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

THE RV-A-10 MISSILE

In the meantime, Thiokol had teamed up with General Electric in the Hermes project to produce a solid-propellant missile (initially

known as the A-2) that was much larger than the Sergeant sounding rocket. It operated on a shoestring budget until canceled, but it still made significant progress in solid-propellant technology.16

The original requirements for the A-2 were to carry a 500-pound warhead to a range as far as 75 nautical miles, but these changed to a payload weighing 1,500 pounds, necessitating a motor with a diameter of 31 inches. Thiokol started developing the motor in May 1950, a point in time that allowed the project to take advantage of work on the Sergeant sounding rocket and of Larry Thackwell’s ex­perience with it. By December 1951, the program had successfully completed a static test of the 31-inch motor. From January 1952 through March 1953, there were 20 more static tests at Redstone Arsenal and four flight tests of the missile at Patrick AFB, Florida.

230 In the process, the missile came to be designated the RV-A-10. The Chapter 6 project encountered unanticipated problems with nozzle erosion and combustion instability that engineers were able to solve.17

The four flight tests achieved a maximum range of 52 miles (on flight one) and a maximum altitude of 195,000 feet (flight two) us­ing a motor case 0.20 inch thick and a propellant grain featuring a star-shaped perforation with broad tips on the star. The propellant was designated TRX-110A. It included 63 percent ammonium per­chlorate by weight as the oxidizer. The propellant took advantage of an air force-sponsored project (MX-105) titled “Improvement of Polysulfide-Perchlorate Propellants" that had begun in 1950 and is­sued a final report (written by Thiokol employees) in May 1951. On test motor number two a propellant designated T13, which con­tained polysulfide LP-33 and ammonium perchlorate, achieved a specific impulse at sea level of more than 195 lbf-sec/lbm at 80°F but also experienced combustion instability. This led to the shift to TRX-110, which had a slightly lower specific impulse but no com­bustion instability.18

Thiokol had arrived at the blunter-tipped star perforation as a re­sult of Thackwell’s experience (at JPL) with the Sergeant test vehi­cle and of photoelastic studies of grains performed at the company’s request by the Armour Institute (later renamed the Illinois Institute of Technology). This, together with a thicker case wall than JPL had used with the Sergeant sounding rocket, eliminated JPL’s prob­lems with cracks and explosions. However, TRX-110 proved not to have enough initial thrust. The solution was to shift the size of the ammonium perchlorate particles from a mixture of coarse and fine pieces to one of consistently fine particles, which yielded not only higher initial thrust but also a more consistent thrust over time—a desirable trait. Meanwhile, the Thiokol-GE team gradually learned

about the thermal environment to which the RV-A-10 nozzles were exposed. Design of the nozzles evolved through subscale and full – scale motor tests employing various materials and techniques for fabrication. The best materials proved to be SAE 1020 steel with carbon inserts, and a roll weld proved superior to casting or forging for producing the nozzle itself.19

Подпись:Another problem encountered in fabricating the large grain for the RV-A-10 was the appearance of cracks and voids when it was cured at atmospheric pressure, probably the cause of a burnout of the liner on motor number two. The solution proved to be twofold: (1) Thiokol poured the first two mixes of propellant into the motor chamber at a temperature 10°F hotter than normal, with the last mix 10°F cooler than normal; then, (2) Thiokol personnel cured the propellant under 20 pounds per square inch of pressure with a layer of liner material laid over it to prevent air from contacting the grain. Together, these two procedures eliminated the voids and cracks.20

With these advances in the art of producing solid-propellant mo­tors, the RV-A-10 became the first known solid-propellant rocket motor of such a large size—31 inches in diameter and 14 feet, 4 inches long—to be flight tested as of February-March 1953. Among its other firsts were scaling up the mixing and casting of polysulfide propellants to the extent that more than 5,000 pounds of it could be processed in a single day; the routine use of many mixes in a single motor; the use of a tubular igniter rolled into coiled plastic tubing (called a jelly roll) to avoid the requirement for a heavy clo­sure at the nozzle end to aid in ignition; and one of the early uses of jet vanes inserted in the exhaust stream of a large solid-propellant rocket to provide thrust vector control.

As recently as December 1945, the head of the Office of Scientific Research and Development during World War II, Vannevar Bush, had stated, “I don’t think anybody in the world knows how [to build an accurate intercontinental ballistic missile] and I feel confident it will not be done for a long time to come." Many people, even in the rocket field, did not believe that solid-propellant rockets could be efficient enough or of long enough duration to serve as long-range missiles. The RV-A-10 was the first rocket to remove such doubts from at least some people’s minds.21 Arguably, it provided a signifi­cant part of the technological basis for the entire next generation of missiles, from Polaris and Minuteman to the large solid boost­ers for the Titan IIIs and IVs and the Space Shuttle, although many further technological developments would be necessary before they became possible (including significant improvements in propellant performance).

The Space Shuttle, 1972-91

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

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

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

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

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

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

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

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

Подпись: 93 U.S. Space-Launch Vehicles, 1958-91 The sophistication of the SSME explained all its problems. “In assessing the technical difficulties that have been causing delays in the development and flight certification of the SSME at full power, it is important to understand that the engine is the most advanced liquid rocket motor ever attempted," wrote an ad hoc committee of the Aeronautics and Space Engineering Board in 1981. “Chamber pressures of more than 3,000 psi, pump pressures of 7,000-8,000 psi, and an operating life of 7.5 hours have not been approached in previous designs of large liquid rocket motors."122

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

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

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

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

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

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

FIG. 2.11

The static test of Solid Rocket Booster (SRB) Demonstration Model 2 (DM-2) at the Thiokol test site near Brigham City, Utah. (Photo courtesy of NASA)

 

The Space Shuttle, 1972-91

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

Подпись: 95 U.S. Space-Launch Vehicles, 1958-91 The air force had a great deal of influence on the requirements for the shuttle because its support had been needed to get the program approved and make it viable economically. NASA needed a com­mitment from the military that all of its launch needs would be carried on the shuttle. To satisfy DoD requirements, the shuttle had to handle payloads 60 feet long with weights of 40,000 pounds for polar orbits or 65,000 pounds for orbits at the latitude of Kennedy Space Center. On July 26, 1972, NASA announced that the Space Transportation Systems Division of North American Rockwell had won the contract for the orbiters.129

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

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

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

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

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

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

Подпись: 97 U.S. Space-Launch Vehicles, 1958-91 Following the accident there was an extensive redesign of many aspects of the shuttle, notably the field joints. This new design al­legedly ensured that the seals would not leak under twice the an­ticipated structural deflection. Following Challenger, both U. S. policy and law changed, essentially forbidding the shuttle to carry commercial satellites and largely restricting the vehicle to missions both using the shuttle’s unique capabilities and requiring people to be onboard. A concomitant result was the rejuvenation of the air force’s expendable launch-vehicle program. Although the Delta II was the only launcher resulting directly from the 32-month hia­tus in shuttle launches following the accident, the air force also ordered more Titan IVs and later, other expendable launch vehicles. The shuttle became a very expensive launch option because its eco­nomic viability had assumed rapid turnaround and large numbers of launches every year. Yet in 1989 it flew only five missions, in­creased to six in 1990 and 1991.135

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

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

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

Titan II Engines

Подпись:Although there were other important upper-stage engines using storable propellants, such as the Bell Agena engine (starting with the 8048 model), the most significant engines with hypergolic pro­pellants were those used in the Titans II, III, and IV, designed and built by Aerojet. Here, the previous experience with Vanguard, Able, and Able-Star undoubtedly were extraordinarily valuable, as was Aerojet’s involvement with the development of UDMH. In the lat­ter development, Aerojet propellant chemist Karl Klager had been an important contributor. Klager held a Ph. D. in chemistry from the University of Vienna (1934) and had come to this country as part of Project Paperclip quite independently of the von Braun group. He worked for the Office of Naval Research in Pasadena during 1949 and started with Aerojet in 1950. The following year, Aerojet received a contract to develop an in-flight thrust-augmentation rocket for the F-86 fighter. The device never went into production, but in develop­ing it, Aerojet engineers conducted a literature search for candidate propellants, did theoretical performance calculations, and measured physical and chemical properties in the laboratory. Together with RFNA, UDMH seemed highly promising but was not available in sufficient quantities to be used. Klager devised (and patented) pro­duction processes that yielded large quantities at reasonable prices, but workers began to get violently ill from the toxic substance. All did recover, and Aerojet learned how to control exposure to the vapors.37

FIG. 4.1

Technical drawing with description of the Agena upper stage as of 1968. (Photo courtesy of NASA)

 

NASA

C-1968-538

 

AGENA

PROJECT

 

Titan II EnginesTitan II Engines

AGENA.. . versatile, upper-stage rocket vehicle employs a single rocket engine which provides 16,000 pounds of thrust. The engine can be shut down and re­started in flight through ground command signals.

Подпись: 160 Chapter 4 Agena and its payload ride into space aboard a large booster rocket. Following staging, the Agena engine "first-burn" maneuvers the vehicle and its payload into an earth – oriented parking orbit. The Agena "second-burn" is geared to each particular mission – for example, an ellip­tical earth orbit or the ejection of a payload on a trajectory to the moon or planets.

This photographic exhibit presents the Agena missions. . . managed by the Lewis Research Center since January 1963.

