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

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

While JPL’s rocket development proceeded, there were several other efforts in the field of rocketry that contributed to the development of U. S. missile and launch-vehicle technology. Some of them started earlier than Malina’s project, notably those associated with what became (in 1934) the American Rocket Society. This organization, first called the American Interplanetary Society, had its birth on April 4, 1930. Characteristically, although Goddard became a mem­ber of the society, founding member Edward Pendray wrote, “Mem – 20 bers of the Society could learn almost nothing about the techni – Chapter 1 cal details of his work." Soon, society members were testing their own rockets with the usual share of failures and partial successes. But their work “finally culminated in. . . a practical liquid-cooled regenerative motor designed by James H. Wyld." This became the

first American engine to apply regenerative cooling (described by Oberth) to the entire combustion chamber. Built in 1938, it was among three engines tested at New Rochelle, New York, Decem­ber 10, 1938. It burned steadily for 13.5 seconds and achieved an exhaust velocity of 6,870 feet per second. This engine led directly to the founding of America’s first rocket company, Reaction Mo­tors, Inc., by Wyld and three other men who had been active in the society’s experiments. Also, it was from Wyld that Frank Ma – lina learned about regenerative cooling for the engines developed at what became JPL, one example of shared information contributing to rocket development.35

Reaction Motors incorporated as a company on December 16, 1941. It had some successes, including engines for tactical missiles; the X-1 and D-558-2 rocket research aircraft; and an early throttle­able engine for the X-15 rocket research airplane that flew to the edge of space and achieved a record speed of 6.7 times the speed of sound (Mach 6.7). Reaction Motors had never been able to develop many rockets with large production runs nor engines beyond the size of the X-15 powerplant. On April 30, 1958, Thiokol, which had become a major producer of solid-propellant rocket motors, merged with Reaction Motors, which then became the Reaction Motors Division of the Thiokol Chemical Corporation. In 1970, Thiokol decided to discontinue working in the liquid-propellant field; and in June 1972, Reaction Motors ceased to exist.36

Despite its ultimate failure as a business, the organization had shown considerable innovation and made lasting contributions to U. S. rocketry besides Wyld’s regenerative cooling. A second impor­tant legacy was the so-called spaghetti construction for combus­tion chambers, invented by Edward A. Neu Jr. Neu applied for a patent on the concept in 1950 (granted in 1965) but had developed the design earlier. It involved preforming cooling tubes so that they became the shells for the combustion chamber when joined to­gether, creating a strong yet light chamber. The materials used for the tubes and the methods of connecting them varied, but the firm used the basic technique on many of its engines on up through the XLR99 for the X-15. By the mid-1950s, other firms picked up on the technique or developed it independently. Rocketdyne used it on the Jupiter and Atlas engines, Aerojet on the Titan engines. Later, Rocketdyne used it on all of the combustion chambers for the Sat­urn series, and today’s space shuttle main engines still use the concept.37

Another important early contribution to later missile and launch- vehicle technology came from a group formed by naval officer Rob-

ert C. Truax. He had already begun developing rockets as an ensign at the Naval Academy. After service aboard ship, he reported to the navy’s Bureau of Aeronautics from April to August 1941 at “the first jet propulsion desk in the Ship Installation Division." There, he was responsible for looking into jet-assisted takeoff for seaplanes. He then returned to Annapolis, where he headed a jet propulsion project at the Naval Engineering Experiment Station (where Robert God­dard was working separately on JATO units nearby). Truax’s group worked closely with Reaction Motors and Aerojet on projects rang­ing from JATOs to tactical missiles. Among the officers who worked under Truax was Ensign Ray C. Stiff, who discovered that aniline and other chemicals ignited spontaneously with nitric acid. This information, shared with Frank Malina, became critical to JPL’s ef­forts to develop a liquid-propellant JATO unit. In another example of the ways technology transferred from one organization or firm to another in rocket development, once Stiff completed his five years of service with the navy, he joined Aerojet as a civilian engineer. He rose to become vice president and general manager of Aerojet’s liq­uid rocket division (1963) and then (1972) president of the Aerojet Energy Conversion Company. In 1969 he became a Fellow of the American Institute of Aeronautics and Astronautics (into which the American Rocket Society had merged) for “his notable contri­butions in the design, development and production of liquid rocket propulsion systems, including the engines for Titan I, II, and III."38

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

MX-774 B

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

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

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

Another major part of the shuttle propulsion system was the exter­nal tank (ET), the only major nonreusable portion of the launch ve­hicle. It was the largest (and, when loaded, the heaviest part of the Space Shuttle), at about 154 feet in length and 27.5 feet in diameter. NASA issued a request for proposals for design and construction to Chrysler, McDonnell Douglas, Boeing, and Martin Marietta on April 2, 1973. All four bidders submitted their proposals on May 17. The source selection board gave the highest technical ratings to 214 Martin Marietta and McDonnell Douglas. Martin argued that it Chapter 5 alone among the bidders had relevant experience, with the Titan III core vehicle being situated between two large solid-rocket motors. Martin’s costs were by far the lowest of the four, although the board recognized that it was bidding below true expected costs—“buying in" as it was called. But as NASA deputy administrator George Low said, “We nevertheless strongly felt that in the end Martin Mari­etta costs would, indeed, be lower than those of any of the other contenders." Consequently, on August 16, 1973, NASA selected Martin Marietta (Denver Division) to negotiate a contract for the design, development, and testing of the external tank, a selection that, this time, the other competitors did not protest. NASA re­quired assembly of the structure at the Marshall-managed Michoud facility near New Orleans.92

