system with dry nitrogen and helium to eliminate moisture (which the propellants could freeze if left in the system), then a slow cooldown using the cryo­genic propellants, full opening of the main fuel valve started the fuel flow that initially occurred from the latent heating and expan­sion that the hardware (still warmer than the liquid hydrogen) im­parted to the cryogenic propellant. However, the flow was pulsating with a pressure oscillation of about two cycles per second (hertz) until chamber pressure in the main thrust chamber stabilized after 1.5 seconds. Then oxidizer flowed to the fuel and oxidizer preburn­ers and the main combustion chamber in carefully timed sequence such that liquid oxygen arrived at the fuel preburner 1.4 seconds af­ter the full opening of the main fuel valve, at the main combustion chamber at 1.5 seconds, and at the oxidizer preburner at 1.6 seconds. Test experience revealed that a key time was 1.25 seconds into the priming sequence. If the speed of turbine revolution in the high – pressure fuel turbopump at that precise moment was not at least 4,600 revolutions per minute, the engine could not start safely. So, 1.25 seconds became a safety checkpoint.

If any “combustor prime" coincided with a downward oscilla­tion (dip) in the fuel flow, excessively high temperatures could re­sult. Other effects of inaccurate timing could be destruction of the high-pressure oxidizer turbopump. Also, a 1 to 2 percent error in valve position or a timing error of as little as a tenth of a second could seriously damage the engine. Because of these problems, the first test to achieve 50 percent of rated thrust occurred at the end of January 1976. The first test to reach the rated power level was in January 1977. Not until the end of 1978 did the engineers achieve a final version of the start sequence that precluded the problems they encountered over more than three years of testing. There were also issues with shutdown sequencing, but they were less severe than those with safe engine start, especially critical because astronauts would be aboard the shuttle when it started.86

One major instance of problems with the high-pressure tur­bopumps occurred on March 12, 1976. Earlier tests of the high – pressure liquid-hydrogen pump, both at Santa Susana and in Mis­sissippi, had revealed significant vibration levels, but not until the March 12 test had engineers recognized this as a major problem. The prototype-engine test on that day was supposed to last 65 sec­onds to demonstrate a 50 percent power level, rising to 65 percent

for a single second. The test did demonstrate 65 percent power for the first time, but engineers had to halt the test at 45.2 seconds because the high-pressure fuel turbopump was losing thrust. After the test, the pump could not be rotated with a tool used to test its torque. Investigation showed that there had been a failure of the turbine-end bearings supporting the shaft. Test data showed a ma­jor loss in the efficiency of the turbines plus a large vibration with a frequency about half the speed of the pump’s rotation. Experts immediately recognized this as characteristic of subsynchronous whirl, an instability in the dynamics of the rotors.

Although recognizing the problem, test personnel evidently did not know what to do about it in a system whose turbine-blade stresses and tip speeds were still close to the limits of technology in 1991 and must have been at the outskirts of the state of the engi­neering art 15 years earlier. In any event, to speed up a solution, the program assembled a team that ultimately included the premier ro – tordynamics experts in government, industry, and academia, from the United States and Great Britain. The pump was centrifugal, 210 driven by a two-stage turbine 11 inches in diameter that was de­Chapter 5 signed to deliver 75,000 horsepower at a ratio of 100 horsepower per pound, an order-of-magnitude improvement over previous tur­bopumps. The team studied previous liquid-hydrogen turbopumps like that on the J-2, which had exhibited subsynchronous whirl. Following a test program involving engine and laboratory tests, as well as those on components and subsystems, the investigators found 22 possible causes; the most likely appeared to be hydro­dynamic problems involving seals that had a coupling effect with the natural frequency of the rotating turbines. Efforts to decrease the coupling effect included damping of the seals and stiffening the shaft. The fixes did not totally end the whirl but did delay its in­ception from 18,000 revolutions per minute, which was below the minimum power level, to 36,000 revolutions per minute, above the rated power level.

As these design improvements increased operating speeds, in­vestigators learned that a mechanism unrelated to subsynchronous whirl was still overheating the turbine bearings, which had no lu­brication but were cooled by liquid hydrogen. The team’s extensive analysis of the cooling revealed that a free vortex was forming at the bottom of the pump’s shaft where coolant flowed. This vortex reduced the pressure, hence the flow of coolant. In a piece of cut- and-try engineering, designers introduced a quarter-sized baffle that changed the nature of the vortex and allowed more coolant to flow. This fix and the elevation of the whirl problem to above the rated

power level permitted long-duration tests of the engine for the first time by early 1977.87