National Aeronautics and Space Administration Lewis Research Center

When the time came in 1960 to begin designing engines for the Titan II, UDMH had a lower specific impulse than the new missile required. Hydrazine had better performance but could detonate if used as a regenerative coolant. Aerojet was the first firm to come up with an equal mixture of hydrazine and UDMH, Aerozine 50. This fuel combination ignited hypergolically with nitrogen tetrox – ide as the oxidizer. Neither was cryogenic. And both could be stored in propellant tanks for extended periods, offering a much quicker response time than Titan I’s 15 minutes for a propulsion system burning liquid oxygen and kerosene. Both Aerojet and Martin, the
overall contractor for Titan I, urged a switch to the new propellants, but it was apparently Robert Demaret, chief designer of the Titan, and others from Martin who proposed the idea to the air force’s Bal­listic Missile Division in early 1958.38

Подпись:In May 1960, the air force signed a letter contract with the Mar­tin Company to develop, produce, and test the Titan II. On Octo­ber 9, 1959, Aerojet had already won approval to convert the Titan I engines to burn storable propellants. Research and development to that end began in January 1960. The Aerojet engineers also worked to achieve the improved performance called for in the April 30, 1960, plan for Titan II. Although the Titan II engines were based on those for Titan I, the new propellants and the requirements in the April 30 plan necessitated considerable redesign. Because the modi­fications did not always work as anticipated, the engineers had to resort to empirical solutions until they found the combinations that worked correctly and provided the necessary performance. Since the propellants were hypergolic, there was no need for an igniter in the Titan II engines (called XLR87-AJ-5 for the two engines in stage one and XLR91-AJ-5 for the single stage-two engine). The injectors for the Titan I engines had used alternating fuel and oxidizer passages with oxidizer impinging on oxidizer and fuel on fuel (called like-on – like) to obtain the necessary mixing of the two propellants. For the Titan II, Aerojet engineers tried fuel-on-oxidizer impingement. This evidently mixed the droplets of propellant better because higher performance resulted. But the improvement led to erosion of the injector face, necessitating a return to the like-on-like pattern.

This older arrangement caused combustion instability in the stage-two engines, and engineers tried several configurations of baf­fles to solve the problem before they came up with one that worked. One potential solution, uncooled stainless-steel baffles, did not last through a full-duration engine test. Copper baffles with both pro­pellants running through them for cooling resulted in corrosion of the copper from the nitrogen tetroxide. An eight-bladed configura­tion cooled by the oxidizer (evidently using another type of metal) yielded poor performance. The final configuration was a six-bladed, wagon-wheel design (with the baffles radiating outward from a cen­tral hub), again cooled by the oxidizer. This solved the problem, at least at for the time being.39

The turbopumps for Titan II were similar to those for Titan I, but differences in the densities of the propellants necessitated greater power and lower shaft speed for the Titan II pumps. This resulted in increased propellant flow rates. But Aerojet engineers had to rede­sign the gears for the turbopumps, making them wider and thus able

Подпись: 162 Chapter 4

Подпись: FIG. 4.2 Generic technical drawing of a liquid- propellant rocket showing some of its components, such as a turbine gas generator and a turbopump. (Photo courtesy of NASA)
Titan II Engines

to withstand greater “tooth pressures" caused by the higher power. From two blades in the inducer for Titan I, the design went to three blades. Engineers also had to redesign the impeller and housing pas­sages to accept the higher flow rates, and there had to be new mate­rials that would not be degraded by the storable propellants.40

A significant innovation for the Titan II replaced the use of pressurized nitrogen (in stage one of Titan I) or helium (in stage two) to initiate propellant flow with a so-called autogenous (self­generating) system. Solid-propellant start cartridges initiated the process by spinning the turbines, whereupon gas generators kept the turbines spinning to pressurize the fuel tanks in both stages and to pump the Aerozine 50 into the thrust chamber. The second – stage oxidizer tank did not need pressurization because acceleration was sufficient to keep the nitrogen tetroxide flowing. In the first – stage propulsion system, though, oxidizer from the pump discharge served to pressurize the tank. The result was a simplified system saving the weight from the pressurized-gas storage tanks in Titan I and requiring no potentially unreliable pressure regulators. A simi­lar increase in reliability and saving in weight came from using the exhaust stream from the turbopump in stage two to provide roll control in place of an auxiliary power-drive assembly used in Titan I for the vernier thrusters. Gimbals (used in Titan I) continued to
provide pitch, roll, and yaw control in the first stage plus pitch and yaw control in the second stage.41

The Titan II propulsion system had significantly fewer parts than its Titan I predecessor. The number of active control components fell from 125 to 30; valves and regulators declined from 91 to 16. The engines had higher thrust and performance, as planned. The Titan II first-stage engines had a combined thrust of 430,000 pounds at sea level, compared to 300,000 for Titan I. Second-stage thrust rose from 80,000 pounds for Titan I to 100,850 pounds for Titan II. Specific impulses rose less dramatically, from slightly above 250 to almost 260 lbf-sec/lbm at sea level for stage one and remained slightly above 310 at altitude (vacuum) for stage two.42

Подпись:A 1965 Aerojet news release on Titan II propulsion credited it to “the efforts of hundreds of men and women" but singled out four of them as leaders of the effort. Three of their biographies illustrate the way that engineers in aerospace migrated from one firm to an­other or from government work to the private sector, carrying their knowledge of various technologies with them. Robert B. Young was the overall manager of the design, development, and production ef­fort for both Titan I and Titan II. He was a chemical engineering graduate of Caltech, where Theodore von Karman had encouraged him to devote his knowledge to rocketry. He had worked for a year as director of industrial liaison on the Saturn program at NASA’s Marshall Space Flight Center, during part of the Titan years, and had risen within Aerojet’s own structure from a project engineer to a vice president and manager of the Sacramento, California, plant. Another leader was Ray C. Stiff Jr., who had discovered aniline as a hypergolic propellant while working at the navy Engineering Exper­iment Station in Annapolis. His role in the use of self-igniting pro­pellants in JATOs to assist a PBY into the air while carrying heavy loads proved a forerunner “of the storable, self-igniting propellants used to such advantage in the Titan II engine systems." Stiff served as Aerojet’s manager of Liquid Rocket Operations near Sacramento and was also a vice president.

A third manager Aerojet mentioned in the release was A. L. Feldman, a rocket engineer educated at Cornell University. While working at Convair, he had been “in on the initial design and devel­opment effort for the Atlas engines." After moving to Aerojet, he served as manager for both the Titan I and Titan II engines and as­sistant manager of Aerojet’s Liquid Rocket Operations. A fourth key manager was L. D. Wilson, who earned a degree in engineering at Kansas State University. “At 28, he managed the painstaking design, development, test and production of the ‘space start’ second-stage

engine for Titan I." Wilson then managed the entire propulsion sys­tem for the Titan II.43 Alone among the four managers, he appeared not to have worked for another rocket effort.

Between March 16, 1962, and April 9, 1963, there were 33 re- search-and-development test flights of Titan II missiles—23 from Cape Canaveral and 10 from Vandenberg AFB. Depending on who was counting, there were variously 8, 9, or 10 failures and partial successes (none of which occurred during the last 13 flights), for a success rate of anywhere from 70 to 76 percent. The problems ranged from failure of electrical umbilicals to disconnect properly on the first, otherwise-successful, Titan II silo launch (from Van­denberg) on February 16, 1963 (pulling missile-guidance cabling with them and causing an uncontrollable roll), to premature en­gine shutdown, an oxidizer leak, a fuel-valve failure, a leak in a fuel pump, and gas-generator failure. But the most serious problem was initially a mystery whose cause was unclear to the engineers in­volved. On the launch of the first Titan II on March 16, 1962, about a minute and a half after liftoff from Cape Canaveral, longitudinal oscillations occurred in the first-stage combustion chambers. They arose about 11 times per second for about 30 seconds. They did not prevent the missile (designated N-2) from traveling 5,000 nautical miles and impacting in the target area, but they were nonetheless disquieting.44

164 The reason for concern was that in late 1961 NASA had reached Chapter 4 an agreement with the Department of Defense to acquire Titan IIs for launching astronauts into space as part of what soon became Project Gemini. Its mission, as a follow-on to Project Mercury and a predecessor to Project Apollo’s Moon flights, was to determine if one spacecraft could rendezvous and dock with another, whether astronauts could work outside a spacecraft in the near weightless­ness of space, and what physiological effects humans would experi­ence during extended flight in space. Pogo (as the longitudinal oscil­lations came to be called ) was not particularly problematic for the missile, but it posed potentially significant problems for an astro­naut who was already experiencing acceleration of about 2.5 times the force of gravity from the launch vehicle. The pogo effect on the first Titan II launch added another ±2.5 Gs, which perhaps could have incapacitated the pilot of the spacecraft from responding to an emergency if one occurred.

Fixing the problem for Project Gemini, however, was compli­cated by an air force reorganization on April 1, 1961, creating Bal­listic Systems Division (BSD) and Space Systems Division (SSD). The problem for NASA lay in the fact that while BSD was intent

on developing Titan II as a missile, SSD would be responsible for simultaneously adapting Titan II as a launch vehicle for the astro­nauts.45 A conflict between the two air force interests would soon develop and be adjudicated by General Schriever, then commanding the parent Air Force Systems Command.

