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

Подпись: A Vanguard launch vehicle undergoing a static test at Cape Canaveral in September 1955. (U.S. Navy photo courtesy of NASA)

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

Подпись: 115 Propulsion with Alcohol and Kerosene Fuels, 1932-72 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

The Beginnings, Goddard and Oberth, 1926-45

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

Подпись: 9 German and U.S. Missiles and Rockets, 1926-66 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

FIG. 1.1

Robert H. Goddard and the first known liquid – propellant rocket ever to have been launched, Auburn, Massachusetts, March 16, 1926. (Photo courtesy of NASA)



Chapter 1


The Beginnings, Goddard and Oberth, 1926-45

FIG. 1.2

Technical drawing of Goddard’s 1926 liquid – propellant rocket. (Photo courtesy of NASA)


The Beginnings, Goddard and Oberth, 1926-45

Подпись: 11 German and U.S. Missiles and Rockets, 1926-66 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-

The Beginnings, Goddard and Oberth, 1926-45

FIG. 1.3 Robert Goddard (left) with his principal technical assistants (left to right: Nils Ljungquist, machinist; Albert Kisk, brother-in-law and machinist; and Charles Mansur, welder) in 1940 at Goddard’s shop in New Mexico. Shown is a rocket without its casing, with (right to left) the two propellant tanks and the extensive plumbing, including turbopumps to inject the propellants into the combustion chamber, where they ignite and create thrust by exhausting through the expansion nozzle (far left). (Photo courtesy of NASA)

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

Подпись: 13 German and U.S. Missiles and Rockets, 1926-66 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.

Propulsion for the Saturn Upper Stages

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


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.

Подпись: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


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

Подпись: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.

Подпись: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 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

Подпись: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

Подпись: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

Подпись: 201 Propulsion with Liquid Hydrogen and Oxygen, 1954-91 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


Подпись: 49 U.S. Space-Launch Vehicles, 1958-91 One major area of difference between missile and launch-vehicle de­velopment lay in the requirement for special safeguards on launch vehicles that propelled humans into space. Except for Juno I and Vanguard, which were short-lived, among the first U. S. space-launch vehicles were the Redstones and Atlases used in Project Mercury and the Atlases and Titan IIs used in Project Gemini to prepare for the Apollo Moon Program. Both Projects Mercury and Gemini re­quired a process called “man-rating" (at a time before there were women serving as astronauts). This process resulted in adaptations of the Redstone, Atlas, and Titan II missiles to make them safer for the human beings carried in Mercury and Gemini capsules.

Man-rating was but one of the ways missiles had to be modi­fied for use as launch vehicles, but the practice carried over to later launch vehicles initially designed as such (rather than as missiles). For Mercury-Redstone, Wernher von Braun’s Development Opera-

A Mercury – Redstone launching Freedom 7 with Astronaut Alan Shepard onboard,

Подпись: FIG. 2.1"Man-Rating&quotMay 5, 1961, from Pad 5 at Cape Canaveral. (Photo courtesy of NASA)

tions Division of the Army Ballistic Missile Agency was respon – 50 sible for the process. Von Braun established a Mercury-Redstone Chapter 2 Project Office to aid in redesigning the Jupiter C version of the Red­stone to satisfy the requirements of the Mercury project. To direct the effort, he chose Joachim P. Kuettner, a flight engineer and test pilot who had worked for Messerschmitt during the Nazi period in Germany.2

Kuettner’s group recognized that the Redstone missile could not satisfy the mission requirements for Project Mercury. These ne­cessitated sufficient performance and reliability to launch a two – ton payload with an astronaut aboard into a flight path reaching

an apogee of 100 nautical miles (115 statute miles). The Jupiter C, with its elongated propellant tanks and a lighter structure, had the required performance but not the safety features necessary for hu­man flight. To add these, Kuettner’s group reverted from the toxic hydyne to alcohol as a fuel. Other changes included an automatic, in-flight abort system with an escape rocket and parachutes to carry the astronauts to a safe landing. To ferret out potential sources of failure, Chrysler, as prime contractor, instituted a special test pro­gram to promote greater reliability. The overall process proved suc­cessful, resulting in the two flights of Alan Shepard and Virgil (Gus) Grissom in May and July 1961.3

Thereafter, Project Mercury switched to Atlas D missiles to pro­pel the astronauts and their capsules into orbit. For this function, the missile required strengthening in its upper section to handle the greater loads the capsule created. Following an explosion on Mer­cury-Atlas (MA) 1 (July 29, 1960), whose cause investigators could not determine, engineers developed an improved structure linking the booster and capsule, resulting in a successful flight of MA-2 on February 21, 1961.4 MA-2 also tested the Atlas abort sensing and implementation system (ASIS) and “escape tower" that were key features in man-rating the Atlas. Besides these two features, there had to be numerous other modifications to convert Atlas to its Mer­cury-Atlas configuration. For example, the Mercury capsule’s sepa­ration rockets potentially could damage the thin “steel-balloon" skin on the liquid-oxygen dome of the Atlas, so General Dynamics (formerly Convair) engineers had to add a fiberglass layer covering the dome. This and other changes, plus increased quality control, caused the Mercury-Atlas launch vehicle to cost 40 percent more than the Atlas missile. After a failure on MA-3 (due to guidance/ control problems), Atlas launch vehicles placed John Glenn, Scott Carpenter, Walter Schirra, and Gordon Cooper in orbit between Feb­ruary 20, 1962, and May 15, 1963.5

Подпись: 51 U.S. Space-Launch Vehicles, 1958-91 For Titan II-Gemini, there were major problems with longitudi­nal oscillations in the engines, known as pogo (from their resem­blance to the gyrations of the then-popular plaything, the pogo stick). These never occurred in flight but appeared in a severe form during static testing of second-stage engines. Surges in the oxidizer feed lines were causing the problem, which Martin engineers and others solved with suppression mechanisms. There was also the is­sue of combustion instability that occurred on only 2 percent of the ground tests of second-stage engines. But for man-rating, even this was too high. Aerojet (the Titan engine contractor) solved the problem with a new injector.6

Подпись: FIG. 2.2 Launch of a Mercury- Atlas vehicle from Cape Canaveral on February 20, 1962. (Photo courtesy of NASA)

For other aspects of man-rating Titan II for Gemini, procedures developed for Project Mercury offered a strong influence, especially 52 as many NASA and Aerospace Corporation engineers who had Chapter 2 worked on Mercury also worked on Gemini. Gemini engineers also benefited from Titan II test launches. As George E. Mueller, NASA’s associate administrator for manned spaceflight from 1963 to 1969, stated in February 1964, the 28 launches of Titan II missiles to that date “provide[d] invaluable launch operations experience and ac­tual space flight test data directly applicable to the Gemini launch vehicle which would [have] be[en] unobtainable otherwise,"7 one example of the symbiotic (though not homogeneous) relationship between missiles and launch vehicles.

FIG. 2.3

Gemini-Titan 12 launch on November 11, 1966, showing the exhaust plume from the engines on the Titan II launch vehicle. (Photo courtesy of NASA)



Подпись: 53 U.S. Space-Launch Vehicles, 1958-91 In addition to a malfunction detection system, features added to the Titan II missile for astronaut safety included redundant com­ponents of the electrical systems. To help compensate for all the weight from the additional components, engineers also deleted ver­nier and retro-rockets, which were not necessary for the Gemini mission.8 From March 23, 1965, to November 11, 1966, Gemini 3 through 12 all carried two astronauts on the Gemini spacecraft. These missions had their problems as well as their triumphs. But with them, the United States finally assumed the lead in the space race with its cold-war rival, the Soviet Union. Despite lots of prob­lems, Gemini had prepared the way for the Apollo Moon landings and achieved its essential objectives.9

Following its successful Gemini missions, the Titan II did not serve again as a space-launch vehicle until the mid-1980s after

it was taken out of service as a missile. Meanwhile, subsequent launch vehicles that required [hu]man-rating, notably the Saturn launch vehicles and the shuttle, included equipment for accommo­dating humans in their original designs, capitalizing on the experi­ences with Mercury-Redstone, Mercury-Atlas, and Gemini-Titan, which NASA passed on to the subsequent programs.