The external tank seemed to some to pose few technological de­mands. James Kingsbury, head of Marshall’s Science and Engineer­ing Directorate, stated, “There was nothing really challenging tech­nologically in the Tank. . . . The challenge was to drive down the cost." Similarly, Larry Mulloy, who was Marshall’s project manager for the solid-rocket booster but also worked on the tank, stated, “There was no technological challenge in the building of the Ex­ternal Tank. The only challenge was building it to sustain the very large loads that it has to carry, and the thermal environment that it is exposed to during ascent" and do so within a weight limit of about 75,000 pounds. As it turned out, however, there was in fact

a major challenge, only fully appreciated 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 an area of the external tank known as the bipod ramp. During reentry into the atmosphere, this breach allowed aerodynamic superheating of the wing’s aluminum structure, melting, and the subsequent breakup of the orbiter under increasing aerodynamic forces.93

The external tank had to carry the cryogenic liquid-hydrogen and liquid-oxygen propellants for the three shuttle main engines. It also served as the “structural backbone" for the shuttle stack and had to withstand substantial heating as the shuttle accelerated to supersonic speeds through the lower atmosphere, where dynamic pressures were high. This heating was much more complex than on a launch vehicle like the Saturn V. At the top, the tank needed only to withstand the effects of high-speed airflow. But further down, the tank’s insulation had to encounter complex shock waves as it passed through the transonic speed range (roughly Mach 0.8 to 1.2). As the airflow became supersonic, shock waves came from the nose of the orbiter, the boosters, and the structural attachments connect­ing the tank, boosters, and orbiter. As the waves impinged on the sides of the external tank, they created heating rates up to 40 British thermal units per square foot per second. This was much smaller than the heating of a nose cone reentering the atmosphere, but it was substantial for the thin aluminum sheeting of which the exter­nal tank was formed to reduce weight.94

As designers examined the requirements for the external tank, they found that not even the arrangement of the hydrogen and oxy­gen tanks involved a simple application of lessons from the Centaur and Saturn. In both, the liquid-hydrogen tank was above the liquid – oxygen tank. Since liquid oxygen was 6 times as heavy as liquid hydrogen, this arrangement made it unnecessary to strengthen the hydrogen tank to support the heavier oxygen during liftoff. Also, with the lighter hydrogen on top, the inertial forces necessary to change the attitude of the vehicle were lower than would have been the case had the reverse arrangement prevailed. For the shuttle, however, the engines were not directly under the tanks, as was the case for the Saturn upper stages and Centaur. Instead, they were off to one side. With the heavy oxygen tank on the bottom of the ex­ternal tank, its weight would have created an inertial force difficult to overcome by gimballing of the SSMEs and the SRB nozzles. Es­pecially after the separation of the solid boosters, the weight of the oxygen tank would have tended to cause the orbiter to spin around

FIG. 5.5 Technical drawing of the Space Shuttle vehicle showing its component parts, including the external tank. (Photo courtesy of NASA)

SPACE SHUTTLE VEHICLE

ORBITER

MS F C 7F> SA 4106—2C

the tank’s center of gravity. Placing the oxygen tank on top moved the shuttle stack’s center of gravity well forward, making steering much more feasible. But it also forced designers to make the liquid – hydrogen tank (and also an intertank structure between it and the oxygen tank) much sturdier than had been necessary on the Saturn upper stages.95

This, in turn, compounded a problem with the ET’s weight. The initial empty weight allowance had been 78,000 pounds, but in 1974, the Johnson Space Center in Houston (renamed from the Manned Spacecraft Center in 1973) reduced the goal to 75,000 pounds. Moreover, NASA asked Martin Marietta if it could not only reduce the weight but do so at no additional cost. In fact, the space agency suggested that it would be helpful actually to reduce the cost. Even though Marshall lowered the safety factor for the ET, the initial tank used on shuttle flights 1-5 and 7 weighed some 77,100 pounds. But through concerted efforts, Martin Marietta was able to achieve a 10,300-pound weight reduction for the lightweight tanks first used on flight 6 of the shuttle. The firm attained the weight re­duction through a variety of design changes, including eliminating some portions of longitudinal structural stiffeners in the hydrogen tank, using fewer circumferential stiffeners, milling some portions of the tank to a lower thickness, using a different type of aluminum that was stronger and thus allowed thinner sections, and redesign­ing anti-slosh baffling.96

The resultant external tank included a liquid-hydrogen tank that constituted 96.66 feet of the ET’s roughly 154 feet in length. It had semi-monocoque design with fusion-welded barrel sections, forward and aft domes, and five ring frames. It operated at a pres­sure range of 32 to 34 pounds per square inch and contained an anti-vortex baffle but no elaborate anti-slosh baffles, because the lightness of the liquid hydrogen made its sloshing less significant than that of liquid oxygen. The feed line from the tank allowed a maximum flow rate of 48,724 gallons per minute from its 385,265- gallon (237,641-pound) capacity. The intertank structure was much shorter at 22.5 feet. Made of both steel and aluminum, it, too, was semi-monocoque in structure with a thrust beam, skin, stringers, and panels. It contained instrumentation and a device called an umbilical plate for supply of purge gas, detection of hazardous gas escaping from the tanks, and boil-off of hydrogen gas while on the ground. The intertank also had a purge system that removed the highly combustible propellants if they escaped from their tanks or plumbing fixtures.