Meanwhile, there were further organizational complications for Project Gemini. SSD assigned development of the Gemini-Titan II launch vehicle to Martin’s plant in Baltimore, whereas the Denver plant was working on the Titan II missile. Assisting SSD in manag­ing its responsibilities was the nonprofit Aerospace Corporation. For Gemini, Aerospace assigned James A. Marsh as manager of its efforts to develop the Titan II as a launch vehicle. BSD established its own committee to investigate the pogo oscillations, headed by Abner Rasumoff of Space Technology Laboratories (STL). For the missile, it found a solution in higher pressure for the stage-one fuel tank, which reduced the oscillations and resultant gravitational forces in half on the fourth launch on July 25, 1962, without STL engineers’ understanding why. Martin engineers correctly thought the problem might lie in pressure oscillations in the propellant feed lines. They suggested installation of a standpipe to suppress surges in the oxidizer lines of future test missiles. BSD and NASA’s Manned Spacecraft Center, which was managing Gemini, agreed.46

Подпись:Although the standpipe later proved to be part of the solution to the pogo problem, initially it seemed to make matters worse. Installed on missile N-11, flying on December 6, 1962, it failed to suppress severe oscillations that raised the gravitational effect from pogo alone to ±5 Gs. This lowered the chamber pressure to the point that instrumentation shut down the first-stage engine prema­turely. The following mission, N-13, on December 19, 1962, did not include the standpipe but did have increased pressure in the fuel tank, which had seemed to be effective against the pogo effect on earlier flights. Another new feature consisted of aluminum oxidizer feed lines in place of steel ones used previously. For reasons not fully understood, the pogo level dropped and the flight was successful.

Missile N-15, evidently with the same configuration as N-13, launched on January 10, 1963. The pogo level dropped to a new low, ±0.6 G, although problems with the gas generator in stage two se­verely restricted the vehicle’s range. This was still not a low enough oscillation level for NASA, which wanted it reduced to ±0.25 G, but it satisfied BSD, which “froze" the missile’s design with regard to further changes to reduce oscillations. Higher pressure in the first-stage fuel tanks plus use of aluminum for the oxidizer lines had reduced the pogo effect below specifications for the missile,

and BSD believed it could not afford the risks and costs of further experimentation to bring the pogo effect down to the level NASA wanted.47

NASA essentially appealed to Schriever, with NASA’s Brainerd Holmes, deputy associate administrator for manned space flight, complaining that no one understood what caused either the pogo oscillations or unstable combustion, another problem affecting Titan II. So Holmes said it was impossible to “man-rate" the mis­sile as a launch vehicle. The result, on April 1, 1963, was the forma­tion of a coordinating committee to address both problems, headed by BSD’s Titan II director and including people from Aerospace Cor­poration and STL.

After engineers from Aerojet, Martin, STL, and Aerospace stud­ied the problems, Sheldon Rubin of Aerospace looked at data from static tests and concluded that as the fuel pumped, a partial vacuum formed in the fuel line, causing resonance. This explained why the oxidizer standpipes had failed to suppress the pogo effect. The solu­tion was to restore the standpipe, keep the increase in tank pressure and the aluminum oxidizer feed lines, and add a fuel surge chamber (also called a piston accumulator) to the fuel lines. Nitrogen gas pressurized the standpipe after the nitrogen tetroxide had filled the oxidizer feed lines. An entrapped gas bubble at the end of the stand­pipe absorbed pressure pulsations in the oxidizer lines. The surge 166 chamber included a spring-loaded piston. Installed perpendicular to Chapter 4 the fuel feed line, it operated like the standpipe to absorb pressure pulses. Finally, on November 1, 1963, missile N-25 carried both of these devices. The successful flight recorded pogo levels of only ±0.11 G, well below NASA’s maximum of ±0.25 G. Subsequent tests on December 12, 1963, and January 15, 1964, included the sup­pression devices and met NASA standards, the January 15 mission doing so even with lower pressures in the fuel tank. This seemed to confirm that the two devices on the propellant lines had fixed the pogo problem.48

Meanwhile, the combustion instability Holmes had mentioned at the same time as the pogo problem, turned out to be another major issue. It never occurred in flight, but it appeared in a severe form during static testing of second-stage engines. Several engines experienced such “hard starts" that the combustion chambers fell from the injector domes as if somebody had cut them away with a laser beam. Engineers examined the test data and concluded that combustion instability at the frequency of 25,000 cycles per second had sliced through the upper combustion chamber near the face of the injector with a force of ± 200 pounds per square inch. This oc-

curred on only 2 percent of the ground tests of second-stage engines, but for Gemini, even this was too high.

Aerojet instituted the Gemini Stability Improvement Program (GEMSIP) in Sept ember 1963 to resolve the problem with com­bustion instability. Apparently, the instability occurred only when second-stage engines were tested in an Aerojet facility that simu­lated the air pressure at 70,000 feet, because there was no combus­tion instability in the first-stage engines at this stage of their devel­opment. The reason was that air pressure at sea level slowed the flow of propellants through the injectors. Aerojet engineers could solve the problem for the second stage by filling the regenerative – cooling tubes that constituted the wall of the combustion chamber with a very expensive fluid that afforded resistance to the rapid flow of propellants similar to that of air pressure at sea level.

Aerojet and the air force finally agreed on a more satisfactory solution, however. This involved a change in injector design with fewer but larger orifices and a modification of the six-bladed baffles radiating out from a hub that Aerojet thought had already solved the combustion-instability problem in the Titan II missile. Aerojet engineers increased the number of baffles to seven and removed the hub for the Gemini configuration. Testing in the altitude chamber at the air force’s Arnold Engineering Development Center proved that the new arrangement worked.49

Подпись:A final problem that occurred in flight testing and required re­engineering involved the gas generator in the stage-two engine. This issue was a matter of concern to the Titan II Program Office at BSD, not just to SSD and NASA. Engineers and managers first became aware there was a problem on the second Titan II test flight (June 7, 1962) when telemetry showed that the second-stage engine had achieved only half of its normal thrust soon after engine start. When the tracking system lost its signal from the vehicle, the range safety officer caused cutoff of the fuel flow, making the reentry ve­hicle splash down well short of its target area. The data from telem­etry on this flight were inadequate for the engineers to diagnose the problem. It took two more occurrences of gas-generator problems to provide enough data to understand what was happening. Particles were partially clogging the small openings in the gas generator’s in­jectors. This restricted propellant flow and resulted in loss of thrust. Technicians very thoroughly removed all foreign matter from com­ponents of the generators in a clean room before assembly, sepa­rated the generators from the engines in transport, and subjected the assemblies to blowdown by nitrogen before each test flight to ensure no foreign particles were present.

These methods did not solve the problem but did narrow the list of sources for the particles. It became apparent that there was no problem with stage-one gas generators because sea-level air pressure did not allow particles from the solid-propellant start cartridges to reach the injector plates for the generators. Aerojet had tested the system for stage two in its altitude chamber, but it could simulate only 70,000 feet of altitude, which engineers assumed was high enough. There had been no problems with gas generators during the tests. It turned out, however, that even at 70,000 feet, enough atmo­sphere was present to cushion the injectors from the particles. At 250,000 feet, where the stage-two engine ignited, the atmosphere was so thin that particles from the start cartridge flowed into the gas generator, sometimes in sufficient quantity to cause difficulties. To rectify the problem, engineers added a rupture disk to the exhaust of the gas generator. This kept enough pressure in the system to cush­ion the flow of particles until ignition start, when the disk ruptured. Gas-generator failure was not a problem after this design change.50

The engines for Titan II went on to form the core of the Titan III and IV space-launch vehicles. Titan IIs also became launch vehicles in their own right, making their engines key propulsion elements in three different series of launch vehicles. Their use through the end of the period covered by this book suggested the importance of storable propellants for the history of space flight. But development 168 of hypergolic propulsion did not end with these rockets.

Chapter 4

The Atlas, Thor, and Jupiter Missiles

Following Redstone and Vanguard, the Atlas, Thor, and Jupiter mis­siles brought further innovations in rocket technology and became the first stages of launch vehicles themselves, with Atlas and Thor having more significance in this role than Jupiter. All three pro­grams illustrated the roles of interservice and interagency rivalry and cooperation that were both key features of rocket development in the United States. They also showed the continued use of both theory and empiricism in the complex engineering of rocket sys­tems. “It was not one important ‘breakthrough’ that enabled this advance; rather, it was a thousand different refinements, a hundred thousand tests and design modifications, all aimed at the develop­ment of equipment of extraordinary power and reliability," accord­ing to Milton Rosen, writing in 1962.61

Atlas was a much larger effort than Vanguard, and it began to create the infrastructure in talent, knowledge, data, and capability necessary for the maturation of launch-vehicle technology in the decade of the 1960s. However, until the air force became serious about Atlas, that service had lagged behind the army and the navy in the development of purely ballistic missiles.62

The process began in a significant way on January 23, 1951, when 32 the air force awarded the Consolidated Vultee Aircraft Corporation Chapter 1 (Convair) a contract for MX-1593, the project that soon became Atlas. (MX-1593 had been preceded by MX-774B and a number of other air force missile contracts in the late 1940s, with a total of $34 million devoted to missile research in fiscal year 1946, much re-

duced in subsequent years.) But the specifications for the MX-1593 missile changed drastically as technology for nuclear warheads evolved to fit more explosive power into smaller packages. This new technology plus the increased threat from the Soviets provided one condition for greater air force support of Atlas and other bal­listic missiles.