Minuteman Propulsion

Development of the propellant for Minuteman began at Wright – Patterson AFB and continued after Lt. Col. Edward N. Hall trans­ferred to the Western Development Division as chief for propulsion development in the liquid-propellant Atlas, Titan, and Thor pro­grams. In December 1954, he invited major manufacturers in the solid-propellant industry (Aerojet, Thiokol, Atlantic Research, Phil-

lips Petroleum Company, Grand Central, and Hercules) to discuss prospects for solids. The result apparently was the Air Force Large [Solid] Rocket Feasibility Program (AFLRP), which involved a com­petition starting in September 1955 with specific companies looking at different technologies. It appears that the propellant for Polaris benefited from Aerojet’s participation in this air force program.51

Redstone Propulsion

When the von Braun group, relocated to Redstone Arsenal, began de­veloping the Redstone missile, it chose North American Aviation’s XLR43-NA-1 liquid-propellant rocket engine, developed for the

air force’s Navaho missile, as the basis for the Redstone propul­sion unit. A letter contract with NAA on March 27, 1951, provided 120 days of research and development to make that engine comply with the ordnance corps’ specifications and to deliver a mockup and two prototypes of the engine (then to be designated NAA 75-110, referring to 75,000 pounds of thrust operating for 110 seconds). Sup­plemental contracts in 1952 and 1953 increased the number of en­gines to be delivered and called for their improvement. These con­tracts included 19 engines, with subsequent powerplants purchased by the prime contractor, Chrysler, through subcontracts.25

Подпись: 117 Propulsion with Alcohol and Kerosene Fuels, 1932-72 The story of how North American Aviation had initially devel­oped the XLR43-NA-1 illustrates much about the ways launch – vehicle technology developed in the United States. NAA came into existence in 1928 as a holding company for a variety of avia­tion-related firms. It suffered during the Depression after 1929, and General Motors acquired it in 1934, hiring James H. Kindelberger, nicknamed “Dutch," as its president—a pilot in World War I and an engineer who had worked for Donald Douglas before moving to NAA. Described as “a hard-driving bear of a man with a gruff, earthy sense of humor—mostly scatological—[who] ran the kind of flexible operation that smart people loved to work for," he re­organized NAA into a manufacturing firm that built thousands of P-51 Mustangs for the army air forces in World War II, the B-25 Mitchell bomber, and the T-6 Texas trainer. With the more cautious but also visionary John Leland Atwood as his chief engineer, Dutch made NAA one of the principal manufacturers of military aircraft during the conflict, its workforce rising to 90,000 at the height of wartime production before it fell to 5,000 after the end of the war. Atwood became president in 1948, when Dutch rose to chairman of the board and General Motors sold its share of the company. The two managers continued to service the military aircraft market in the much less lucrative postwar climate, when many competitors shifted to commercial airliners.26

Despite the drop in business, NAA had money from wartime production, and Kindelberger hired a top-notch individual to head a research laboratory filled with quality engineers in the fields of elec­tronics, automatic control with gyroscopes, jet propulsion, and mis­siles. He selected William Bollay, a former von Karman student. Fol­lowing receipt of his Ph. D. in aeronautical engineering at Caltech, Bollay had joined the navy in 1941 and been assigned to Annapolis where the Bureau of Aeronautics (BuAer) was working on experimen­tal engines, including the JATOs Robert Truax was developing. At war’s end, Bollay was chief of the Power Plant Development Branch

for BuAer. As such, he was responsible for turbojet engines, and at the time, NAA was developing the FJ-1 Fury, destined to become one of the navy’s first jet fighters. Bollay came to work for NAA during the fall of 1945 in a building near the Los Angeles airport, where he would create what became the Aerophysics Laboratory.27

On October 31, 1945, the army air forces’ Air Technical Service Command released an invitation for leading aircraft firms to bid on studies of guided missiles. NAA proposed a surface-to-surface rocket with a range of 175-500 miles that it designated Navaho (for North American vehicle alcohol [plus] hydrogen peroxide and oxygen). The proposal resulted in a contract on March 29, 1946, for MX-770, the designation of the experimental missile. Other con­tracts for the missile followed. The Navaho ultimately evolved into a complicated project before its cancellation in 1958. It included a rocket booster and ramjet engines with a lot of legacies passed on to aerospace technology, but for the Redstone, only the rocket engine that evolved to become NAA 75-10 is relevant.28

NAA did not originally intend to manufacture the engine. As an early employee of the firm recalled, the company was “forced into the engine business—we had the prime contract for Navaho and couldn’t find a subcontractor who would tackle the engine for it, so we decided to build it for ourselves." NAA’s plans for developing the engine began with the German V-2 as a model but soon led to “an en­tirely new design rated at 75,000 pounds thrust" (as compared with about 56,000 for the V-2). In the spring of 1946, Bollay and his as­sociates had visited Fort Bliss to conduct numerous interviews with many of the Germans who had worked on the V-2, including von Braun, Walter Riedel, and Konrad Dannenberg. By the middle of June 118 1946, Bollay’s team began redesigning the V-2 engine with the aid of

Chapter 3 drawings and other documents obtained from the Peenemunde files. In September, the firm secured the loan of a complete V-2.29

The NAA engineers also conferred with JPL, GE, Bell Aircraft Corp., the National Advisory Committee for Aeronautics’ labora­tory in Cleveland (later Lewis Research Center), the Naval Ord­nance Test Station at Inyokern, and Aerojet about various aspects of rocket technology.30 Thus, the heritage of the Redstone engine went well beyond what NAA had learned from the V-2.

By October 1947, the Astrophysics Laboratory had grown to more than 500 people. This necessitated a move to a plant in nearby Downey in July 1948. By the following fall, the engineers had taken apart and reconditioned the V-2 engine, examining all of its parts carefully. The team had also built the XLR41-NA-1, a rocket engine like the V-2 but using U. S. manufacturing techniques and design

standards, some improved materials, and various replacements of small components. Then, by early 1950, the team had redesigned the engine to a cylindrical shape, replacing the spherical contour of the V-2, which produced efficient propulsion but was hard to form and weld. Bollay’s people kept the propellants for the V-2 (75 per­cent alcohol and liquid oxygen). But in place of the 18-pot design of the V-2, which had avoided combustion instability, NAA engineers developed two types of flat injectors—a doublet version in which the alcohol and liquid oxygen impinged on one another to achieve mixing, and a triplet, wherein two streams of alcohol met one of liquid oxygen. They tested subscale versions of these injectors in small engines fired in the parking lot. Their methodology was purely empirical, showing the undeveloped state of analytical capabilities in this period. They, too, encountered combustion instability. But they found that the triplet type of injector provided slightly higher performance due to improved mixing of the propellants.31