Above the intertank was the liquid-oxygen tank. Its 49.33 feet of length, combined with those of the intertank and the liquid – hydrogen tank, exceeded the total length of the ET because it and its liquid-hydrogen counterpart extended into the intertank. The liquid-oxygen tank was an aluminum monocoque structure oper­ating with a pressure range of 20-22 pounds per square inch. It al­lowed a maximum of 2,787 pounds (19,017 gallons) of liquid oxy­gen to flow to the main engines when they were operating at 104 percent of their rated thrust. Containing both anti-slosh and anti­vortex mechanisms, the tank had a capacity of 143,351 gallons, or 1,391,936 pounds, of oxidizer.97

The thermal-protection system for the external tank had to withstand the complex aerodynamic heating generated by the shut­tle structure and keep the cryogenic propellants from boiling. The tank was coated with an inch of foam similar to that used on the Saturn S-II. Unlike the S-II insulation, however, which had to pro­tect only against boil-off and not against the formation of ice on the foam from the liquid hydrogen and the liquid oxygen, that on the ET could not permit ice formation, because if ice came off the tank during launch, it could easily damage the critical and delicate thermal-protection system on the orbiter. Thus, the external tank’s insulation had to be thicker than that on the S-II. It was in fact so effective that despite the extreme temperatures inside the tanks, the surface of the insulation felt “only slightly cool to the touch." For the first two shuttle flights, there was a white, fire-retardant

latex coating on top of the foam, but thereafter, following testing to determine that the foam alone provided sufficient protection during ascent, the shuttle team dispensed with this coating, saving 595 pounds and leaving the orange foam to add its distinctive color to the white of the orbiter and solid-rocket boosters at launch.98

Like the main engines, the external tank underwent extensive testing before the first shuttle launch. The entire propulsion sys­tem was, of course, designed under Marshall oversight, with Cen­ter Director Lucas continuing von Braun’s practice of using weekly notes for overall communication and systems engineering. In view of this, the Columbia Accident Investigation Board was perhaps un­fairly critical in 2003 when it wrote:

In the 1970s, engineers often developed particular facets of a design (structural, thermal, and so on) one after another and in relative isolation from other engineers working on different facets. Today, engineers usually work together on all aspects of a design as an inte­grated team. The bipod fitting [in the area where foam separated on 218 Columbia’s last flight] was designed from a structural standpoint,

Chapter 5 and the application process for foam (to prevent ice formation) and

Super Lightweight Ablator (to protect from high heating) were devel­oped separately.

However, the board went on to note in all fairness:

It was—and still is—impossible to conduct a ground-based, simul­taneous, full-scale simulation of the combination of loads, airflows, temperatures, pressures, vibration, and acoustics the External Tank experiences during launch and ascent. Therefore, the qualification testing did not truly reflect the combination of factors the bipod would experience during flight. Engineers and designers used the best methods available at the time: test the bipod and foam under as many severe combinations as could be simulated and then in­terpolate the results. Various analyses determined stresses, thermal gradients, air loads, and other conditions that could not be obtained through testing.99

Design requirements specified that the Space Shuttle system not shed any debris, but on the first shuttle flight, the external tank produced a shower of particles, causing engineers to say they would have been hard-pressed to clear Columbia for flight if they had known this would happen. When the bipod ramp lost foam on shuttle flight 7, wind-tunnel testing showed that the ramp area was

designed with an aerodynamically too steep angle, and designers changed the ramp angle from 45 degrees to a shallower 22 to 30 degrees. However, this and a later “slight modification to the ramp impingement profile" failed to prevent the destruction of Space Shuttle Columbia on February 1, 2003. It is beyond the scope of this history to discuss the Columbia accident further, but despite advances in analytical capabilities until 2003, the board was unable to pinpoint the “precise reasons why the left bipod foam ramp was lost."100

This was so even though the board included a staff of more than 120 people aided by about 400 NASA engineers in a lengthy and extensive investigation lasting months. The reasons a definitive ex­planation was impossible included the fact that foam did not “have the same properties in all directions" or the “same composition at every point." It was “extremely difficult to model analytically or to characterize physically. . . in even relatively static conditions, much less during the launch and ascent of the Shuttle." Factors that may have caused the foam to separate and damage the wing in­cluded “aerodynamic loads, thermal and vacuum effects, vibrations, stress in the External Tank structure, and myriad other conditions" including “wind shear, associated Solid Rocket Booster and Space Shuttle Main Engine responses, and liquid oxygen sloshing in the External Tank." Even in 2003, “Non-destructive evaluation tech­niques for determining External Tank foam strength have not been perfected or qualified." 101

With statements such as, “In our view, the NASA organizational culture had as much to do with this accident as the foam," the ac­cident investigation board clearly implicated more than technol­ogy in the causes of the Columbia accident. But a major cause was NASA and contractor engineers’ failure to understand the reasons for and full implications of foam shedding from the external tank. As well-known space commentator John Pike said, “The more they study the foam, the less they understand it." And as a newspaper article stated, “Getting every ounce of the foam to stick to the ex­ternal tank has bedeviled NASA engineers for 22 years. . . . Why foam falls off any area of the tank remains a scientific mystery." In the more sober language of the CAIB report, “Although engineers have made numerous changes in foam design and application in the 25 years the External Tank has been in production, the problem of foam-shedding has not been solved." 102

Whatever the larger causes of the accident, from the perspective of this book, this was but one more instance in which engineers did not have the design, development, and operation of rockets “down

to a science." Despite countless billions of dollars spent on research­ing, developing, and operating a large number of missiles and rock­ets; despite a great deal of effort on NASA’s and contractors’ parts to understand and correct this particular problem, there were aspects of rocketry (including this one) that eluded the understanding of engineers and even scientists such as investigation board member Douglas D. Osheroff, a Nobel Prize-winning physicist from Stan­ford University. Osheroff had conducted some simple experiments with foam that helped him understand the “basic physical proper­ties of the foam itself" but also demonstrated “the difficulty of un­derstanding why foam falls off the external tank." As he said, “At­tempts to understand [the] complex behavior and failure modes" of the components of the shuttle stack were “hampered by their strong interactions with other systems in the stack."103

Meanwhile, at the end of World War II, U. S. Army Ordnance became especially interested in learning more about missile technology from the German engineers under von Braun; from their technical documents; and from firing actual V-2 missiles in the United States. According to von Karman, the army air forces (AAF) had a chance to get involved in long-range missile development before Army Ord­nance stepped into the breach. In 1943, Col. W. H. Joiner, liaison officer for the AAF Materiel Command at Caltech, suggested a re­port on the possibilities for long-range missiles. Frank Malina and his Chinese colleague Hsue-shen Tsien prepared the report, stating that a 10,000-pound liquid-propellant rocket could carry a projectile 22 75 miles. Issued November 20, 1943, this was the first report to

Chapter 1 bear the rubric JPL, according to von Karman. But the AAF did not follow up on the opportunity. Army Ordnance then contracted with JPL to conduct research resulting in the Corporal missile; it also initiated Project Hermes on November 15, 1944.