But it also took two heterogeneous engineers to nudge the new­est armed service and the Department of Defense (DoD) in a new direction. One of them was Trevor Gardner, assistant for research and development to Secretary of the Air Force Harold Talbott in the Eisenhower administration. The other key promoter of ballistic missiles was the “brilliant and affable" polymath, John von Neu­mann, who was research professor of mathematics at Princeton’s Institute for Advanced Study and also director of its electronic com­puter project. In 1953, he headed a Nuclear Weapons Panel of the Air Force Scientific Advisory Board, which confirmed beliefs that in the next six to eight years, the United States would have the capability to field a thermonuclear warhead weighing about 1,500 pounds and yielding 1 megaton of explosive force. This was 50 times the yield of the atomic warhead originally planned for the Atlas missile, fit in a much lighter package. This and a report (dated February 1, 1954) for von Neumann’s Teapot Committee set the stage for extraordi­nary air force support for Atlas.63

Подпись: 33 German and U.S. Missiles and Rockets, 1926-66 In May 1954 the air force directed that the Atlas program be­gin an accelerated development schedule, using the service’s high­est priority. The Air Research and Development Command within the air arm created a new organization in Inglewood, California, named the Western Development Division (WDD), and placed Brig. Gen. Bernard A. Schriever in charge. Schriever, who was born in Germany but moved to Texas when his father became a prisoner of war there during World War I, graduated from the Agricultural and Mechanical College of Texas (since 1964, Texas A&M Univer­sity) in 1931 with a degree in architectural engineering. Tall, slen­der, and handsome, the determined young man accepted a reserve commission in the army and completed pilot training, eventu­ally marrying the daughter of Brig. Gen. George Brett of the army air corps in 1938. Placed in charge of the WDD, Schriever in es­sence took over from Gardner and von Neumann the role of het­erogeneous engineer, promoting and developing the Atlas and later missiles.64

Gardner had been intense and abrasive in pushing the develop­ment of missiles. Schriever was generally calm and persuasive. He selected highly competent people for his staff, many of them be-

coming general officers. An extremely hard worker, like von Braun he demanded much of his staff; but unlike von Braun he seemed somewhat aloof to most of them and inconsiderate of their time— frequently late for meetings without even realizing it. Good at plan­ning and organizing, gifted with vision, he was poor at management, often overlooking matters that needed his attention—not surprising because he spent much time flying back and forth to Washington, D. C. His secretary and program managers had to watch carefully over key documents to ensure that he saw and responded to them. One of his early staff members (later a lieutenant general), Otto J. Glasser, said Schriever was “probably the keenest planner of any­body I ever met" but he was “one of the lousiest managers."

Later Lt. Gen. Charles H. Terhune Jr., who became Schriever’s deputy director for technical operations, called his boss a “superb front man" for the organization, “very convincing. . . . He had a lot of people working for him [who] were very good and did their jobs, but Schriever was the one who pulled it all together and represented them in Congress and other places." Glasser added, “He was just superb at. . . laying out the wisdom of his approach so that the Congress wanted to ladle out money to him." Glasser also said he was good at building camaraderie among his staff.65

To facilitate missile development, Schriever received from the air force unusual prerogatives, such as the Gillette Procedures. Designed by Hyde Gillette, a budgetary expert in the office of the secretary of the air force, these served to simplify procedures for managing intercontinental ballistic missiles (ICBMs). Schriever had complained that there were 40 different offices and agencies he had to deal with to get his job done. Approval of his annual develop­ment plans took months to sail through all of these bodies. With the new procedures (granted November 8, 1955), Schriever had to deal with only two ballistic missile committees, one at the secretary – of-defense and the other at the air-force level. Coupled with other arrangements, this gave Schriever unprecedented authority to de­velop missiles.66

Another key element in the management of the ballistic mis­sile effort was the Ramo-Wooldridge Corporation. Simon Ramo and Dean Wooldridge, classmates at Caltech, each had earned a Ph. D. there at age 23. After World War II, they had presided over an 34 electronics team that built fire-control systems for the air force at Chapter 1 Hughes Aircraft. In 1953, they set up their own corporation, with the Thompson Products firm buying 49 percent of the stock. For a variety of reasons, including recommendations of the Teapot Com­mittee, Schriever made Ramo-Wooldridge into a systems engineer-

ing-technical direction contractor to advise his staff on the man­agement of the Atlas program. The air force issued a contract to the firm for this task on January 29, 1955, although it had begun working in May 1954 under letter contract on a study of how to redirect the Atlas program. This unique arrangement with Ramo- Wooldridge caused considerable concern in the industry (especially on the part of Convair) that Ramo-Wooldridge employees would be in an unfair position to use the knowledge they gained to bid on other contracts, although the firm was not supposed to produce hardware for missiles. To ward off such criticism, the firm created a Guided Missile Research Division (GMRD) and kept it separate from other divisions of the firm. Louis Dunn, who had served on the Teapot Committee, became the GMRD director, bringing sev­eral people with him from JPL. This arrangement did not put an end to controversy about Ramo-Wooldridge’s role, so in 1957, GMRD became Space Technology Laboratories (STL), an autonomous divi­sion of the firm, with Ramo as president and Dunn as executive vice president and general manager.67

Some air force officers on Schriever’s staff objected to the con­tract with Ramo-Wooldridge, notably Col. Edward Hall, a propul­sion expert. Hall had nothing good to say about Ramo-Wooldridge (or Schriever), but several engineers at Convair concluded that the firm made a positive contribution to Atlas development.68

The Ramo-Wooldridge staff outnumbered the air force staff at WDD, but the two groups worked together in selecting contrac­tors for components of Atlas and later missiles, overseeing their performance, testing, and analyzing results. For such a large under­taking as Atlas, soon joined by other programs, there needed to be some system to inform managers and allow them to make decisions on problem areas. The WDD, which became the Air Force Ballistic Missile Division on June 1, 1957, developed a management control system to collect information for planning and scheduling.

Подпись: 35 German and U.S. Missiles and Rockets, 1926-66 Schriever and his program directors gathered all of this data in a program control room, located in a concrete vault and kept under guard at all times. At first, hundreds of charts and graphs covered the walls, but WDD soon added digital computers for tracking infor­mation. Although some staff members claimed Schriever used the control room only to impress important visitors, program managers benefited from preparing weekly and monthly reports of status, be­cause they had to verify their accuracy and thereby keep abreast of events. Separate reports from a procurement office the Air Force Air Materiel Command assigned to the WDD on August 15, 1954, pro­vided Schriever an independent check on information from his own

managers. The thousands of milestones—Schriever called them inchstones—in the master schedule kept him and his key manag­ers advised of how development matched planning. All of the infor­mation came together on “Black Saturday" meetings once a month starting in 1955. Here program managers and department heads presented problem areas to Schriever, Ramo, and Brig. Gen. Ben I. Funk, commander of the procurement office. As problems arose, discussion sometimes could resolve them in the course of the meet­ing. If not, a specific person or organization would be assigned to come up with a solution, while the staff of the program control room tracked progress. Sometimes, Ramo brought in outside experts from industry or academia to deal with particularly difficult problems.69

Because the process of developing new missile systems entailed considerable urgency when the Soviet threat was perceived as great and the technology was still far from mature, Schriever and his team used a practice called concurrency that was not new but not routinely practiced in the federal government. Used on the B-29 bomber, the Manhattan Project, and development of nuclear vessels for the U. S. Navy, it involved developing all subsystems and the facilities to test and manufacture them on overlapping schedules; likewise, the systems for operational control and the training sys­tem for the Strategic Air Command, which took over the missiles when they became operational.

Schriever claimed that implementing concurrency was equiva­lent to requiring a car manufacturer to build the automobile and also to construct highways, bridges, and filling stations as well as teach drivers’ education. He argued that concurrency saved money, but this seems doubtful. Each model of the Atlas missile from A to F involved expensive improvements, and the F models were housed in silos. Each time the F-model design changed, the Army Corps of Engineers had to reconfigure the silo. There were 199 engineering change orders for the silos near Lincoln, Nebraska, and these raised the costs from $23 million to more than $50 million dollars—to give one example of costs added by concurrency. What concurrency did achieve was speed of overall development and the assurance that all systems would be available on schedule.70

A further tool in WDD’s management portfolio was parallel de­velopment. To avoid being dependent on a single supplier for a sys – 36 tem, Schriever insisted on parallel contractors for many of them.

Chapter 1 Eventually, when Thor and Titan I came along, the testing program became overwhelming, and Glasser argued that Ramo-Wooldridge just ignored the problem. He went to Schriever, who directed him to come up with a solution. He decided which systems would go on

Atlas, which on Titan and Thor, in the process becoming the deputy for systems management and the Atlas project manager.71

A final component of the management structure for Schriever’s west-coast operation consisted of the nonprofit Aerospace Corpora­tion. It had come into existence on June 4, 1960, as a solution to the problems many people saw in Space Technology Laboratories’ serv­ing as a systems-engineering and technical-direction contractor to the air force while part of Thompson Ramo Wooldridge (later, just TRW), as the company had become following an eventual merger of Ramo-Wooldridge with Thompson Products. STL continued its operations for programs then in existence, but many of its person­nel transferred to the Aerospace Corporation for systems engineer­ing and technical direction of new programs. Further complicat­ing the picture, a reorganization occurred within the air force on April 1, 1961, in which Air Force Systems Command (AFSC) re­placed the Air Research and Development Command. On the same date, within AFSC, the Ballistic Missile Division split into a Ballis­tic Systems Division (BSD), which would retain responsibility for ballistic missiles (and would soon move to Norton Air Force Base [AFB] east of Los Angeles near San Bernardino); and a Space Systems Division (SSD), which moved to El Segundo, much closer to Los Angeles, and obtained responsibility for military space systems and boosters. There would be further reorganizations of the two offices, but whether combined or separated, they oversaw the development of a variety of missiles and launch vehicles, ranging from the Atlas and Thor to Titans I through IV.72