This problem with the fuel pump had delayed the program, but it was not as diabolical as explosions in the high-pressure liquid – oxygen turbopump. If a fire started in the presence of liquid oxygen under high pressure, it incinerated the metal parts, usually remov­ing all evidence that could lead to a solution. After solution of the fuel-pump-whirl problem, there were four fires in the high-pressure oxygen turbopump between March 1977 and the end of July 1980. This turbopump was on the same shaft as the low-pressure oxygen turbopump that supplied liquid oxygen to the preburners. The com­mon shaft rotated at a speed of nearly 30,000 revolutions per min­ute. The high-pressure pump was centrifugal and provided as much as 7,500 gallons of liquid oxygen at a pressure higher than 4,500 pounds per square inch. An essential feature of the pump’s design was to keep the liquid oxygen fully separated from the hydrogen – rich gas that drove its turbines. To ensure separation, engineers and technicians had used various seals, drains, and purges.

Despite such precautions, on March 24, 1977, an engine caught fire and burned so severely it removed most physical evidence of its cause. Fortunately, investigators used data from instrumentation to determine that the fire started near a complex liquid-oxygen seal. Since it was not evident what a redesign should involve, testing on other engines resumed, indicating that one of the purges did not prevent the mixing of liquid oxygen and fluids draining from hot gas. On July 25, 1977, engineers tried out a new seal intended as an interim fix. But it worked so well it became the permanent solu­tion, together with increasing the flow rate of the helium purge and other measures.88

On September 8, 1977, there was another disastrous fire originat­ing in the high-pressure oxygen turbopump. Data made it clear that the problem involved gradual breakdown of bearings on each end of the turbopump’s shaft, but there was no clear indication of the cause. Fixes included enhanced coolant flow, better balance in the rotors, heavier-duty bearings, and new bearing supports. The other two fires did not involve design flaws but did entail delays. In 1972, the shuttle program had expected to launch a flight to orbit by early March 1978. The engine and turbopump problems and many others involving the propulsion system were but some of the causes for not making that deadline, but engines would have kept the shuttle from flying that early if everything else had gone as planned.

By March 1978, the expected first-flight date had slipped to March 1979, but an engine fire and other problems caused even a Septem-

ber 1979 launch to be postponed. By early 1979, turbopumps were demonstrating longer periods between failures. By 1980, engines were expected to reach 10,000 seconds of testing apiece, a figure it had taken the entire program until 1977 to reach for all engines com­bined. But there continued to be failures in July and November 1980. Thus, not until early 1981 was the space shuttle main engine fully qualified for flight. Problems had included turbine-blade failures in the high-pressure fuel turbopump, a fire involving the main oxidizer valve, failures of nozzle feed lines, a burnthrough of the fuel pre­burner, and a rupture in the main-fuel-valve housing. But finally on April 12, 1981, the first Space Shuttle lifted off. After much trouble­shooting and empirical redesign, the main engines finally worked.89

The large number of problems encountered in the development of the space shuttle main engines resulted from its advanced design. The high chamber and pump pressures as well as an operating life of 7.5 hours greatly exceeded those of any previous engine. Each shut­tle had three main engines, which could be gimballed 10.5 degrees in each direction in pitch and 8.5 degrees in yaw. The engines could be throttled over a range from 65 to 109 percent of their rated power level (although there had been so many problems trying to demon­strate the 109 percent level in testing that it was not available on a routine basis until 2001). Moreover, the 65 percent minimum power
level (changed from the original 50 percent level) was unavailable at sea level because of flow separation. During launch, the three main engines ignited before the solid-rocket boosters. When computers and sensors verified that they were providing the proper thrust level, the SRBs ignited. To reduce vehicle loads during the period of maxi­mum dynamic pressure (reached at about 33,600 feet some 60 sec­onds after liftoff) and to keep vehicle acceleration at a maximum of 3 Gs, the flight-control system throttled back the engines during this phase of the flight. Throttling also made it feasible to abort the mis­sion either with all engines functioning or with one of them out.90

At 100 percent of the rated power level, each main engine pro­vided 375,000 pounds of thrust at sea level and 470,000 pounds at altitude. The minimum specific impulse was more than 360 lbf-

FIG. 5.4

The space shuttle main engine firing during a test at the National Space Technologies Laboratories (later the Stennis Space Center), Janu­ary 1, 1981, showing the regenerative cooling tubes around the circumference of the combustion chamber.

(Photo courtesy of NASA)

sec/lbm at sea level and 450 lbf-sec/lbm at altitude. This was sub­stantially higher than the J-2 Saturn engine, which had a sea-level specific impulse of more than 290 lbf-sec/lbm and one at altitude of more than 420 lbf-sec/lbm. The J-2’s thrust levels were also substan­tially lower at 230,000 pounds at altitude. Not only were the SSMEs much more powerful than the earlier engines using liquid-hydrogen technology but they were also vastly more sophisticated.91