Under contract to the army for this effort, the General Electric Company (GE) agreed to perform research and development on guided missiles. The Ordnance Department sought to provide the company’s engineers with captured V-2s for study and test firing. Through Project Overcast, which became Project Paperclip in 1946, von Braun and a handful of his associates flew to the United States on September 18, 1945. Three groups (about 118) who had worked at Peenemunde arrived in the United States by ship between Novem­ber 1945 and February 1946, ending up at Fort Bliss in Texas, across the state border from White Sands Proving Ground in New Mexico, where the V-2s would be launched once they were assembled from parts captured in Germany.39

Using U. S.-manufactured parts when German ones were dam­aged or not available in sufficient quantity, between April 16, 1946, and September 19, 1952, GE and the army launched 73 V-2s at White Sands. The last GE flight was on June 28, 1951, with the ensuing five flights conducted by the army alone. Depending upon the crite­ria of failure, some 52 to 68 percent of the flights conducted under GE auspices succeeded, but many of the failures still yielded useful information. Many of the V-2 flights carried scientific experiments, with the areas of experimentation ranging from atmospheric phys­ics to cosmic radiation measurements. The explorations helped spawn the field of space science that blossomed further after the birth of NASA in 1958.40

One of the major contributions to launch-vehicle technology by Project Hermes came from the Bumper-WAC project. This combined the V-2s with JPL’s WAC Corporal B rockets in a two-stage con­figuration. The flights sought to provide flight tests of vehicle sepa­ration at high speeds, to achieve speeds and altitudes higher than could then otherwise be obtained, and to investigate such phenom­ena as aerodynamic heating at high speeds within the atmosphere. Since a number of groups were already looking into launching sat­ellites, these matters were of some near-term importance because more than one rocket stage was commonly required to put a satel­lite in orbit.41

On the fifth Bumper-WAC launch from White Sands (February 24, 1949), the second stage reached a reported altitude of 244 miles and a maximum speed of 7,553 feet per second. These constituted the greatest altitude and the highest speed reached by a rocket or missile until that date. The highly successful launch demonstrated the va­lidity of the theory that a rocket’s velocity could be increased with a second stage. Also shown was a method of igniting a rocket engine at high altitude, offering a foundation for later two – and multiple-

stage vehicles. The final two launches of the Bumper program oc­curred at the newly activated Joint Long Range Proving Grounds in Florida, which later became Cape Canaveral Air Force Station and had previously been the site of the Banana River Naval Air Station until taken over by the air force on September 1, 1948.42

While preparations were being made for the Bumper-WAC proj­ect, six GE engineers assisted in launching a V-2 from the fantail of the aircraft carrier USS Midway on September 6, 1947. The missile headed off at an angle from the vertical of 45 degrees even though the ship had almost no pitch or roll. After the missile nearly hit the bridge and then straightened out momentarily, it began tumbling. This test, called Operation Sandy, could hardly have been encour­aging for those in the Bureau of Aeronautics who wanted the navy to develop ballistic missiles for shipboard use. Even less so were the results of two other V-2s, loaded with propellants, that were 24 knocked over on purpose aboard the Midway in part of what was Chapter 1 called Operation Pushover. They detonated, as expected. Finally, on December 3, 1948, a V-2 with propellants burning was also toppled at White Sands on a mocked-up ship deck. This produced a huge blast during which structural supports cracked. There was a rupture

in the deck itself, whereupon alcohol and liquid oxygen ran through the hole and ignited. These tests contributed significantly to the navy’s negative attitude toward the use of liquid propellants aboard ship in what became the Polaris program.43

With two exceptions (discussed later), the tactical Hermes mis­siles GE developed appear not to have had a large influence on launch-vehicle technology. They began as weapons projects but be­came simply test vehicles in October 1953 and were canceled by 1954. Nevertheless, according to a GE publication written in 1965, the engineers who worked on Project Hermes “formed a unique nucleus of talent that was fully realized when they took over many top management slots in General Electric missile and space efforts in later years." GE also later supplied the first-stage engine for the Vanguard launch vehicle and the second stage for the short-lived Atlas-Vega vehicle, among others. Finally, once Project Hermes ended in 1954-55, the people who worked on it shifted their focus to a ballistic reentry vehicle for the air force.44

Even apart from GE’s direct role in launch-vehicle technology, clearly there was a technical legacy of considerable importance from Project Hermes, including the firing of the V-2s. But opinions have been mixed about the significance of this legacy.45 One em­phatic proponent of its importance, Julius H. Braun, had worked on the project while serving in the army. He spoke of “a massive tech­nology transfer to the U. S. rocket and missile community" from the V-2. “There was a steady flow of visitors from industry, govern­ment labs, universities and other services," he added. Referring to the Germans under von Braun’s leadership, he observed, “Propul­sion experts from the team traveled to North American Aviation [NAA] in Southern California to assist in formulating a program to design, build and test large liquid propellant rocket engines. This program led to the formation of the NAA Rocketdyne division." Braun opined that “the rapid exploitation and wide dissemination of captured information" from the Germans “saved the U. S. at least ten years during the severe R&D [research and development] cut­backs of the postwar period." He listed a great many U. S. missiles and rockets that “incorporated components. . . derived from the V-2 and its HERMES follow-on programs."46

As the quotations from Braun suggest, the V-2 technology was a starting point for many efforts by both the Germans and U. S. engi­neers to develop more advanced technology for the rockets and mis­siles that followed. But the Americans did not simply copy V-2 tech­nology; they went beyond it. Even in assembling the V-2s for firing in this country, GE engineers and others had to develop modifications

to German technology in making replacement parts. Firms that con­tracted to make the parts undoubtedly learned from the effort.