To return specifically to the Atlas program, under the earlier (1946-48) MX-774B project, Convair had developed swiveling of en­gines (a precursor of gimballing); monocoque propellant tanks that were integral to the structure of the rockets and pressurized with nitrogen to provide structural strength with very little weight pen­alty (later evolving into what Convair called a steel balloon); and separable nose cones so that the missile itself did not have to travel with a warhead to the target and thus have to survive the aerody­namic heating from reentering the atmosphere.73

Подпись: 37 German and U.S. Missiles and Rockets, 1926-66 Other innovations followed under the genial leadership of Karel (Charlie) Bossart. Finally, on January 6, 1955, the air force awarded a contract to Convair for the development and production of the Atlas airframe, the integration of other subsystems with the airframe and one another, their assembly and testing. The contractor for the At­las engines was North American Aviation, which built upon earlier research done on the Navaho missile. NAA’s Rocketdyne Division, formed in 1955 to handle the requirements of Navaho, Atlas, and

Redstone, developed one sustainer and two outside booster engines for the Atlas under a so-called stage-and-a-half arrangement, with the boosters discarded after they had done their work. Produced in

1957 and 1958, the early engines ran into failures of systems and components in flight testing that also plagued the Thor and Jupiter engines, which were under simultaneous development and shared many component designs with the Atlas.74

But innovation continued, partly through engineers making “the right guess or assumption" or simply learning from problems. De­spite repeated failures and (trial-and-error) modifications to elimi­nate their causes, development proceeded from Atlas A through At­las F with a total of 158 successful launches for all models against 69 failures—a success rate of only 69.6 percent. The Atlas D became the first operational version in September 1959, with the first E and F models following in 1961. All three remained operational until 1965, when they were phased out of the missile inventory, with many of them later becoming launch-vehicle stages.75

Meanwhile, fearing (unnecessarily) that an ICBM like the Atlas could not be deployed before 1962, a Technology Capabilities Panel headed by James R. Killian Jr., president of MIT, issued a report in mid-February 1955 recommending the development of both sea – and land-based intermediate-range ballistic missiles (IRBMs). In Novem­ber 1955, the Joint Chiefs of Staff recommended, in turn, that the air force develop the land-based version while the army and navy collaborate on an IRBM that could be both land and sea based. Thus were born the air force’s Thor and the army’s Jupiter, with the navy eventually developing the solid-propellant Polaris after initially try­ing to adapt the liquid-propellant Jupiter to shipboard use.76

Arising out of this decision was the “Thor-Jupiter Controversy," which the House of Representatives Committee on Government Operations called a “case study in interservice rivalry." The Thor did not use the extremely light, steel-balloon structure of Atlas but a more conventional aluminum airframe. Its main engine consisted essentially of half of the booster system for Atlas. In 1957 and 1958, it experienced 12 failures or partial successes out of the first 18 launches. Before the air force nevertheless decided in September

1958 that Thor was ready for operational deployment, problems with the turbopumps (common to the Atlas, Thor, and Jupiter) and

38 differences of approach to these problems had led to disagreement Chapter 1 between the Thor and Jupiter teams.77

Von Braun’s engineers, working on the Jupiter for the army, di­agnosed the problem first and had Rocketdyne design a bearing re­tainer for the turbopump that solved the problem, which the Thor

program would not admit at first, suspecting another cause. Once the Jupiters resumed test flights, they had no further turbopump problems. Meanwhile, failures of an Atlas and a Thor missile in April 1958, plus subsequent analysis, led the air force belatedly to accept the army’s diagnosis and a turbopump redesign. The first Thor squadron went on operational alert in Great Britain in June 1959, with three others following by April 1960. When the Atlas and Titan ICBMs achieved operational readiness in 1960, the last Thors could be removed from operational status in 1963, making them available for space-launch activities.78

While the Western Development Division and the successor Air Force Ballistic Missile Division were developing the Thor in con­junction with contractors, von Braun’s group at what had become the Army Ballistic Missile Agency (ABMA) in Alabama and its con­tractors were busily at work on Jupiter without a clear indication whether the army or the air force would eventually deploy the mis­sile. At ABMA, the forceful and dynamic Maj. Gen. John B. Medaris enjoyed powers of initiative roughly analogous to those of Schriever for the air force. On December 8, 1956, the navy left the Jupiter program to develop Polaris, but not before the sea service’s require­ments had altered the shape of the army missile to a much shorter and somewhat thicker contour than the army had planned. With Chrysler the prime contractor (as on the Redstone), Medaris reluc­tantly accepted the same basic engine North American Rocketdyne was developing for the Thor except that the Jupiter engine evolved from an earlier version of the powerplant and ended as somewhat less powerful than the air force counterparts.79

Подпись: 39 German and U.S. Missiles and Rockets, 1926-66 With a quite different vernier engine and guidance/control sys­tem, the Jupiter was a decidedly distinct missile from the Thor. The first actual Jupiter (as distinguished from the Jupiter A and Jupiter C, which were actually Redstones) launched on March 1, 1957, at Cape Canaveral. Facing the usual developmental problems, including at least one that Medaris blamed on the thicker shape resulting from the navy’s requirements, the Jupiter nevertheless achieved 22 sat­isfactory research-and-development flight tests out of 29 attempts. The air force, instead of the army, deployed the missile, with initial operational capability coming on October 20, 1960. Two squadrons of the missile became fully operational in Italy as of June 20, 1961. A third squadron in Turkey was not operational until 1962, with all of the missiles taken out of service in April of the following year. Three feet shorter, slightly thicker and heavier, the Jupiter was more accurate but less powerful than the Thor, with a comparable range. The greater average thrust of the Thor may have contributed

to its becoming a standard first-stage launch vehicle, whereas Jupi­ter served in that capacity to only a limited degree. Another factor may have been that there were 160 production Thors to only 60 Ju­piter missiles.80

Although much has rightly been made of the intense interservice rivalry between the army and the air force over Thor and Jupiter, even those two programs cooperated to a considerable extent and exchanged much data. Medaris complained about the lack of infor­mation he received from the air force, but Schriever claimed that his Ballistic Missile Division had transmitted to ABMA a total of 4,476 documents between 1954 and February 1959. By his count, BMD withheld only 28 documents for a variety of reasons, including con­tractors’ proprietary information.81 This was one of many examples showing that—although interservice and interagency rivalry helped encourage competing engineers to excel—without sharing of infor­mation and technology, rocketry might have advanced much less quickly than in fact it did.

Double-Base Propellants during and Soon after World War II

At the beginning of recent solid-propellant development during World War II, the vast majority of rockets produced for use in com­bat employed extruded double-base propellants. These were limited in size by the nature of the extrusion process used at that time to produce them. In extrusion using a solvent, nitrocellulose was sus­pended in the solvent, which caused the nitrocellulose to swell. It was formed into a doughlike composition and then extruded (forced) through dies to form it into grains. This process of production lim­ited the size of the grains to thin sections so the solvent could evap­orate, and the elasticity of the grain was too low for bonding large charges to the motor case. With a solventless (or dry) process, there 232 were also limitations on the size of the grain and greater hazards of Chapter 6 explosion than with extrusion using a solvent.

These factors created the need for castable double-base propel­lants. But before a truly viable process for producing large castable propellants could be developed, the United States, because it was at war, needed a variety of rockets to attack such targets as ships (including submarines), enemy fortifications, gun emplacements, aircraft, tanks, and logistical systems. The development of these weapons did not lead directly to any launch-vehicle technology, but the organizations that developed them later played a role in furthering that technology. Two individuals provided the leader­ship in producing the comparatively small wartime rockets with extruded grains. One was Clarence Hickman, who had worked with Goddard on rockets intended for military applications during World War I. He then earned a Ph. D. at Clark University and went to work at Bell Telephone Laboratories. After consulting with God­dard, in June 1940 Hickman submitted a series of rocket proposals to Frank B. Jewett, president of Bell Labs and chairman of a divi­sion in the recently created National Defense Research Committee (NDRC). The upshot was the creation of Section H (for Hickman) of the NDRC’s Division of Armor and Ordnance. Hickman’s section had responsibility for researching and developing rocket ordnance. Although Section H was initially located at the Naval Proving Ground at Dahlgren, Virginia, it worked largely for the army.

Hickman chose to use wet-extruded, double-base propellants (em­ploying a solvent) because he favored the shorter burning times they afforded compared with dry-extruded ones. He and his associates worked with this propellant at Dahlgren, moved to the Navy Pow­der Factory at Indian Head, Maryland, and finally to Allegany Ord­nance Plant, Pinto Branch, on the West Virginia side of the Potomac

River west of Cumberland, Maryland. There at the end of 1943 they set up Allegany Ballistics Laboratory, a rocket-development facility operated for Section H by George Washington University. By using traps, cages, and other devices to hold the solvent-extruded, double­base propellant, they helped develop the bazooka antitank weapon, a 4.5-inch aircraft rocket, JATO devices with less smoke than those produced by Aerojet using Parson’s asphalt-based propellant, and a recoilless gun.22 Under different management, ABL later became an important producer of upper stages for missiles and rockets.

Подпись:Hickman’s counterpart on the West Coast was physics professor Charles Lauritsen of Caltech. Lauritsen was vice chairman of the Division of Armor and Ordnance (Division A), and in that capacity he had made an extended trip to England to observe rocket devel­opments there. The English had developed a way to make solvent­less, double-base propellant by dry extrusion. This yielded a thicker grain that would burn longer than the wet-extruded propellant but required extremely heavy presses for the extrusion. However, the benefits were higher propellant loading and the longer burning time that Lauritsen preferred.