Meanwhile, NAA searched for a place where it could test larger engines. It found one in the Santa Susana Mountains northwest of Los Angeles in Ventura County, California. The firm obtained a permit in November 1947 for engine testing there. It leased the land and built rocket-testing facilities in the rugged area where Tom Mix had starred in western movies, using company funds for about a third of the initial costs and air force funding for the rest. By early 1950, the first full-scale static test on XLR43-NA-1 took place.32

Подпись: 119 Propulsion with Alcohol and Kerosene Fuels, 1932-72 Full-scale engine tests with the triplet injector revealed severe combustion instability, so engineers reverted to the doublet injec­tor that partly relieved the problem. Although the reduced combus­tion instability came at a cost of lower performance, the XLR43- NA-1 still outperformed the V-2, enabling use of the simpler and less bulky cylindrical combustion chamber that looked a bit like a farmer’s milk container with a bottom that flared out at the nozzle. The engine delivered 75,000 pounds of thrust at a specific impulse 8 percent better than that of the V-2. Further enhancing perfor­mance was a 40 percent reduction in weight. The new engine re­tained the double-wall construction of the V-2 with regenerative and film cooling. Tinkering with the placement of the igniter plus injection of liquid oxygen ahead of the fuel solved the problem with combustion instability. The engine used hydrogen peroxide – powered turbopumps like those on the V-2 except that they were smaller and lighter. It also provided higher combustion pressures.

Like the V-2, the XLR43-NA-1 began ignition with a preliminary stage in which the propellants flowed at only some 10 to 15 percent of full combustion rates. If observation suggested that the engine

was burning satisfactorily, technicians allowed it to transition to so-called main stage combustion. To enable the engineers to ob­serve ignition and early combustion, von Braun, who was working with the NAA engineers by this time, suggested rolling a small, surplus army tank to the rear of the nozzle. By looking at the com­bustion process from inside the tank, engineers could see what was happening while protected from the hot exhaust, enabling them to reduce problems with rough starts by changing sequencing and im­proving purges of the system in a trial-and-error process. Through such methods, the XLR43-NA-1 became the basis for the Redstone missile’s NAA 75-110.33

Having supervised the development of this engine and the ex­pansion of the Aerophysics Laboratory to about 2,400 people on staff, Bollay left North American in 1951 to set up his own com­pany, which built army battlefield missiles. In 1949, he had hired Samuel K. Hoffman, who had served as a design engineer for Fair­child Aircraft Company, Lycoming Manufacturing Company, and the Allison Division of General Motors. He then worked his way up from project engineer with the Lycoming Division of the Aviation Corporation to become its chief engineer, responsible for the design, development, and production of aircraft engines. In 1945 he became a professor of aeronautical engineering at his alma mater, Penn State University, the position he left in 1949 to head the Propulsion Sec­tion of what became NAA’s Aerophysics Laboratory.

As Hoffman later recalled, Bollay had hired him for his practi­cal experience building engines, something the many brilliant but young engineers working in the laboratory did not possess. Hoffman succeeded Bollay in 1951. Meanwhile, he and Bollay had overseen 120 the development of a significantly new rocket engine. Although it Chapter 3 had used the V-2 as a starting point and bore considerable resem­blance to the cylindrical engine developed at Peenemunde before the end of the war, it had advanced substantially beyond the Ger­man technology and provided greater thrust with a smaller weight penalty. Also, it marked the beginnings of another rocket-engine manufacturing organization that went on to become the Rocket- dyne Division of NAA in 1955, destined to become the foremost producer of rocket engines in the country.34

Development of the NAA 75-110 engine for the Redstone mis­sile did not stop in 1951. Improvements continued through seven engine types, designated A-1 through A-7. Each of these engines had fundamentally the same operational features, designed for identical performance parameters. The engines were interchangeable, requir­ing only minor modifications in their tubing for them to be installed

Подпись: TABLE 3.1. Comparison of Components in Pneumatic Control System for the Redstone A-1 and A-7 Engines Components A-1 Engine A-7 Engine Regulators 4 1 Relief valves 2 1 Solenoid valves 12 4 Pressure switches 6 2 Check valves 3 0 Test connections 4 1 Pneumatic filter 0 1 Total 31 10

in the Redstone missile. All of them except A-5 flew on Redstone tests between August 20, 1953, and November 5, 1958, with A-1 be­ing the prototype and A-2, for example, having an inducer added to the liquid-oxygen pump to prevent cavitation (bubbles forming in the oxidizer, causing lower performance of the turbopump and even damage to hardware as the bubbles imploded).35

Подпись: 121 Propulsion with Alcohol and Kerosene Fuels, 1932-72 During the course of these improvements, the Chrysler Corpo­ration had become the prime contractor for the Redstone missile, receiving a letter contract in October 1952 and a more formal one on June 19, 1953. Thereafter, it and NAA had undertaken a product – improvement program to increase engine reliability and reproduc­ibility. A comparison of the numbers of components in the pneu­matic control system for the A-1 and A-7 engines, used respectively in 1953 and 1958, illustrates the results (see table 3.1).36

Obviously, the fewer components needed to operate a complex system like the engine for a large missile, the fewer things there are that can go wrong in its launching and flight. Thus, this threefold reduction in components on a single system for the engine must have contributed significantly to the reliability of its operation. This was especially true since Rocketdyne engineers (as they be­came after 1955) tested each new component design both in the lab­oratory and in static firings before qualifying it for production. They also simulated operating conditions at extreme temperatures, levels of humidity, dust, and the like, because the Redstone was scheduled for deployment and use by the army in the field. Static engine tests showed reliability higher than 96 percent for the engines.

This was a remarkably high figure, considering that Rocketdyne purchased about half of Redstone engine components (or parts thereof) from outside suppliers; but the parts had to be built to a higher standard than those used in conventional aircraft. The rea­son was that the stresses of an operating rocket engine were greater than those for an airplane. All welds for stressed components had to undergo radiographic inspection to ensure reliability. The army then required a minimum of four static engine tests to prove each new model worked satisfactorily before the service would accept the system. Two of these tests had to last for 15 seconds each, and a third was for the full rated duration. This, presumably, was 110 sec­onds, but according to Chrysler’s publication on the Redstone, the engine ultimately produced 78,000 pounds of thrust for a duration of 117 seconds.37

Kummersdorf, Peenemunde, and the V-2

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

Подпись: 15 German and U.S. Missiles and Rockets, 1926-66 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.


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


FLIGHT TESTINGThe 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

Подпись: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

The Thor and Delta Family of Launch Vehicles, 1958-90

The Thor missiles did not remain in operational use very long, but even before the air force retired them in 1963, it had begun to use the Thor’s airframe and propulsive elements (including its vernier engines) as the first stages of various launch vehicles. With a series of upratings and modifications, the Thor remained in use with such upper stages as the Able, Able-Star, Agena, Burner I, Burner II, and Burner IIA until 1980. In addition, NASA quickly chose the Thor as the first stage of what became its Thor-Delta (later, just Delta) launch-vehicle family, which has had an even longer history than the air force’s Thor series. The Delta launch vehicles initially drew upon Vanguard upper stages, as did the Thor-Able used by the air force.10

Throughout its history, the Delta evolved by uprating existing components or adopting newer ones that had proven themselves. It used a low-risk strategy to improve its payload capacity through the Delta II at the end of the period covered by this history. But it did not stop there, evolving through a Delta III, first launched (unsuc­cessfully) in 1998, and a Delta IV that finally had its successful first launch on November 20, 2002. (To be sure, the Delta IV used an entirely new first stage, making it in some senses a new launch ve­hicle, but the design emphasized reliability and low cost, hallmarks of the Delta program from the beginning.) The unsuccessful first (and second) launch(es) of the Delta III and numerous delays in the launch of Delta IV because of both software and hardware problems 54 suggested, however, that the design of new launch vehicles was still Chapter 2 not something engineers had “down to a science," even in the 21st century.11