Although many visitors surely picked up a great deal of useful in­formation from the Germans, as did the GE engineers themselves, JPL propellant chemist Martin Summerfield, who questioned von Braun and others about the rocket engine for the V-2 on April 19, 1946, evidently learned little from the interchange. Summerfield had already learned from his own research at JPL the kinds of tech­nical lessons that the Germans imparted. Others from JPL had a good chance to look over the V-2s while testing the Corporal and firing the Bumper-WAC. According to Clayton Koppes, “They con­cluded there was relatively little they wanted to apply to their proj – ects."47 Of course, even the Corporal did borrow some technology from the V-2, but JPL had developed much else independently.

Other visitors to the German engineers had less experience with rocketry and probably benefited greatly from talking with them, as Braun said. Projects like the Bumper-WAC added to the fund of engineering data. However, as Braun had mentioned, funding for rocket-and-missile development was limited in the immediate post-World War II United States. Firing the V-2s had cost about $1 million per year through 1951, but until the United States became alarmed about a threat from the Soviet Union’s missiles and war­heads, there would not be a truly major U. S. effort to go much be­yond the technologies already developed.48

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

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

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

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

Castable composite propellants grew out of a grant Theodore von Karman and Frank Malina arranged with the National Academy of Sciences (NAS) Committee on Army Air Corps Research in Janu­ary 1939. With the $1,000 allotment, Malina and

his associates at GALCIT studied jet-assisted takeoff (JATO) of air­craft and prepared a proposal for research on the subject. This led to an NAS contract for $10,000, effective July 1, 1939, and to sub­sequent contracts with the army air corps and the navy for JATO units (with both liquid and solid propellants). JATOs actually used rocket (rather than jet) thrust to help heavily loaded aircraft take off on a short runway.1

The key individual in the development of the first castable com­posite propellant was John W. Parsons, a propellant chemist on Ma – lina’s team. Parsons, largely self-taught, had taken some chemistry courses from the University of Southern California in 1935-36 but did not graduate. He worked as a chemist for Hercules Powder Com­pany in Los Angeles from 1932 to 1934 and then was chief chemist for Halifax Explosives Company in Saugas, California, from 1934 to 1938. In 1939-40, Parsons sought a solution to the problem of controlled burning for many seconds in a solid-propellant rocket motor. This was critical to the development of a JATO unit. It was he, apparently, who conceived the concept of “cigarette-burning" at only one end of the propellant. But repeated tests of powder, com­pressed into a chamber and coated with a variety of substances to form a seal with the chamber wall, resulted in explosions. Authori­ties von Karman consulted advised that a powder rocket could burn for only two or three seconds.

Not satisfied with this expert opinion, von Karman characteris­tically turned to theory for a solution. He devised four differential equations describing the operation of the rocket motor and handed them to Malina for solution. In solving them, Malina discovered that, theoretically, if the combustion chamber were completely filled by the propellant charge, if the physical properties of the pro­pellant and the ratio of the area of burning propellant to the throat area of the chamber’s nozzle remained constant, thrust also would do likewise and there would be no explosions. Encouraged by these findings, Parsons and others came up with a compressed powder design that worked effectively (after one initial explosion) for 152 successive motors used in successful flight tests of JATO units on an Ercoupe aircraft in August 1941, convincing the navy to contract for a variety of assisted takeoff motors.

After storage under varying temperatures, however, the motors usually exploded. Parsons then found a solution to that problem. Apparently watching a roofing operation about June 1942, he con­cluded that asphalt as a binder and fuel mixed with potassium per­chlorate as an oxidizer would yield a stable propellant. This proved to be true. Thus the theory of von Karman and Malina combined with

the practical knowledge and imagination of Parsons to produce a castable, composite solid propellant that, with later improvements, made large solid-propellant rockets possible. A fundamental tech­nological breakthrough, this formed the basis of many later castable propellants with much higher performance than asphalt-potassium perchlorate.2

Meanwhile, to produce its JATOs (both solid and liquid), five members of the GALCIT project (Malina, von Karman, Martin Summerfield, Parsons, and Edward S. Forman), plus von Karman’s lawyer, Andrew G. Haley, formed the Aerojet Engineering Corpora­tion in March 1942 (Aerojet General Corporation after its acquisi­tion by General Tire and Rubber Company in 1944-45). Aerojet did much business with the army air forces and navy for JATO units during the war and became by 1950 the largest rocket engine manu­facturer in the world and a leader in research and development of rocket technology. Until the acquisition by General Tire, Aerojet and the GALCIT project maintained close technical relations.3