Convinced of the superiority of this kind of extrusion and be­lieving that the United States needed a larger rocket program than Section H could provide with its limited facilities, Lauritsen argued successfully for a West Coast program. Caltech then set up opera­tions in Eaton Canyon in the foothills of the San Gabriel Moun­tains northeast of the campus in Pasadena. It operated from 1942 to 1945 and expanded to a 3,000-person effort involving research, development, and pilot production of rocket motors; development of fuses and warheads; and static and flight testing. The group pro­duced an antisubmarine rocket 7.2 inches in diameter, a 4.5-inch barrage rocket, several retro-rockets (fired from the rear of airplanes at submarines), 3.5- and 5-inch forward-firing aircraft rockets, and the 11.75-inch “Tiny Tim" rocket that produced 30,000 pounds of thrust and weighed 1,385 pounds. (This last item later served as a booster for the WAC Corporal.)

By contrast with Section H, Section L (for Lauritsen) served mainly the navy’s requirements. In need of a place to test and eval­uate the rockets being developed at Eaton Canyon, in November 1943 the navy established the Naval Ordnance Test Station (NOTS) in the sparsely populated desert region around Inyokern well north of the San Gabriel Mountains. Like the Allegany Ballistics Labora­tory, NOTS was destined to play a significant role in the history of U. S. rocketry, mostly with tactical rockets but also with contribu­tions to ballistic missiles and launch vehicles.

One early contribution was the “White Whizzer" 5.0-inch rocket developed by members of the Caltech team who had already moved to NOTS but were still under direction of the university rather than the navy. By about January1944, combustion instability had become a problem with the 2.25-inch motors for some of the tactical rock­ets. These rockets used tubular, partially internal-burning charges of double-base propellant. Radial holes in the grain helped solve se­vere pressure excursions—it was thought, by allowing the gas from the burning propellant to escape from the internal cavity. Edward W. Price, who had not yet received his bachelor’s degree but would later become one of the nation’s leading experts in combustion in­stability, suggested creating a star-shaped perforation in the grain for internal burning. He thought this might do a better job than the 234 radial holes in preventing oscillatory gas flow that was causing the Chapter 6 charges of propellant to split. He tested the star perforation, and it did produce stable burning.

In 1946, Price applied this technique to the White Whizzer, which featured a star-perforated, internal-burning grain with the outside of the charge wrapped in plastic to inhibit burning there. This ge­ometry allowed higher loading of propellant (the previous design having channels for gas flow both inside and outside the grain). And since the grain itself protected the case from the heat in the internal cavity, the case could be made of lightweight aluminum, providing better performance than heavier cases that were slower to accelerate because of the additional weight. Ground-launched about May 1946, the White Whizzer yielded a speed of 3,200 feet per second, then a record for solid rockets. The internal-burning, aluminum-cased design features later appeared in the 5.0-inch Zuni and Sidewinder tactical missiles. The internal-burning feature of the design also came to be applied to a great many other solid rock­ets, including ballistic missiles and stages for launch vehicles. This apparently was the first flight of a rocket using such a grain design in the United States, preceding JPL’s use of a similar design, known as the Deacon, and also flight testing of the first member of the Vicar family to be flown.23

Propulsion for the A-4 (V-2)

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

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

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

Подпись: 103 Propulsion with Alcohol and Kerosene Fuels, 1932-72 Dornberger said Thiel assumed “complete charge of propul­sion, with the aim of creating a 25-ton motor" (the one used for the A-4, providing 25 metric tons of thrust). Because Thiel remained at Kummersdorf until 1940 instead of moving to Peenemunde with the rest of the von Braun group in 1937, testing facilities limited him to engines of no more than 8,000 pounds of thrust from 1936 to 1940. Although Thiel was “extremely hard-working, conscientious, and systematic," Dornberger said he was difficult to work with. Ambitious and aware of his abilities, he “took a superior attitude and demanded. . . devotion to duty from his colleagues [equal to his own]." This caused friction that Dornberger claimed he had to mol­lify. Martin Schilling (chief of the testing laboratory at Peenemunde for Thiel’s propulsion development office and, later, head of the of­fice after Thiel died in a bombing raid in 1943) noted that Thiel was “high strung." He said, “Thiel was a good manager of such a great and risky development program. He was a competent and dynamic leader, and a pusher. At the same time, he was no match to von Braun’s or Steinhoff’s vision and optimism." (Ernst Steinhoff was chief of guidance and control.) 3

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

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

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

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

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

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

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

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

Подпись: 105 Propulsion with Alcohol and Kerosene Fuels, 1932-72 In conjunction with the shortening of the combustion chamber, Thiel also converted the shape from cylindrical to spherical to en­compass the greatest volume in available space. This also served to reduce pressure fluctuations and increase the mixing of the propel­lants. Until he could use a larger test stand at Peenemunde, how­ever, Thiel was restricted in scaling up these innovations in the 1.5-ton engine to the full 25 tons. He thus went to an intermediate size of 4.2 tons that he could test at Kummersdorf, and he moved from one injector in the smaller engine to three in the larger one. Each had its own “mixing compartment" or “pot," and the clus­tering actually increased the efficiency of combustion further. But to go from that arrangement to one for the 25-ton engine created considerable problems of scaling up and of arranging the 18 “pots" that the researchers designed for the A-4 combustion chamber. At first, Thiel and his associates favored an arrangement of six or eight larger injectors around the sides of the chamber, but von Braun sug-

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

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

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

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

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

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

Подпись: 107 Propulsion with Alcohol and Kerosene Fuels, 1932-72 The problems with the pumps included warping of the pump housing because of the temperature difference between the steam and the liquid oxygen; cavitation because of bubbles in the propel­lants; difficulties with lubrication of the bearings; and problems with seals, gaskets, and choice of alloys (all problems that would recur in later U. S. missiles and rockets). The cavitation problem was especially severe since it could lead to vibrations in the com­bustion chamber, resulting in explosions. The solution came from redesigning the interior of the pumps and carefully regulating the internal pressure in the propellant tanks to preclude the formation of the bubbles.11

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

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

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

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

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

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

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

Подпись: 109 Propulsion with Alcohol and Kerosene Fuels, 1932-72 One final innovation, of undetermined origin, involved the igni­tion process, which used a pyrotechnic igniter. In the first step of the process, the oxygen valve opened by means of an electrically activated servo system, followed by the alcohol valve. Both opened to about 20 percent of capacity, but since the propellants flowed only as a result of gravity (and slight pressure in the oxygen tank), the flow was only about 10 percent of normal. When lit by the ig­niter, the burning propellants produced a thrust of about 2.5 tons. When this stage of ignition occurred, the launch team started the turbopump by opening a valve permitting air pressure to flow to the hydrogen peroxide and sodium permanganate tanks. The perman­ganate solution flowed into a mixing chamber, and as soon as pres­sure was sufficient, a switch opened the peroxide valve, allowing peroxide to enter the mixing chamber. When pressure was up to 33 atmospheres as a result of decomposing the hydrogen peroxide, the oxygen and alcohol valves opened fully, and the pressure on the tur­bines in the pumps caused them to operate, feeding the propellants into the combustion chamber. It required only about three-quarters of a second from the time the valve in the peroxide system was elec­trically triggered until the missile left the ground.17

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

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

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

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

Propulsion for the A-4 (V-2)
Подпись: Walt Disney (left), with his hand on a model of the V-2 rocket, and Wernher von Braun in 1954 during a period in which von Braun worked with Disney Studios to promote spaceflight on television, an example of his heterogeneous engineering. (Photo courtesy of NASA)

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

Подпись: 111 Propulsion with Alcohol and Kerosene Fuels, 1932-72 Under difficult, wartime conditions, in-house contributions and those from technical institutes and industry came together through discussions among the contributors at Kummersdorf and Peene – munde. The pooling of their expertise probably contributed in innu­merable ways to the progress of technological development, but the

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

Transtage

An important step forward occurred with the third liquid-propellant stage for the Titan III, known as Transtage, for which the air force decided on a pressure-fed engine that would use the same nitrogen tetroxide as oxidizer and Aerozine 50 for fuel as stages one and two. As planned, it would have two gimballed thrust chambers, each pro­ducing 8,000 pounds of thrust, and a capability of up to three starts over a six-hour period. Aerojet won this contract, with a Phase I agreement signed in early 1962 and a Phase II (development) award issued on January 14, 1963.51

Aerojet designed the Transtage engine (designated AJ10-138) at about the same time as a larger propulsion unit for the Apollo service module. The two engines used basically the same design, featuring the same propellants, ablatively cooled thrust chambers, and a radiatively cooled nozzle assembly. Since the Apollo service module’s engine bore the designation AJ10-137, its development ap­parently began earlier in 1962, but it also lasted longer. Although

Aerojet designed and built them both, and more information is available about the development of the spacecraft engine, it is not clear that any of the latter’s problems and solutions are relevant to the Transtage engine, which was less than half as powerful and roughly half the length and diameter of its sibling.52

Apparently, these two engines were not the only ones with ab­latively cooled combustion chambers in this period, because an important NASA publication on liquid-propellant rocket engines issued in 1967 stated that such “thrust chambers have many ad­vantages for upper-stage applications. They are designed to meet accumulated duration requirements varying from a few seconds to many minutes." Although construction could vary, in one ex­ample (unspecified), the ablative liner used a high-silica fabric im­pregnated with phenolic resin and then tape-wrapped on a mandrel. Asbestos impregnated with phenolic served as an insulator on the outer surface of the liner. A strong outer shell consisted of layers of one-directional glass cloth to provide longitudinal strength. Cir­cumferential glass filaments “bonded with epoxy resin" provided “hoop strength."53 This appears to have described the Transtage combustion chamber (as well as others?).54