For the Able upper stage, the air force and its contractors used many features of the Vanguard second stage but added a control compartment, skirts and structural elements to mate it with the Thor, a tank venting and pressurization safety system, new electri­cal components, and a roll-control system. Used for reentry test­ing, the first Thor-Able failed because of a faulty turbopump in the Thor, but the second launch on July 9, 1958, was successful.12

FIG. 2.4

A Thor-Able launch vehicle with the Pioneer 1 spacecraft as its payload. (Photo courtesy of NASA)


The Thor and Delta Family of Launch Vehicles, 1958-90

Succeeding versions of Thor-Able modified both second and third stages of Vanguard. The final Thor-Able launch on April 1, 1960, placed the Tiros 1 meteorological satellite in orbit. In the 16 Thor-Able launches, all of the stages worked satisfactorily on 10 of the missions, whereas at least one stage failed or was only partly successful on 6 flights. Although this was only a 62.5 percent suc­cess rate, it was sufficiently good for this early period that the air force could refer to Thor-Able as “an extremely capable and reliable vehicle combination."13

Подпись: 55 U.S. Space-Launch Vehicles, 1958-91 Long before the final launch of the Thor-Able, nevertheless, the Department of Defense’s Advanced Research Projects Agency had

issued an order on July 1, 1959, calling for the development of the Able-Star upper stage, derived from the Able but possessing two and one half times its total impulse plus the capability to shut down its propulsion in space, coast, and restart. This ability would permit a more precise selection of the orbit for a satellite than was possible before. Once Able-Star became operational in January 1960, it ef­fectively replaced the Able as an upper stage.14

The Able-Star engine was a derivative of the several Able engines except that it had the added restart capability plus the capacity to provide attitude control during coasting periods and to burn longer than the earlier engines. Following a rapid but not unproblematic development, an Able-Star upper stage on a Thor booster launched the Transit 1B navigation satellite on April 13, 1960, marking the first programmed restart of a rocket engine in flight. Although the coast attitude-control system worked, a malfunction in the Able – Star ground guidance system resulted in a still-useful elliptical rather than a circular orbit. Sloshing in stage-two propellant tanks for Transit 2A—launched on June 22, 1960—again produced an el­liptical orbit because it caused roll forces that the guidance/control system could not overcome. Again the orbit was useful. Following placement of anti-slosh baffles in both Able-Star propellant tanks, the Thor booster failed on the attempted launch of Transit 3A, No­vember 30, 1960. Then on February 21, 1961, a Thor/Able-Star failed to place Transit 3B in a usable orbit because a part malfunctioned in the programmer before it could signal the restart of stage two from its coasting orbit. Substantially the same launch vehicle as on Transit 3A successfully launched Transit 4A into a nearly circular orbit on June 28, 1961.15 Through August 13, 1965, including the launches just discussed, the Thor/Able-Star completed a total of 20 missions with 5 failures, for a success rate of 75 percent. Quite successful for an early launch vehicle, the Thor/Able-Star also marked a step forward in satellite-launching capability.16

Even before the first launch of Thor/Able-Star, the air force had 56 begun using an Agena upper stage with the Thor, and this combi – Chapter 2 nation became a preferred choice for a great many often-classified missions, including those to place a family of reconnaissance satel­lites in orbit under what began as the WS-117L program. Initially, the Agena upper stages flew on basic Thors, but in three versions from A to D (without a C), the Agena also operated with uprated Thors, Atlases, and Titan IIIs to orbit a great many military and NASA spacecraft until 1987.17

The air force began developing the Agena in July 1956. On Oc­tober 29, 1956, that service selected Lockheed Missile Systems

Division as the prime contractor for both the WS-117L reconnais­sance satellite system and an associated upper stage that became the Agena. The engine for the Lockheed upper stage was a modified version of the Hustler propulsion unit (model 117) that Bell Aero­space had developed for the B-58 bomber’s air-to-surface missile, designated the Powered Disposable Bomb Pod. The air force can­celed the missile, but Lockheed contracted with Bell in the fall of 1957 to develop the engine for Agena.18

One change from the Hustler engine was the addition of gim – balling. Another was a nozzle closure to ensure that the Agena started in space after cutoff of the first-stage engine. The Agena stage with this engine, known as the Bell 8001, flew only once, on February 28, 1959, for the launch of Discoverer 1 by a Thor-Agena A. (Discoverer was the name publicly released for the secret Corona reconnaissance satellites, which had separated from the WS-117L program by this time.) Accounts differ as to the outcome of this first launch into a polar orbit from Vandenberg AFB, California—some claiming the launch itself was successful, and others that it was not.19

The Agena nevertheless had an extensive career as an upper stage. The Agena A operated successfully on 78 percent of its 14 launches by September 13, 1960 (all by Thors; all but one with a new Bell model 8048 engine for the Agena), with 3 failures. A more capable Thor-Agena B appears to have had 39 successful performances on 48 launches from October 26, 1960, to May 15, 1966, an 81 percent success rate, mostly launching Corona satellites. With a thrust- augmented Thor, the Agena B could launch much heavier satellites, added thrust coming from solid-propellant strap-on boosters.20

Подпись: 57 U.S. Space-Launch Vehicles, 1958-91 Meanwhile, in the fall of 1959 Bell began designing the engine for the Agena D, which became the standard Agena propulsion unit. From June 28, 1962, to May 25, 1972, a large number of Thor-Agena D launches occurred, but because of the classified nature of many of the payloads, a reliable and precise tally is not available. During this period, the basic booster changed from the thrust-augmented to the long-tank, thrust-augmented Thor (called Thorad) with increased burning time and improved strap-on boosters.21

Another series of upper stages used with the Thor first stage in­cluded Burner I, Burner II, and Burner IIA. Burner I actually bore little relation to Burners II and IIA. Information about it is sparse, but sources refer to it as the Altair, a derivative of the Vanguard third stage developed by Hercules Powder Company at the Allegany Ballistics Laboratory. The first launch of the Thor-Burner I occurred on January 18, 1965, with the last one taking place on March 30, 1966. There apparently were only four such launches, all from Van-

An Agena upper stage, used also for many satellite launches, serving here as the Gemini 8 target vehicle for docking. (Photo courtesy of NASA)

Подпись: FIG. 2.5The Thor and Delta Family of Launch Vehicles, 1958-90denberg AFB into sun-synchronous orbits, one of them a mission failure. The spacecraft were classified at the time but appear to have been Block 4A Defense Satellite Applications Program weather sat­ellites used to inform the U. S. military of weather conditions for launching reconnaissance satellites and other defense purposes, such as mission planning during the conflict in Vietnam. In 1973, the program became the Defense Meteorological Satellite Program (DMSP) and was no longer classified.22

58 Burner I was little used because of the development of Burner Chapter 2 II. Conceiving a need for a guided upper stage that would be low in cost and usable with more than one first-stage vehicle, on Septem­ber 2, 1961, the air force’s Space Systems Division (SSD) awarded study contracts to the Boeing Company and Ling-Temco-Vought, Inc., pursuant to development of what became Burner II. As a result of its initial work, Boeing won a fixed-price contract on April 1, 1965, to provide one ground-test and three flight versions of the new upper stage. By September 15, Maj. Gen. Ben I. Funk, commander of SSD, could announce the development of the new stage, which

became the smallest maneuverable upper-stage vehicle in the air force inventory.23