The initial asphalt-potassium perchlorate propellant—known as GALCIT 53—did not have a particularly impressive performance
compared, for example, with ballistite (a double-base composition). But it operated effectively at temperatures down to 40°F. At even lower temperatures, however, GALCIT 53 cracked. It also melted in the tropical sun and was very smoky when burning. This last char­acteristic restricted visibility for the takeoff of second and follow-on aircraft using JATO units on a single runway. Consequently, re­searchers at GALCIT and its successor, JPL, began searching for an elastic binder with storage limits beyond GALCIT 53’s extremes of -9°F to 120°F. In particular, a young engineer named Charles Bartley, who was employed at JPL from June 1944 to August 1951, began ex­amining synthetic rubbers and polymers, eventually hitting upon a liquid polysulfide compound designated LP-2 as a solid-propellant binder. The Thiokol Chemical Corporation made it for sealing air­craft tanks and other applications.4

Like many innovations, LP-2 had resulted from an initial, inad­vertent discovery. In 1926, Joseph C. Patrick, a physician who found chemistry more interesting than medicine, sought an inexpensive way to produce antifreeze from ethylene using sodium polysulfide as a hydrolyzing agent. His procedure yielded a synthetic rubber in­stead of antifreeze. It led him to cofound Thiokol, which marketed the material in the form of gaskets, sealants, adhesives, and coatings (the polysulfide polymer being resistant to weather, solvents, and electrical arcing). Then in 1942, Patrick and an employee, H. L. Fer­guson, found a way to make the first liquid polymer that included no volatile solvent yet could be cured to form a rubberlike solid. During World War II it was used to seal fuel tanks in aircraft, gun turrets, fuselages, air ducts, and the like.5

Before learning of LP-2, Bartley and his associates at JPL had tried a variety of moldable synthetic rubbers as both binders and fuels, in­cluding Buna-S, Buna-N, and neoprene. Neoprene had the best prop­erties for use as a binder and burned the best of the lot, but mold­ing it required high pressures. Like the extrusion process used with double-base propellants (forcing them through a die), this made the production of large propellant grains (masses of propellant) imprac­tical. Meanwhile, Thiokol chemists had begun to release data about LP-2. At a meeting of the American Chemical Society, Bartley asked about a liquid that would polymerize to a solid elastomer (rubber­like substance). Frank M. McMillan, who represented Shell Oil in the San Francisco area, knew about Thiokol’s product and shared the information with Bartley, who acquired small quantities from Walt Boswell, Thiokol’s representative for the western United States.6

With encouragement from Army Ordnance and the navy, Bartley—joined by John I. Shafer, a JPL design engineer, and

H. Lawrence Thackwell Jr., a specialist in aircraft structures—be­gan in 1947 to develop a small rocket designated Thunderbird, with a 6-inch diameter. They used it for testing whether polysulfide pro­pellants could withstand the forces of high acceleration that a po­tential large launch vehicle might encounter. Bartley had already found that an end-burning grain of polysulfide propellant did not produce steady thrust but burned faster at first and then leveled off. He attributed this to accelerated burning along the case to form a convex cone, a hypothesis he confirmed by quenching the flame partway through the burn.

To solve the problem of unsteady thrust, the three JPL engineers adopted a grain design that had been developed in Great Britain in the late 1930s but was similar to one developed independently in 226 1946 for double-base propellants by an American, Edward W. Price.

Chapter 6 It featured an internal-burning, star-shaped cavity. This design pro­tected the case from excess heat because the burning was in the middle of the propellant grain. It also provided a constant level of thrust because as the star points burned away, the internal cavity became a cylinder with roughly the same surface area as the initial star. Bartley had read about the star design from a British report he did not specify and had instructed Shafer to investigate it. Shafer found that the government-owned, contractor-operated Allegany Ballistics Laboratory had used the British design in the uncom­pleted Vicar rocket and a scaled-down version named the Curate. Using equations from the ABL report on the projects, Shafer began developing a number of star designs in 1947. Combining a polysul­fide propellant with the star design and casting it in the case so that it bonded thereto, the team under Bartley produced the successful Thunderbird rocket that passed its flight tests in 1947-48.7

Another significant development was the replacement of potas­sium perchlorate as an oxidizer by ammonium perchlorate, which offered higher performance (specific impulse) and less smoke. Ap­parently, the Thunderbird used a propellant designated JPL 100, which contained a mixture of ammonium perchlorate and potas­sium perchlorate in a polysulfide binder. In 1947, however, JPL had developed a JPL 118 propellant that used only ammonium perchlo­rate as an oxidizer together with polysulfide as the binder and a couple of curing agents. Although this propellant had yet to be fully investigated in 1947, by mid-1948 JPL had tested it and showed that it had a specific impulse of at least 198 lbf-sec/lbm at sea level, using an expansion ratio of 10 for the rocket nozzle. This was still relatively low compared with a typical performance of double-base

PROPELLANT

RUBBER-BASE LINER OR RESTRICTION BONDED TO CHAMBER

About this time, under contract to the army, Thiokol produced a T-40 motor intended for use as a JATO unit. As a propellant, it used JPL 100 (rechristened T-10 by Thiokol) in a case-bonded motor de­sign. Also in 1949, Thiokol designed the T-41 motor for the Hughes Aircraft Company’s Falcon missile under development for the air force. This was a shorter version of JPL’s Thunderbird motor. It be­gan production at Elkton, then moved to Huntsville, where a larger version called the T-42 evolved from it.11

According to Edward N. Hall, later an air force colonel who was important in promoting the development of solid propellants, the Falcon tactical (air-to-air) missile contributed “quality control techniques for rubber-base propellants, design data for case-bonded 228 grains, [and] aging characteristics of rubber-based propellants" to Chapter 6 the evolving store of knowledge about solid-propellant technology.