Подпись:On July 23, 1963, Aerojet had successfully operated a Transtage engine for 4 minutes, 44 seconds, considered “a long duration fir­ing." During that static test, the engine started and stopped three times, demonstrating the restart capability. However, a more criti­cal test of this crucial capability (which would allow it to place multiple satellites into different orbits on a single launch or to po­sition a single satellite in a final orbit, such as a geostationary or­bit, without a need for a separate kick motor) would occur in the simulated-altitude test chamber at Arnold Engineering Develop­ment Center in Tullahoma, Tennessee. In August 1963, tests at that center confirmed suspicions from the July 23 test that the combus­tion chamber would burn through before completing a full-duration firing (undefined). In addition, gimballing of the engine in a cold environment revealed a malfunction of a bipropellant valve (that fed propellants to the combustion chamber) and a weakness in the nozzle extension, made of aluminide-coated columbium and radia­tion cooled with an expansion ratio of 40:1. Information about how Aerojet solved these problems is not available in any of the sources for this book, with the official history of the Titan III merely stating that “by the close of 1963, an extensive redesign and testing pro­gram was underway to eliminate these difficulties so the contractor could make his first delivery of flight engine hardware—due in mid – December 1963." 55

One Aerojet source does not comment on these particular diffi­culties but does refer to “the error of trying to develop in a produc­tion atmosphere." The source explained in this connection that de­velopment of this small engine occurred while Titan I was starting into production, causing management and engineers/technicians to pay less attention to it. But presumably, the speed required in Transtage’s development was also a factor in these particular prob­lems. Obviously, engineers had not expected them and had to adjust designs to correct the difficulties. In any event, engine deliveries did not occur in mid-December, as initially planned, but started in April 1964.

Aerojet engineer and manager Ray Stiff recalled that after engine deliveries began, the air force started to impose new requirements. Because Transtage needed to perform a 6.5-hour coast while in orbit and then be capable of “a variety of firing, coast, and refire combi­nations," there had to be “unique insulation requirements," to pro­tect propellants from freezing in the extreme cold of orbit in space, especially when shaded from the sun. But this insulation retained the heat from combustion, which built up around the injector with presumed dire consequences for continued performance. Stiff does not reveal how Aerojet solved this problem, stating only that the engine’s injector was “baffled for assurance of stable combustion."56

Other sources reveal that the injector used an “all-aluminum flat 170 faced design" with a “concave spherical face, [and] multiple-orifice Chapter 4 impinging patterns." The baffle was fuel cooled, so perhaps an ad­justment in this feature solved the heating problem. According to an Aerojet history written by former employees and managers, “The injector design has undergone two performance upgrade programs which resulted in the very high specific impulse value of 320 lbf-sec/ lbm, and the design has been carried over into later versions of the Delta."57 (Most sources do not rate the specific impulse this high.)

In any event, the two initial Transtage engines each yielded 8,000 pounds (lbf) of thrust with a specific impulse of more than 300 lbf-sec/lbm. Pressurized by cold helium gas, each of the hyper – golic propellants was stored in tanks of a titanium alloy that the prime contractor, Martin, machined in its Baltimore Division. The titanium forgings came from the Ladish Company of Cudahy, Wis­consin. Although titanium was difficult to machine, it was gaining increasing use for liquid-propellant tanks. With a fuel tank about 4 feet in diameter by 13.5 feet in length and an oxidizer tank mea­suring about 5 by 11 feet, Transtage’s propellant containers were hardly huge but were reportedly some of the largest yet produced from titanium. Each overall engine was 6.8 feet long with its diam-

eter ranging from 25.2 to 48.2 inches. Its rated burning time was a robust 500 seconds, and its total weight was only 238 pounds.58

Transtage advanced storable-propellant technology but also rep­resented a further example of trial-and-error engineering. Other up­per stages used the technology developed for Transtage and for the Apollo service module’s engine.

Polaris and Minuteman

Jupiter, Thor, and Atlas marked a huge step forward in the matura­tion of U. S. rocketry, but before the technology from those missiles came to significant use in launch vehicles, the navy’s development of the Polaris inaugurated a solid-propellant breakthrough in mis­sile technology that also profoundly affected launch vehicles.82 Un­til Polaris A1 became operational in 1960, all intermediate-range and intercontinental missiles in the U. S. arsenal had employed liq­uid propellants. These had important advantages in terms of perfor­mance but required extensive plumbing and large propellant tanks that made protecting them in silos difficult and expensive. Such factors also virtually precluded their efficient use onboard ships, especially submarines. Once Minuteman I became operational in 1962, the U. S. military began to phase out liquid-propellant stra­tegic missiles. To this day, Minuteman III and the solid-propellant fleet ballistic missiles continue to play a major role in the nation’s strategic defenses because they are simpler and cheaper to operate than liquid-propellant missiles.

Because of the higher performance of some liquid propellants and 40 their ability to be throttled as well as turned off and on by the use of Chapter 1 valves, they remained the primary propellants for space-launch ve­hicles. However, since solid-propellant boosters could be strapped on the sides of liquid-propellant stages for an instant addition of high thrust (because their thrust-to-weight ratio is higher, allowing

faster liftoff), solid-propellant boosters became important parts of launch-vehicle technology. The technologies used on Polaris and Minuteman transferred to such boosters and also to upper stages of rockets used to launch satellites. Thus, the solid-propellant break­through in missiles had important implications for launch-vehicle technology. By the time that Polaris got under way in 1956 and Minuteman in 1958, solid-propellant rocketry had already made tremendous strides from the use of extruded double-base propel­lants in World War II tactical missiles. But there were still enor­mous technical hurdles to overcome before solid-propellant missiles could hope to launch strategic nuclear warheads far and accurately enough to serve effectively as a deterrent or as a retaliatory weapon in case of enemy aggression.83

With a much smaller organization than the army or air force, a navy special projects office under the leadership of Capt. (soon-to-be Rear Adm.) William F. Raborn pushed ahead to find the right tech­nologies for a submarine-launched, solid-propellant missile, a daunt­ing task since a solid propellant with the necessary performance did not yet exist. Capt. Levering Smith—who, at the Naval Ord­nance Test Station (NOTS), had led the effort to develop a 50-foot solid-propellant missile named “Big Stoop" that flew 20 miles in 1951—joined Raborn’s special projects office in April 1956. Smith contributed importantly to Polaris, but one key technical discovery came from the Atlantic Research Corporation (ARC), a chemical firm founded in 1949 with which the Navy Bureau of Ordnance had contracted to improve the specific impulse of solid propellants (the ratio of thrust a rocket engine or motor produced to the amount of propellant needed to produce that thrust).84

ARC’s discovery that the addition of comparatively large quan­tities of aluminum to solid propellants significantly raised perfor­mance, together with the work of Aerojet chemists, led to successful propellants for both stages of Polaris A1. The addition of aluminum to Aerojet’s binder essentially solved the problem of performance for both Polaris (and, as it turned out, with a different binder, for Minuteman). Other key technical solutions relating to guidance and an appropriate warhead led to the directive on December 8, 1956, that formally began the Polaris program.85

Подпись: 41 German and U.S. Missiles and Rockets, 1926-66 Flight testing of Polaris at the air force’s Cape Canaveral (be­ginning in 1958 in a series designated AX) revealed a number of problems. Solutions required considerable interservice cooperation. On July 20, 1960, the USS George Washington launched the first functional Polaris missile. The fleet then deployed the missile on November 15, 1960.86

The navy quickly moved forward to Polaris A2. It increased the range of the fleet ballistic missile from 1,200 to 1,500 miles. Flight testing of the A2 missiles started in November 1960, with the first successful launch from a submerged submarine occurring on Octo­ber 23, 1961. The missile became operational less than a year later in June 1962. Polaris A3 was still more capable, with a range of 2,500 miles. It incorporated many other new technologies in both propulsion and guidance/control, becoming operational on Septem­ber 28, 1964. All three versions of Polaris made significant contri­butions to launch-vehicle technology, such as the Altair II motor, produced by the Hercules Powder Company under sponsorship of the Bureau of Naval Weapons and NASA and used as a fourth stage for the Scout launch vehicle.87

While Polaris was still in development, the air force had officially begun work on Minuteman I. Its principal architect was Edward N. Hall, a heterogeneous engineer who helped begin the air force’s involvement with solid propellants as a major at Wright-Patterson AFB in the early 1950s. As Karl Klager, who worked on both Polaris and Minuteman, has stated, Hall “deserves most of the credit for maintaining interest in large solid rocket technology [during the mid-1950s] because of the greater simplicity of solid systems over liquid systems." Hall’s efforts “contributed substantially to the Polaris program," Klager added, further illustrating the extent to which (unintended) interservice cooperation and shared informa­tion contributed to the solid-propellant breakthrough. Hall moved to the WDD as the chief for propulsion development in the liq­uid-propellant Atlas, Titan, and Thor programs, but he continued his work on solids, aided by his former colleagues back at Wright – Patterson AFB.88

Despite this sort of preparatory work for Minuteman, the mis­sile could not begin its formal development until the air force se­cured final DoD approval in February 1958, more than a year later than Polaris. Hall and others at WDD had a difficult job convincing Schriever in particular to convert to solids. Without their heteroge­neous engineering, the shift to solids might never have happened. They were aided, however, by development of Polaris because it provided what Harvey Sapolsky has dubbed “competitive pressure" for the air force to develop its own solid-propellant missile.89

Soon after program approval, Hall left the Ballistic Missile Divi­sion. From August 1959 to 1963, the program director was Col. (soon promoted to Brig. Gen.) Samuel C. Phillips. Hall and his coworkers deserve much credit for the design of Minuteman and its support by the air force, whereas Phillips brought the missile to completion.