The primary propulsion for Burner II came from the Thiokol TE- M-364-2 (Star 37B) motor, a spherical design promoted by NASA engineer Guy Thibodeaux. Between September 15, 1966, and Feb­ruary 17, 1971, Thor-Burner II vehicles launched four Block 4A, three Block 4B, and three Block 5A Defense Satellite Applications Program weather satellites from the Western Test Range. During this same general period, the Thor-Burner II also launched scientific satellites as part of the Department of Defense’s Space Experiments Support Program managed by SSD.24

The Block 5B versions of the Defense Satellite Applications Pro­gram weather satellites were about twice as heavy as the 5A ver­sions, necessitating increased thrust for Burner II. So the air force’s Space and Missile Systems Organization (created July 1, 1967, to bring the SSD and its sister Ballistics Systems Division into a single organization headquartered in Los Angeles at the former SSD loca­tion) contracted with Boeing for an uprated Burner II that became Burner IIA. Boeing did the uprating with a minimum of modifi­cations by adding a Thiokol TE-M-442-1 motor to form a second upper stage. With the Burner IIA, a Thor first stage launched five Block 5B and two Block 5C meteorological satellites (in what be­came the DMSP) from October 14, 1971, to May 24, 1975. A final Thor-Burner IIA launch on February 18, 1976, failed because the Thor prematurely ceased firing. This last use of the Burner IIA did not spell the end of the DMSP program, however, because a Thor coupled with a Thiokol TE-M-364-15 (Star 37S) motor that had a titanium case (rather than the steel used on the Star 37B) launched 4 improved Block 5D weather satellites between Sept ember 11, 1976, and June 6, 1979. Then Atlas Es and Titan IIs launched 10 more DMSP satellites by 1999.25

Подпись: 59 U.S. Space-Launch Vehicles, 1958-91 A final major Thor launch vehicle was the Thor-Delta. Whereas the other Thor-based launch vehicles were primarily air force as­sets sometimes used by NASA, Delta was a NASA-developed space – launch vehicle used on occasion by the air force until near the end of the period covered by this book, when the air force began to make extensive use of Delta IIs. Since it was conceived by NASA in 1959 as an interim vehicle to lift medium payloads by using existing technology, modified only as needed for specific missions, Delta has enjoyed a remarkably long career, attesting to its success.26

The initial idea for Thor-Delta apparently came from Milton Rosen. He was working at NASA Headquarters in the Office of Space Flight Development, headed by Abe Silverstein. His imme-

diate supervisor was Abraham Hyatt, who had become the assis­tant director for propulsion following a decade of work at the navy’s Bureau of Aeronautics. At Silverstein’s behest, Rosen worked with Douglas Aircraft Company to develop the vehicle. Using compo­nents already proven in flight, NASA and Douglas eliminated the need for developmental flights. Their contract set a very ambitious goal (for 1959) of an initial 50 percent reliability with a final rate of 90 percent.27

An important asset in Delta’s development consisted of the per­sonnel from the Vanguard program, including Rosen, who brought their experience to decision-making positions at the new Goddard Space Flight Center within NASA, as well as at NASA Headquarters. At Goddard, William R. Schindler, who had worked on Vanguard, headed a small technical group that provided direction and tech­nical monitoring for the Delta program, which initially borrowed technology from the Vanguard, Thor-Able, and Titan programs. On November 24, 1962, NASA converted this technical direction to formal project management for Delta.28

On May 13, 1960, an attempt to launch the spherical, passive reflector satellite Echo with the first Thor-Delta failed when the third-stage propellants did not ignite because a small chunk of sol­der in a transistor broke loose in flight and shorted out a semicon­ductor that had passed all of its qualification tests. A similar but less costly problem with another transistor on the third Delta launch led NASA to change its specifications and testing of such components. Meanwhile, on August 12, 1960, the second Delta launch success­fully placed Echo 1 into orbit. And the remainder of the original 12 Deltas all had successful launches of a variety of payloads from the Tiros 2 through 6 weather satellites to the Telstar 1 communica­tions satellite, the first commercial spacecraft launched by NASA (on July 10, 1962, the last of the original 12 launches by a Delta).29

From this beginning, the Delta went through a long and compli­cated series of modifications and upgrades. The initial Delta could 60 launch 100 pounds of payload to geostationary (also called geosyn – Chapter 2 chronous) transfer orbit. Starting in 1962, Delta evolved through a series of models with designations such as A, B, C, D, E, J, L, M, M-6, N, 900, 904, 2914, 3914, 3910/PAM (for Payload Assist Mod­ule), 3920/PAM, 6925 (Delta II), and 7925 (also a Delta II), the last of them introduced in 1990. The payload capabilities of these versions of the vehicle increased, at first gradually and then more rapidly, so that the 3914 introduced in 1975 could lift 2,100 pounds to geo­stationary transfer orbit and the 7925 could lift 4,010 pounds (40.1 times the original capability).30

FIG. 2.6

The Thor and Delta Family of Launch Vehicles, 1958-90Подпись: 61 U.S. Space-Launch Vehicles, 1958-91 Подпись: To achieve this enormous growth in payload from 1960 to 1990, the Delta program augmented the capabilities of the booster and upper stages, lengthened and enlarged the tanks of the first two liquid-propellant stages, enlarged and upgraded third-stage motors, improved guidance systems, and introduced increasingly large and numerous strap-on solids to provide so-called zero-stage boost. During this period, the program generally continued to follow Rosen's initial approach of introducing only low-risk modifications or ones involving proven systems. This enabled, on average, a launch every 60 days with a reliability over the 30 years of 94 percent (189 successes out of 201 attempts, the last one through the end of 1990 occurring on November 26, 1990).31 Delta launches and improvements continued beyond this period. Because the Thor and Delta rockets were not so much innovators
A Delta 1910 vehicle launching Orbiting Solar Observatory 8 on June 21, 1975, showing the Castor II strap-on boosters at the base of the vehicle to add to the thrust. (Photo courtesy of NASA)

as borrowers of new technology from other programs, they experi­enced fewer birth pangs than other missiles and rockets, showing the value of shared information. They nevertheless did experience some unexpected problems that required redesign. But by (mostly) using components already tested and proven, the Delta achieved a high reliability that made it an enduring member of the launch – vehicle family. From an interim launch vehicle in 1959, it became one of the few that lasted into the 21st century, a distinction shared with the Atlas family.32


Подпись:Despite this preparatory work for Minuteman, the missile did not begin formal development until the air force secured final DoD ap­proval in February 1958. Meanwhile, civilian engineers employed by the air force at the Non-Rotating Engine Branch at Wright-Patterson AFB had continued efforts to develop large, solid-propellant motors. Perhaps from sources at Aerojet, they learned about adding large quantities of aluminum to a solid propellant to increase its perfor­mance. They combined this with other information they had been gathering on solid propellants while Hall was still with the branch. Bill Fagan from the branch carried this information to Hall at the Ballistic Missile Division (as the WDD had become).52

The specifics of Minuteman technology continued to evolve, but its basic concept took advantage of the reduced complexity of solids over liquids. This cut the number of people needed to launch it. Each missile could be remotely launched from a control center us­ing communications cables, without a crew of missileers in atten­dance. The air force initially considered using mobile Minutemen but ultimately decided to launch them from silos. Because solids were more compact than liquids, Minuteman IA (by one account) was only 53.8 feet long to Atlas’s 82.5, Titan I’s 98, and Titan II’s 108 feet in length. Its diameter was slightly over half that of the other three missiles (5.5 feet to the others’ 10), and its weight was only 65,000 pounds to Atlas’s 267,136, Titan I’s 220,000, and Titan II’s 330,000 pounds. All of this made the costs of silos much lower and substantially reduced the thrust needed to launch the missile. Consequently, the initial costs of Minuteman were a fifth and the annual maintenance costs a tenth those of Titan I. Moreover, a crew of two could launch 10 Minutemen, whereas it took six people to launch each Titan I—a 30-fold advantage in favor of Minuteman.53