It appears that Thiokol did not make these contributions on its own. JPL provided considerable assistance in an early example of technol­ogy transfer. In October 1947, Charles Bartley of JPL was present at a meeting of Thiokol personnel, representatives of Army Ordnance, and the navy’s Bureau of Aeronautics to discuss the kind of work Thiokol was expected to do in the further development of poly­sulfide propellants. The next day, Bartley met again with Thiokol personnel to relate JPL’s experience with polysulfide-perchlorate propellants.12

In about January or February 1949, a trip report by a Thiokol employee discussed a visit to JPL’s Solid Rocket Section, of which Bartley was the chief. The report covered such matters as the grease used for extracting the mandrel to create the internal cavity in the grain once it cured and igniters that employed black or a special igniter powder. Also discussed were grinding ammonium and po­tassium perchlorate, combining them with the liquid polymer in a vertical mixer, pouring the propellant, preparing the liner for the combustion chamber, and testing. Also reported was a visit to Western Electrochemical Company, which supplied the perchlo­rate. The document concluded with some recommended changes in Thiokol’s operating procedures at Elkton.13

As helpful as JPL’s assistance was, however, the contributions Hall mentions seem to have come also from work done indepen­dently at Thiokol’s Elkton and Huntsville plants. For example, Thiokol discovered that the size of perchlorate particles was impor­tant in motor operation and propellant castability, so it introduced a micromerograph to measure particle size. To reduce the deleteri­ous absorption of moisture by the perchlorates, Thiokol installed

air conditioning in the grinding rooms. The firm determined the optimal mixing time for the propellant and replaced a barium grease JPL had used to extract the mandrel from the middle of the cast propellant after curing with a Teflon coating. This latter step was necessary not only because the grease-affected part of the propellant had to be sanded after extraction but also because the grease had en­folded into the propellant, causing weak areas. Thiokol also intro­duced a “temperature-programmed cure cycle," pressurized curing, and a method of casting that eliminated propellant voids resulting from shrinkage and air bubbles.14 These details provide early exam­ples of information—rare in the literature about rocketry—about which firms introduced specific innovations, illustrating the ways technology sometimes transferred.

One Hermes test missile that did lead to a follow-on effort was the C1. General Electric envisioned it as a three-stage vehicle. How­ever, the company did not continue immediately with its original conception in 1946 because at that time it had insufficient data to design such a missile. In October 1950, Army Ordnance directed the firm to proceed with a feasibility study of Hermes C1 in con­junction with the Germans under von Braun. Meanwhile, in April 1948 Col. Holger Toftoy, heading the Rocket Branch in the Army Ordnance Department, recommended establishment of a rocket laboratory. The result was the reactivation of the World War II Red­stone Arsenal in Huntsville, Alabama, announced by the chief of 26 ordnance on November 18, 1948.49 This became the site for most of Chapter 1 the development of the C1 missile as well as the reason for renam­ing it Redstone.

On October 28, 1949, the secretary of the army approved a move of the guided missile group—formerly known as the Ordnance Research

and Development Division Sub-Office (Rocket)—from Fort Bliss, Texas, to Redstone Arsenal. It became the Ordnance Guided Mis­sile Center. The army officially established the center on April 15, 1950. It took about six months for the people and equipment to be transferred. Those who moved included the German rocket group and some 800 other people, among them civil servants, GE employ­ees, and about 500 military personnel.50

Despite the outbreak of the Korean War in June 1950, the C1/ Redstone missile initially lacked significant funding. On July 10, 1950, Army Ordnance had told the center to study the possibility of developing a tactical missile with a 500-mile range and an ac­curacy that would place half of the warheads in a circle of 1,000 yards in radius (circular error probable, or CEP). Von Braun served as project engineer and put together a preliminary study, which he presented to the Department of Defense’s Research and Develop­ment Board in the fall of 1950. Army Ordnance did not send the center the missile’s initial funding in the amount of $2.5 million until May 1, 1951. Once development began in May, it continued for seven and a half years until the flight test of the last research – and-development vehicle. Proposed military characteristics for the missile changed over time. Using a modified North American Avia­tion (NAA) engine originally designed for the U. S. Air Force’s Na – vaho missile (canceled in 1958), the Redstone incorporated much American as well as German rocket technology. With other Ameri­can companies besides North American working on the missile, it became substantially different vehicle from the V-2.51

Following the usual high number of failures for early rockets dur­ing flight testing, the army deployed the 69-foot, 4-inch Redstone on June 18, 1958, and deactivated it in June 1964, when the faster and more mobile Pershing, also developed at Redstone Arsenal, re­placed it. Costing more per missile than the Corporal and Sergeant, the Redstone also had much higher performance in terms of range, thrust, and payload than its older cousins among army missiles (see table 1.1).52

More than just a missile, however, the Redstone also became the basis for the first stage of the first U. S. true launch vehicle. Known as the Jupiter C because of its use in testing Jupiter-missile compo­nents, the variant of the elongated Redstone used in this launch ve­hicle incorporated larger fuel and oxidizer tanks and a change from alcohol to hydyne (unsymmetrical dimethylhydrazine and diethyl­ene triamine) as the fuel to increase performance. For the upper stages, a combined Redstone Arsenal-JPL team employed six-inch – diameter scale models of the Sergeant missile already used as upper

stages in the Jupiter C tests. With three scaled-down-Sergeant upper stages, the Jupiter C became the Juno I launch vehicle. On Janu­ary 31, 1958, it lifted off from Cape Canaveral and placed the first U. S. satellite, the Explorer I, into orbit. Subsequently, somewhat modified Juno Is placed Explorer III and Explorer IV in orbit, with three other Explorer launches failing for various reasons.53