Facing many technical hurdles, Phillips succeeded as brilliantly as had Levering Smith with Polaris in providing technical manage­ment of a complex and innovative missile. Often using trial-and – error engineering, his team working on the three-stage Minuteman I overcame problems with materials for nozzle throats in the lower stages, with firing the missile from a silo, and with a new binder for the first stage called polybutadiene-acrylic acid-acrylonitrile (PBAN), developed by the contractor, Thiokol Chemical Corpora­tion. Incorporating substantial new technology as well as some bor­rowed from Polaris, the first Minuteman I wing became operational in October 1962.90

Minuteman II included a new propellant in stage two, known as carboxy-terminated polybutadiene and an improved guidance/con – trol system. The new propellant yielded a higher specific impulse, and other changes (including increased length and diameter) made Minuteman II a more capable and accurate missile than Minute – man I. The newer version gradually replaced its predecessor in mis­sile silos after December 1966.91

In Minuteman III, stages one and two did not change from Min – uteman II, but stage three became larger. Aerojet replaced Hercules as the contractor for the new third stage. With the larger size and a different propellant, the third stage more than doubled its total impulse. These and other modifications allowed Minuteman IIIs to achieve their initial operational capability in June 1970. As a result of the improvements, the range of the missile increased from about 6,000 miles for Minuteman I to 7,021 for Minuteman II, and 8,083 for Minuteman III.92

Подпись: 43 German and U.S. Missiles and Rockets, 1926-66 The deployment of Minuteman I in 1961 marked the completion of the solid-propellant breakthrough in terms of its basic technol­ogy, though innovations and improvements continued to occur. But the gradual phaseout of liquid-propellant missiles followed almost inexorably from the appearance on the scene of the first Minuteman. The breakthrough in solid-rocket technology required the extensive cooperation of a great many firms, government laboratories, and uni­versities, only some of which could be mentioned here. It occurred on many fronts, ranging from materials science and metallurgy through chemistry to the physics of internal ballistics and the mathematics and physics of guidance and control, among many other disciplines. It was partially spurred by interservice rivalries for roles and mis­sions. Less well known, however, was the contribution of interser­vice cooperation. Necessary funding for advances in and the sharing of technology came from all three services, the Advanced Research Projects Agency, and NASA. Technologies such as aluminum fuel,

methods of thrust vector control, and improved guidance and con­trol transferred from one service’s missiles to another. Also crucial were the roles of heterogeneous engineers like Raborn, Schriever, and Hall. But a great many people with more purely technical skills, such as Levering Smith and Sam Phillips, ARC, Thiokol, and Aerojet engineers made vital contributions.

The solid-propellant breakthrough that these people and many others achieved had important implications for launch vehicles as well as missiles. The propellants for the large solid-rocket boosters on the Titan III, Titan IVA, and the Space Shuttles were derived from the one used on Minuteman, stage one. Without ARC’s dis­covery of aluminum as a fuel and Thiokol’s development of PBAN as a binder, it is not clear that the huge Titan and shuttle boosters would have been possible. Many other solid-propellant formula­tions also used aluminum and other ingredients of the Polaris and Minuteman motors. Although some or all of them might have been developed even if there had been no urgent national need for solid – propellant missiles, it seems highly unlikely that their development would have occurred as quickly as it did without the impetus of the cold-war missile programs and their generous funding.

CASTABLE DOUBLE-BASE PROPELLANT

The next major development in double-base propellants was a method for casting (rather than extruding) the grain. The company that produced the first known rocket motor using this procedure was the Hercules Powder Company, which had operated the govern­ment-owned Allegany Ballistics Laboratory since the end of World War II. The firm came into existence in 1912 when an antitrust suit

against its parent company, E. I. du Pont de Nemours & Company, forced du Pont to divest some of its holdings. Hercules began as an explosives firm that produced more than 50,000 tons of smokeless powder during World War I. It then began to diversify into other uses of nitrocellulose. During World War II, the firm supplied large quantities of extruded double-base propellants for tactical rockets. After the war, it began casting double-based propellants by beginning with a casting powder consisting of nitrocellulose, nitroglycerin, and a stabilizer. Chemists poured this into a mold and added a cast­ing solvent of nitroglycerin plus a diluent and the stabilizer. With heat and the passage of time, this yielded a much larger grain than could be produced by extrusion alone.

Подпись:Wartime research by John F. Kincaid and Henry M. Shuey at the National Defense Research Committee’s Explosives Research Laboratory at Bruceton, Pennsylvania (operated by the Bureau of Mines and the Carnegie Institute of Technology), had yielded this process. Kincaid and Shuey, as well as other propellant chemists, had developed it further after transferring to ABL, and under Hercu­les management, ABL continued work on cast double-base propel­lants. This led to the flight testing of a JATO using this propellant in 1947. The process allowed Hercules to produce a propellant grain that was as large as the castable, composite propellants that Aero­jet, Thiokol, and Grand Central were developing in this period but with a slightly higher specific impulse (also with a greater danger of exploding rather than burning and releasing the exhaust gases at a controlled rate).24

The navy had contracted with Hercules for a motor to be used as an alternative third stage on Vanguard (designated JATO X241 A1). The propellant that Hercules’ ABL initially used for the motor was a cast double-base formulation with insulation material between it and the case. This yielded a specific impulse of about 250 lbf-sec/ lbm, higher both than Grand Central’s propellant for its Vanguard third-stage motor and the specification of 245 lbf-sec/lbm for both motors. A key feature of the motor was its case and nozzle, made of laminated fiberglass. ABL had subcontracted work on the case and nozzle to Young Development Laboratories, which developed a method during 1956 of wrapping threads of fiberglass soaked in epoxy resin around a liner made of phenolic asbestos. (A phenol is a compound used in making resins to provide laminated coatings or form adhesives.) Following curing, this process yielded a strong, rigid Spiralloy (fiberglass) shell with a strength-to-weight ratio 20 percent higher than the stainless steel Aerojet was using for its propellant tanks on stage two of Vanguard.25

In 1958, while its third-stage motor was still under development, Hercules acquired this fiberglass-winding firm. Richard E. Young, a test pilot who had worked for the M. W. Kellogg Company on the Manhattan Project, had founded it. In 1947, Kellogg had designed a winding machine under navy contract, leading to a laboratory in New Jersey that built a fiberglass nozzle. It moved to Rocky Hill, New Jersey, in 1948. There, Young set up the development labora­tories under his own name and sought to develop lighter materials for rocket motors. He and the firm evolved from nozzles to cases, seeking to improve a rocket’s mass fraction (the mass of the propel­lant divided by the total mass of a stage or rocket), which was as important as specific impulse in achieving high velocities. In the mid-1950s, ABL succeeded in testing small rockets and missiles us – 236 ing cases made with Young’s Spiralloy material.26

Chapter 6 This combination of a cast double-base propellant and the fiber­glass case and nozzle created a lot of problems for Hercules engi­neers. By February 1957, ABL had performed static tests on about 20 motors, 15 of which resulted in failures of insulation or joints. Combustion instability became a problem on about a third of the tests. Attempting to reduce the instability, Hercules installed a plastic paddle in the combustion zone to interrupt the acoustic pat­terns (resonance) that caused the problem. This did not work as well as hoped, so the engineers developed a suppressor of thicker plastic. They also improved the bond between insulator and case, then cast the propellant in the case instead of just sliding it in as a single piece. Nine cases still failed during hydrostatic tests or static firings. The culprits were high stress at joints and “severe combus­tion instability."27

In February 1958, ABL began developing a follow-on third-stage motor designated X248 A2 in addition to X241. Perhaps it did so in part to reduce combustion instability, because 3 percent of the propellant in the new motor consisted of aluminum, which burned in the motor and produced particles in the combustion gases that suppressed (damped) high-frequency instabilities. But another moti­vation was increased thrust. The new motor was the one that actu­ally flew on the final Vanguard mission, September 18, 1959. As of August 1958, ABL had developed a modification of this motor, X248 A3, for use as the upper stage in a Thor-Able lunar probe. By this time, ABL was testing the motors in an altitude chamber at the air force’s Arnold Engineering Development Center and was experienc­ing problems with ignition and with burnthroughs of the case the last few seconds of the static tests.28

The X248 solid-rocket motor consisted of an epoxy-fiberglass case filled with the case-bonded propellant. The nozzle was still made of epoxy fiberglass, but with a coating of "ceramo-asbestos." By November 11, 1958, wind-tunnel static tests had shown that the X248 A2 filament-wound exit cone was adequate. By this time also, the motor had a sea-level theoretical specific impulse of about 235, which extrapolated to an impulse at altitude of some 255 lbf-sec/ lbm, and designers had overcome the other problems with the mo­tor. The X248 offered a "considerable improvement in reliability and performance over the X241 contracted for originally," according to Kurt Stehling. He also said the ABL version of the third stage suc­cessfully launched the Vanguard III satellite weighing 50 pounds, whereas Grand Central Rocket’s third stage could orbit only about 30 pounds.29