Minuteman development used multiple approaches to arrive at individual technologies, as had been true with Atlas. Ballistic Mis­sile Division (BMD) contracted separately with Aerojet and Thiokol to work on all three stages of the missile. A later contract with Her­cules assigned it to work on the third stage, too. As firms developed technologies, parallel development gave way to specific responsi-

bilities, with Thiokol building the first stage, Aerojet the second, and Hercules the third. Space Technology Laboratories retained its role in systems engineering and technical direction at BMD. These and other companies and organizations all sent representatives to frequent program-review meetings and quarterly gatherings of top officials. They examined progress and identified problems requiring solution. Then, the relevant organizations found solutions to keep the program on track.54

One major technical challenge involved materials for the nozzle throats and exit cones. The addition of aluminum to the propellant provided a high enough specific impulse to make Minuteman fea­sible, as it had done for Polaris, and its combustion produced alu­minum oxide particles that damped instabilities in the combustion 248 chamber. But the hot product flow degraded the nozzle throats and Chapter 6 other exposed structures. It seemed that it might not be possible to design a vectorable nozzle that would last the 60 seconds needed for the missile to reach its ballistic trajectory so it could arrive accu­rately on target. Solving this problem required “many months and many dollars. . . spent in a frustrating cycle of design, test, failure, redesign, retest, and failure." The Minuteman team used many dif­ferent grades and exotic compounds of graphite, which seemed the most capable material, but all of them experienced blowouts or per­formance-degrading erosion. One solution was a tungsten throat in­sert, a compromise in view of its high weight and cost. For the exit cones, the team tried Fiberite molded at high pressure and loaded with silica and graphite cloth. It provided better resistance to ero­sion than graphite but still experienced random failures.55

Another significant problem concerned the vectorable feature of the nozzles. Polaris had solved its steering problem with jetavators, but flight-control studies for Minuteman showed a need for the stage-one nozzles to vector the thrust eight degrees, more than Min – uteman engineers thought jetavators could deliver. Thiokol’s test of the motor in 1959—the largest solid-propellant powerplant yet built—resulted in the ejection of all four nozzles after 30 millisec­onds of firing, well before the full stage had ignited. Five successive explosions of the motors and their test stands occurred in October 1959, each having a different failure mode. BMD halted first-stage testing in January 1960. Discussion among BMD, STL, and Thiokol personnel revealed two problem areas, internal insulation and the nozzles themselves, as masking other potential problems.

There followed two concurrent programs of testing. Firings with battleship-steel cases tested movable nozzles, while partici­pants used flight-weight cases to test a single, fixed nozzle massive

enough to sustain a full-duration firing. Thiokol solved the problem with insulation by summer but did not resolve the nozzle issue un­til fall. This was close to the date set for all-up testing of the entire missile. However, General Phillips had ordered Thiokol to begin manufacturing the first stage except the nozzles, permitting instal­lation of the nozzles as soon as their problem was solved.56

Подпись:Among many other difficulties, a major concern was launching from a silo. The first successful launch of the missile occurred in early February 1961, which Phillips referred to as “December 63rd" since the planned date had been in December 1960. It and two suc­ceeding launches took place from a surface pad and were so suc­cessful that the team advanced the first silo launch to August 1961. The missile blew up in the silo, giving credence to critics in STL who had argued that firing a missile from a silo was impossible. As Phillips said, the missile “really came out of there like a Roman candle."57

Fortunately, team members recovered enough of the guidance system from the wreckage to find that the problem was not the silo launch itself but quality control. Solder tabs containing con­nections had vibrated together, causing all of the stages to ignite simultaneously. Knowing this, the team was able to prepare the fifth missile for flight with the problem solved by mid-November, when it had a successful flight from the silo. Previous testing of silo launches aided this quick recovery. In early 1958, BMD and STL engineers had arranged for development of underground silos. They divided the effort into three phases, with the first two using only subscale models of Minuteman. The third tested full-scale models. Most testing occurred at the air force’s rocket site on Edwards AFB in the remote Mojave Desert, but Boeing (contractor for missile as­sembly and test) tested subscale models in Seattle.

At the rocket site on Leuhman Ridge at Edwards, subscale silos investigated heat transfer and turbulence in some 56 tests by No­vember 1958. Meanwhile, Boeing modeled the pressures that the rocket’s exhaust gases imparted to the missile and silo. It also ex­amined acoustic effects of the noise levels generated by the rocket motor in the silo on delicate systems such as guidance. Armed with such data, engineers at Edwards began one-third-scale tests in Feb­ruary 1959. Full-scale tests initially used a mock-up made of steel plate with ballast to match the weight and shape of the actual mis­sile and only enough propellant (in the first-stage motor) to provide about three seconds of full thrust—enough to move the missile, on a tether, out of the silo and to check the effects of the thrust on the silo and missile. They configured the tether so that the missile

would not drop back on the silo and damage it. These tests ensured that the second silo launch at Cape Canaveral on November 17, 1961, was fully successful.58

The all-up testing on Minuteman was itself a significant inno­vation later used on other programs, including the Saturn launch vehicle for the Apollo program. It differed from the usual practice for liquid-propellant missiles—gradually testing a missile’s differ­ent capabilities over a series of ranges and tests (for first-stage or booster propulsion, for second-stage or sustainer-engine propulsion, for guidance/control, and so forth). For the first time, it tested all of the missile’s functions at once over the full operational range. Al­though the practice accorded well with general procedures at BMD, such as concurrency, its use came about in an odd way. According 250 to Otto Glasser, he was briefing Secretary of the Air Force James H.

Chapter 6 Douglas on Minuteman, with Gen. Curtis LeMay, vice chief of staff of the air force, sitting next to Douglas. Douglas insisted Glasser had moved the first flight of the missile to a year later than the original schedule. Glasser protested (to no avail) that this was not the case, and the only way he could conceive to cut a year out of the develop­ment process was all-up testing. “Boy, the Ramo-Wooldridge crowd came right out of the chair on that," Glasser said. They protested “a test program. . . with that sort of lack of attention to all normal, sensible standards." But all-up testing “worked all the way."59

Overcoming these and other problems, BMD delivered the first Minuteman I to the Strategic Air Command in October 1962, al­most exactly four years after the first contracts had been signed with contractors to begin the missile’s development. This was a year earlier than initially planned because of the speeded-up sched­ule. The missile in question was the A model of Minuteman I, later to be succeeded by a B model. The former was 53.7 feet long and consisted of three stages plus the reentry vehicle. Thiokol’s first stage included a new propellant binder developed by the company’s chemists from 1952 to 1954. Thiokol first tried a binder called poly­butadiene-acrylic acid, or PBAA, an elastomeric (rubberlike) copo­lymer of butadiene and acrylic acid that allowed higher concentra­tions of solid ingredients and greater fuel content than previous propellants. It had a higher hydrogen content than earlier Thiokol polysulfide polymers. With PBAA, a favorable reaction of oxygen with the aluminum generated significant amounts of hydrogen in the exhaust gases, reducing the average molecular weights of the combustion products (since hydrogen is the lightest of elements). This added to the performance, with Minuteman being the first rocket to use the new binder.60