Explorer I was the United States’ rejoinder to the shocking launch by the Soviet Union on October 4, 1957, of Sputnik I, the world’s first artificial satellite. Soon after Sputnik, the United States en­tered a competition with the Soviets to place a human being in orbit as part of an emerging space race between the two superpow­ers. Known as Project Mercury, the U. S. effort planned to use the air force’s Atlas missiles as launch vehicles for the orbiting cap­sules containing astronauts, but testing of the capsules in subor­bital flight employed modified Redstones as launchers. For these early launches, the Redstone was the only trustworthy booster in the American inventory. The military was still testing the Atlas as well as the Thor and Jupiter missiles. However, it required exten­sive modifications—some 800 in all—to make the Redstone safe for an astronaut in the Mercury capsule. Called the Mercury-Redstone, this version of the missile boosted Astronaut Alan Shepard into a successful suborbital flight on May 5, 1961, followed by Virgil I. Grissom on July 21 of the same year, concluding the involvement of Redstone in Project Mercury and its role as a launch vehicle.54

For the Vanguard launch vehicle, the prime contractor, the Martin Company, chose General Electric (GE) on October 1, 1955, to de­velop the first-stage engine. Although GE had earlier developed an A3-B engine that burned alcohol and liquid oxygen, the firm decided to use kerosene and liquid oxygen for the Vanguard (X-405) engine. To achieve the performance needed to launch satellites, the X-405 114 featured a chamber pressure of 616 pounds per square inch and a Chapter 3 146-second propellant burn. The engine achieved a specific impulse of roughly the 254 lbf-sec/lbm called for in the specifications for the powerplant. The X-405 was regeneratively cooled, the propellants fed by decomposition of hydrogen peroxide to provide the speci­fied chamber pressure. GE was able to deliver the first production engine (P-1) on October 1, 1956. But during static testing, dam­age occurred to the lining of the combustion chambers in engines P-2 and P-3. When chamber liners also failed in the P-4 engine, the schedule had to be delayed to fix the problem. A redesign entailed adjustments to the cooling system and careful attention to injector specifications to prevent combustion instability and local hot spots. GE had to test 15 injectors and six variations in design between January and April 1956 before the firm’s engineers found one that worked. Obviously, the state of the art of injector design did not

allow a clear-cut, quick solution, but the overall result of design and testing was a relatively uncomplicated engine with a minimal number of relays and valves. Redesign had worked, the engine never experiencing a burnthrough in flight.23

The November 1955 Vanguard schedule specified that six test vehicles would launch between September 1956 and August 1957, with the first satellite-launching vehicle lifting off in October 1957. If the project had remained on schedule, conceivably the navy could have launched a satellite about the same time as the Soviet Sput­nik. Unfortunately, problems with both the first – and second-stage engines caused delays. On October 23, 1957, a Vanguard test ve-

hicle (without a satellite but with a prototype first-stage engine) did launch successfully almost three weeks after the Soviet satellite began orbiting. Because of the tremendous pressure from the launch of Sputnik, the navy decided to launch the next test vehicle with a minimal, 3.4-pound satellite aboard. When the White House an­nounced this test, the press seized upon it as the United States’ an­swer to the Soviets. This test (TV-3) was the first with three “live" stages. The intent was for it to test the three stages and, if all went well, launch the satellite.

On December 6, 1957, the launch began. The first stage ignited, but the vehicle rose slowly, “agonizingly hesitated a moment. . . and. . . began to topple [as] an immense cloud of red flame from burning propellants engulfed the whole area." GE and Martin Com­pany technicians pored over records from ground instrumentation, films of the failed launch, and the two seconds of telemetered data from the toppling inferno. Martin concluded that there had been an “improper engine start" because of low fuel-tank pressure. GE said the start had not been improper and blamed the failed launch on a loose fuel-line connection. As it turned out, Martin was correct in part, but the problem was more extensive than low fuel-tank pres­sure. Telemetry data indicated that there had been a high-pressure spike on engine start that GE had not noticed on testing because it had used low-response instrumentation. The pressure spike had destroyed a high-pressure fuel line, resulting in the rocket’s destruc­tion. To solve the problem, engineers increased the period of oxy­gen injection into the combustion chamber (ahead of the fuel) from three to six seconds. With this correction and an increase in the minimum pressure in the fuel tank by 30 percent, the first-stage 116 engine worked without problems in 14 static and flight tests fol – Chapter 3 lowing the disaster. Although the engine was largely successful af­ter its first failure, however, it appears to have contributed only experience and data to later launch-vehicle technology. In March 1959, NASA contracted with General Dynamics and GE to adapt the Vanguard first stage as an upper stage (called the Vega) for the Atlas launch vehicle, but in December 1959, the space agency can­celed the contracts in favor of the DoD-sponsored Agena B upper stage. Thus ended further use of the GE engine.24

While these developments were occurring at Thiokol, JPL worked with a test vehicle named the Sergeant, not to be confused with the later missile of that name. Army Ordnance had authorized the de­velopment of this test vehicle, a sounding rocket with a diameter of 15 inches, which was quite large for the day (although only an inch larger than an aircraft rocket developed at the Naval Ordnance Test Station, Inyokern, California, in 1945 and called the “Big Richard"). Designed with an extremely thin steel case of 0.065 inch and a star­shaped perforation, the Sergeant test vehicle was expected to attain an altitude of up to 700,000 feet while carrying a 50-pound payload. Static tests with a thicker case in February 1949 showed that a poly­sulfide grain of that diameter could function without deformation.

But when the JPL researchers (including Bartley, Shafer, and Thackwell) shifted to the thinner case, the result was 12 succes­sive explosions through April 27, 1950. At this point, JPL director Louis G. Dunn canceled the project for the sounding rocket and re­duced all solid-propellant work at the laboratory to basic research. The researchers soon determined that the causes of the explosions included a chamber pressure that was too high for the thin case and points on the star configuration that were too sharp, causing cracks as pressure built up. An easy solution would have been a thicker case and rounded points on the star. When Dunn canceled the proj­ect, Thackwell took his knowledge of solid-propellant rocketry to Thiokol’s Redstone Division in Huntsville.15