But testing showed PBAA had a lower tear strength than poly­sulfide, so Thiokol added 10 percent acrylonitrile, creating poly­butadiene-acrylic acid-acrylonitrile (PBAN). The binder and curing agent constituted only 14 percent of the propellant, with ammonium perchlorate (oxidizer) and aluminum (fuel) the two other major in­gredients. The combination yielded a theoretical specific impulse of more than 260 lb-sec/lb, with the actual specific impulse at sea level at 70°F somewhat lower than 230.61

Подпись:For stage two of Minuteman I, Aerojet used the polyurethane binder employed in Polaris, with ammonium perchlorate as the oxidizer and aluminum powder the major fuel. It used two slightly different propellant grains, with a faster-burning inner grain and a slower-burning outer one. The combination resulted in a conversion of the four-point, star-shaped, internal-burning cavity to a cylindri­cal one as the propellant burned, avoiding slivers of propellant that did not burn. The propellant yielded a vacuum specific impulse of nearly 275 lbf-sec/lbm at temperatures ranging from 60°F to 80°F.62

For stage three, the Hercules Powder Company used a glass- filament-wound case instead of the steel employed on stages one and two, plus a very different propellant than for the first two stages. The third stage featured four phenolic-coated aluminum tubes for thrust termination and a grain consisting of two separate composi­tions. The one used for the largest percentage of the grain included the high-explosive HMX, combined with ammonium perchlorate, nitroglycerin, nitrocellulose, aluminum, a plasticizer, and a stabi­lizer. The second composition had the same basic ingredients mi­nus the HMX and formed a horizontal segment at the front of the motor. A hollow core ran from the back of the motor almost to the segment containing the non-HMX composition. It was roughly cone shaped before tapering off to a cylinder. This motor yielded a specific impulse of more than 275 lbf-sec/lbm at temperatures ranging from 60°F to 80°F. The four nozzles for stage three of Min – uteman I rotated in pairs up to four degrees in one plane to provide pitch, yaw, and roll control.63 Minuteman I, Wing I became opera­tional at Malmstrom AFB, Montana, in October 1962.64

It would be tedious to follow the evolution of Minuteman through all the improvements in its later versions, but some discussion of the major changes is appropriate. Wings II through V of Minuteman I (each located at a different base) featured several changes to increase the missile’s range. This had been shorter than initially planned be­cause of the acceleration of the Minuteman I schedule. The shorter range was not a problem at Malmstrom because it was so far north (hence closer to the Soviet Union), but range became a problem

starting with Wing II. Consequently, for it and subsequent missiles, more propellant was added to the aft dome of stage one, and the exit cone included contouring that made the nozzle more efficient. In stage two, the material for the motor case was changed from steel to titanium. Titanium is considerably lighter than steel but more expensive. Since each pound of reduced weight yielded an extra mile of range, use of titanium seemed worth the extra cost. The nozzles also were lighter. Overall, the reduction in weight totaled slightly less than 300 pounds despite an increase in propellant weight. The increase in propellant mass plus the decrease in weight yielded a range increase of 315 miles to a figure usually given as 6,300 nauti­cal miles. There were no significant changes to stage three.65

Подпись: 252 Chapter 6 MINUTEMAN II

For Minuteman II, the major improvements occurred in Aerojet’s stage two. There had been problems with cracking and ejection of graphite from the nozzles and aft closure of stage one. An air force reliability improvement program solved these difficulties. There had also been problems with insulation burning through in the aft dome area of stage three. Unspecified design changes inhibited the flow of hot gases in that region. Stage two, however, featured an entirely new rocket motor with a new propellant, a slightly greater length, a substantially larger diameter, and a single fixed nozzle that used a liquid-injection thrust-vector-control system for directional control.66

The new propellant was carboxy-terminated polybutadiene (CTPB), which propellant companies other than Aerojet had devel­oped. Some accounts attribute its development to Thiokol, which first made the propellant in the late 1950s and converted it into a useful propellant in the early 1960s. Initially, Thiokol chemists used an imine known as MAPO and an epoxide in curing the CTPB. It turned out that the phosphorous-nitrogen bond in the imine was susceptible to hydrolysis, causing degradation and softening of the propellant. According to Thiokol historian E. S. Sutton, “The post­curing problem was finally solved by the discovery that a small amount of chromium octoate (0.02%) could be used to catalyze the epoxide-carboxyl reaction and eliminate this change in properties with time." A history of Atlantic Research Corporation agrees that Thiokol produced the CTPB but attributes the solution of the curing problem to ARC, which is not incompatible with Sutton’s account. According to the ARC history, “ARC used a complex chromium compound, which would accelerate the polymer/epoxy reaction,

paving the way for an all epoxy cure system for CTPB polymer." The result was “an extremely stable binder system."67

Подпись:It frequently happens in the history of technology that innova­tions occur to different people at about the same time. This ap­pears to have been the case with CTPB, which Aerojet historians attribute to Phillips Petroleum and Rocketdyne without providing details. These two companies may have been the source for infor­mation about the CTPB that Aerojet used in Minuteman II, stage two. Like Thiokol, in any event, Aerojet proposed to use MAPO as a cross-linking agent. TRW historians state that their firm’s labo­ratory investigations revealed the hydrolysis problem. They state that “working with Aerojet’s research and development staff," they developed “a formulation that eliminated MAPO. . . ." The CTPB that resulted from what apparently was a multicompany develop­ment effort had better fuel values than previous propellants, good mechanical properties such as the long shelf life required for silo – based missiles, and a higher solids content than previous binders. The propellant consisted primarily of CTPB, ammonium perchlo­rate, and aluminum. It yielded a vacuum specific impulse more than 15 lbf-sec/lbm higher than the propellant used in stage two of Minuteman I, Wing II.68

Although CTPB marked a significant step forward in binder tech­nology, it was not as widely used as it might have been because of its higher cost compared with PBAN. Another factor was the emer­gence in the late 1960s of an even better polymer with lower viscos­ity and lower cost, hydroxy-terminated polybutadiene (HTPB). It became the industry standard for newer tactical rockets. HTPB had many uses as an adhesive, sealant, and coating, but to employ it in a propellant required, among other things, the development of suit­able bonding agents. These tightly linked the polymer to such solid ingredients as ammonium perchlorate and aluminum. Without such links, the propellant could not withstand the temperature cy­cling, ignition pressure, and other forces that could cause the solid particles to separate from the binder network. This would produce voids in the grain that could result in cracks and structural failure.

A key figure in the development of HTPB for use as a binder was Robert C. Corley, who served as a research chemist and project manager at the Air Force Rocket Propulsion Laboratory at Edwards AFB from 1966 to 1978 and rose through other positions to become the lab’s chief scientist from 1991 to 1997. But many other people from Thiokol, Aerojet, the army at Redstone Arsenal, Atlantic Re­search, Hercules, and the navy were also involved. Even HTPB did

not replace PBAN for all uses, including the Titan III, Titan IVA, and Space Shuttle solid-rocket motors, because PBAN could be pro­duced for the comparatively low cost of $2.50 per pound at a rate in the 1980s of 4 million pounds per year, much higher than for any other propellant.69

To return to Minuteman II, however, the second major change in the stage-two motor was the shift to a single nozzle with liquid thrust vector control replacing movable nozzles for control in pitch and yaw. Static firings had shown that the same propellants pro­duced seven to eight points less specific impulse when fired from four nozzles than from a single one. With the four nozzles, liquid particles agglomerated in their approach sections and produced exit-cone erosion, changing the configuration of the exit cone in 254 an unfavorable way. The solution was not only a single nozzle on Chapter 6 Minuteman II’s second stage but also the change in thrust vector control. The navy had begun testing a Freon system for thrust vec­tor control in the second stage of Polaris A3 in September 1961, well before the Minuteman II, stage-two program began in Febru­ary 1962. The system was low in weight, was insensitive to pro­pellant flame temperature, and posed negligible constraints on the design of the nozzle. The Minuteman engineers adopted it—but one more example of borrowings back and forth between the Polaris and Minuteman projects despite the air force’s view of Polaris as a threat to its roles and missions.

Despite this pioneering work by the navy and its contractors, ac­cording to TRW historians, their firm still had to determine how much “vector capability" stage two of Minuteman II would require. TRW analyzed the amount of injectant that could be used before sloshing in the tank permitted the ingestion of air, and it determined the system performance requirements. Since Aerojet was involved in the development of the system for Polaris, probably its participa­tion in this process was also important. In any event, the Minute – man team, like the navy, used Freon as the injectant, confining it in a rubber bladder inside a metal pressure vessel. Both TRW and Aerojet studied the propensity of the Freon to “migrate" through the bladder wall and become unavailable for its intended purpose. They found that only 25 of 262 pounds of Freon would escape, leav­ing enough to provide the necessary control in pitch and yaw. A separate solid-propellant gas generator provided roll control. In addi­tion to these changes, stage two of Minuteman II increased in length from 159.2 inches for Minuteman I to 162.32 inches. The diameter increased from 44.3 to 52.17 inches, resulting in an overall weight increase from 11,558.9 to 15,506 pounds. Some 3,382.2 of this ad-

ditional 3,947.1 pounds consisted of propellant weight. Even so, the propellant mass fraction decreased slightly from 0.897 to 0.887.70

The Strategic Air Command put the first Minuteman II squadron on operational alert in May 1966, with initial operational capability declared as of December 1966. In the next few years, the air force began replacing Minuteman Is with Minuteman IIs.71


Подпись:Minuteman III featured multiple, independently targetable reentry vehicles with a liquid fourth stage for deployment of this payload. This last feature was not particularly relevant to launch-vehicle de­velopment except that the added weight required for it necessitated higher booster performance. Stages one and two did not change from Minuteman II, but stage three became larger. Hercules lost the contract for the larger motor to Aerojet. Subsequently, Thiokol and a new organization, the Chemical Systems Division of United Technologies Corporation, won contracts to build replacement mo­tors. Stage three featured a fiberglass motor case, the same basic propellant Aerojet had used in stage two only in slightly different proportions, a single nozzle that was fixed in place and partially submerged into the case, a liquid-injection thrust-vector-control system for control in pitch and yaw, a separate roll-control system, and a thrust-termination system.

Aerojet had moved its filament-wound case production to Sacra­mento. It produced most of the Minuteman fiberglass combustion chambers there but ceased winding filament in 1965. Meanwhile, Young had licensed his Spiralloy technology to Black, Sivalls, and Bryson in Oklahoma City, which became a second source for the Minuteman third-stage motor case. This instance and the three firms involved in producing the third stage illustrate the extent to which technology transferred among the contractors and subcon­tractors for government missiles and rockets.

The issue of technology transfer among competing contractors and the armed services, which were also competing over funds and missions, is a complex one about which a whole chapter—even a book—could be written. To address the subject briefly, there had been a degree of effort to exchange knowledge about rocket propul­sion technology beginning in 1946 when the navy provided fund­ing for a Rocket Propellant Information Agency (RPIA) within the Johns Hopkins University’s Applied Physics Laboratory. The army added support in 1948, and the RPIA became the Solid Propellant Information Agency (SPIA). The air force joined the other services in 1951. After the Sputnik launch, the newly created NASA be-

gan participating in SPIA activities in 1959. Meanwhile, the navy created the Liquid Propellant Information Agency (LIPA) in 1958. The SPIA and LPIA combined on December 1, 1962, to create the Chemical Propulsion Information Agency (CPIA).

With the further development of rockets and missiles, the need had become obvious by 1962 for a better exchange of information. So the DoD created an Interagency Chemical Rocket Propulsion Group in November 1962, the name later changing to the Joint Army/Navy/NASA/Air Force (JANNAF) Interagency Propulsion Committee. Together with CPIA, JANNAF effectively promoted sharing of technology. In addition, “joint-venture" contracts, pio­neered by Levering Smith of the navy, often mandated the sharing of manufacturing technology among companies. These contracts 256 served to eliminate the services’ dependence for a given technology Chapter 6 on sole sources that could be destroyed by fire or possible enemy targeting. It also provided for competitive bidding on future con­tracts. The air force had a similar policy.72

Meanwhile, the propellant for Aerojet’s third stage of Minuteman had less CTPB and more aluminum than the second stage. The grain configuration consisted of an internal-burning cylindrical bore with six “fins" radiating out in the forward end. The igniter used black – powder squibs to start some of the CTPB propellant, which in turn spread the burning to the grain itself. The 50 percent submerged nozzle had a graphite phenolic entrance section, a forged tungsten throat insert, and a carbon-phenolic exit cone. As compared with the 85.25-inch-long, 37.88-inch-diameter third stage of Minuteman II, that for Minuteman III was 91.4 inches long and 52 inches in di­ameter. The mass fraction improved from 0.864 to 0.910, and with a nearly 10-lbf-sec/lbm greater specific impulse, the new third stage had more than twice the total impulse of its predecessor—2,074,774 as compared with 1,006,000 pounds force per second.73

The thrust-vector-control system for the new stage three was similar to that for stage two except that strontium perchlorate was used instead of Freon as the injectant into the thrust stream to pro­vide control in the pitch and yaw axes. Helium gas provided the pressure to insert the strontium perchlorate instead of the solid – propellant gas generator used in the second stage. Roll control again came from a gas generator supplying gas to nozzles pointing in opposite directions. When both were operating, there was neutral torque in the roll axis. When roll torque was required, the flight – control system closed a flapper on one of the nozzles, providing unbalanced thrust to stop any incipient roll.

To ensure accuracy for the delivery of the warheads, Minute – man had always required precise thrust termination for stage three, determined by the flight-control computer. On Minuteman I, the thrust-termination system consisted of four thick carbon-phenolic tubes integrally wound in the sidewall of the third-stage case and sealed with snap-ring closures to form side ports. Detonation of explosive ordnance released a frangible section of the snap ring, thereby venting the combustion chamber and causing a momentary negative thrust that resulted in the third stage dropping away from the postboost vehicle.

Подпись:The system for Minuteman III involved six circular-shaped charges on the forward dome. Using data from high-speed films and strain gauges, the Minuteman team learned that this arrange­ment worked within 20 microseconds, cutting holes that resulted in a rupture of the pressure vessel within 2 additional milliseconds. But the case developed cracks radiating from the edge of the holes. TRW used a NASTRAN computer code to define propagation of the cracks. It then determined the dome thickness needed to eliminate the failure of the fiberglass. Aerojet wound “doilies" integrally into the dome of the motor case under each of the circular charges. This eliminated the rupturing, allowing the system to vent the pressure in the chamber and produce momentary negative thrust.74

Minuteman Ills achieved their initial operational capability in June 1970, the first squadron of the upgraded missiles turned over to an operational wing at Minot AFB, North Dakota, in January 1971. By July 1975, there were 450 Minuteman Ils and 550 Minute – man Ills deployed at Strategic Air Command bases.75