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

Simultaneously with the development of Polaris and then Minute – man, the air force continued work on two liquid-propellant missiles, the Titans I and II. The Titan II introduced storable propellants into the missile inventory and laid the groundwork for the core portion of the Titans III and IV space-launch vehicles. Titan I began as es­sentially insurance for Atlas in case the earlier missile’s technology proved unworkable. The major new feature of the first of the Titans was demonstration of the ability to start a large second-stage engine at a high altitude.93 The WAC Corporal had proved the viability of the basic process involved, and Vanguard would develop it further (after Titan I was started). But in 1955, using a full second stage on a ballistic missile and igniting it only after the first-stage engines had exhausted their propellants seemed risky.

The air force approved development of Titan I on May 2, 1955. Meanwhile, the Western Development Division had awarded a 44 contract on January 14, 1955, to Aerojet for engines burning liq – Chapter 1 uid oxygen and a hydrocarbon fuel for possible use on Atlas. These soon evolved into engines for the two-stage missile. Even though the Aerojet engines burned the same propellants as Atlas, there were problems with development, showing that rocket engineers

still did not have the process of design “down to a science." Despite the change in propellants, the Titan II used a highly similar design for its engines, making Aerojet’s development for that missile less problematic than it might otherwise have been (although still not without difficulties), with technology then carrying over into the Titans III and IV core launch vehicles. Meanwhile, the air force de­ployed the Titan Is in 1962. They quickly deactivated in 1965 with the deployment of Minuteman I and Titan II, but Titan I did provide an interim deterrent force.94

The history of the transition from Titan I to Titan II is compli­cated. One major factor stimulating the change was the 15 minutes or so it took to raise Titan I from its silo, load the propellants, and launch it. Another was the difficulty of handling Titan I’s extremely cold liquid oxygen used in Titan I inside a missile silo. One solution to the twin problems would have been conversion to solid propel­lants like those used in Polaris and Minuteman, but another was storable propellants. Under a navy contract in 1951, Aerojet had begun studying hydrazine as a rocket propellant. It had good perfor­mance but could detonate. Aerojet came up with a compromise so­lution, an equal mixture of hydrazine and unsymmetrical dimethyl hydrazine, which it called Aerozine 50. With nitrogen tetroxide as an oxidizer, this fuel mixture ignited hypergolically (upon contact with the oxidizer, without the need for an ignition device), offering a much quicker response time than for Titan I.95 As a result of this and other issues and developments, in November 1959 the Depart­ment of Defense authorized the air force to develop the Titan II. The new missile would use storable propellants, in-silo launch, and an all-inertial guidance system.96

On April 30, 1960, the Air Force Ballistic Missile Division’s de­velopment plan for Titan II called for it to be 103 feet long (compared to 97.4 feet for Titan I), have a uniform diameter of 10 feet (whereas Titan I’s second stage was only 8 feet across), and have increased thrust over its predecessor. This higher performance would increase the range with the Mark 4 reentry vehicle from about 5,500 nauti­cal miles for Titan I to 8,400. With the new Mark 6 reentry vehicle, which had about twice the weight and more than twice the yield of the Mark 4, the range would remain about 5,500 nautical miles. Because of the larger nuclear warhead it could carry, the Titan II served a different and complementary function to Minuteman I’s in the strategy of the air force, convincing Congress to fund them both. It was a credible counterforce weapon, whereas Minuteman I served primarily as a countercity missile, offering deterrence rather than the ability to destroy enemy weapons in silos.97

In May 1960, the air force signed a letter contract with the Mar­tin Company to develop, produce, and test the Titan II. It followed this with a contract to General Electric to design the Mark 6 reentry vehicle. In April 1959, AC Spark Plug had contracted to build an in­ertial guidance system for a Titan missile, although it was not clear at the time that this would be the Titan II.98

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 new designs did not always work as anticipated, the engineers had to resort to empirical solu­tions until they found the combinations that provided the necessary performance. Even with other changes to the Titan I engine designs, the Titan II propulsion system had significantly fewer parts than its Titan I predecessor, reducing chances for failure during operation. Despite the greater simplicity, the engines had higher thrust and higher performance, as planned.99

Flight testing of the Titan II had its problems, complicated by plans to use the missile as a launch vehicle for NASA’s Project Gem­ini, leading to the Project Apollo Moon flights. However, the last 13 flights in the research-and-development series were successful, giving the air force the confidence to declare the missile fully opera­tional on the final day of 1963. Between October and December 1963, the Strategic Air Command deployed six squadrons of nine Titan IIs apiece. They remained a part of the strategic defense of the United States until deactivated between 1984 and 1987. By that time, fleet ballistic missiles and smaller land-based, solid-propellant ballistic missiles could deliver (admittedly smaller) warheads much more accurately than could the Titan IIs. Deactivation left the former operational Titan II missiles available for refurbishment as space – launch vehicles.100

Development of Titan I and Titan II did not require a lot of new technology. Instead, it adapted technologies developed either ear­lier or simultaneously for other missile or launch-vehicle programs. Nevertheless, the process of adaptation for the designs of the two Titan missiles generated problems requiring engineers to use their fund of knowledge to find solutions. These did work, and Titan II became the nation’s longest-lasting liquid-propellant missile with the greatest throw weight of any vehicle in the U. S. inventory.

It would appear that the Caltech group had not made great progress on liquid propulsion as of July 1940 when it was joined by Malina’s roommate, Martin Summerfield, who completed work on his Ph. D. during 1941 in the physics department at Caltech. After joining Malina’s group, he went to the Caltech library, consulted the litera­ture on combustion-chamber physics, and found a text with infor­mation on the speed of combustion. Using it, he calculated—much in the fashion of Thiel at Kummersdorf—that the combustion chamber being used by the GALCIT team was far too large, result­ing in heat transfer that degraded performance. So he constructed a smaller chamber of cylindrical shape that yielded a 20 percent increase in performance. Von Karman believed that roughly 25 to 30 percent of the heat in the combustion chamber would be lost, based on information about reciprocating engines. The eminent aerodynamicist had therefore concluded that it would be impossi­ble for rocket engines to be self-cooling, restricting both their light­ness and length of operation. Summerfield’s calculations showed 146 these assumptions about heat transfer to be far too high, indicating Chapter 4 that it was possible for a self-cooling engine to operate for a sus­tained period. Subsequent tests confirmed Summerfield’s calcula­tions, and Malina learned about the technique of regenerative cool­ing from James H. Wyld of Reaction Motors during one of his trips back East.2

For the moment, the group worked with uncooled engines burn­ing RFNA and gasoline. Successive engines of 200, 500, and 1,000 pounds of thrust with various numbers of injectors provided some successes but presented problems with throbbing or incomplete ini­tial ignition, which led to explosions. After four months of efforts to improve combustion and ignition, Malina paid a visit to the Na­val Engineering Experiment Station in Annapolis in February 1942. There he learned that chemical engineer Ray C. Stiff had discovered in the literature of chemistry that aniline ignited hypergolically with nitric acid. Malina telegrammed Summerfield to replace the gasoline with aniline. He did so, but it took three different injector designs to make the 1,000-pound engine work. The third involved eight sets of injectors each for the two propellants, with the stream of pro­pellants washing against the chamber walls. Summerfield recalled

that after 25 seconds of operation, the heavy JATO units glowed cherry red. But they worked on a Douglas A-20A bomber for 44 suc­cessive firings in April 1942, the first successful operation of a liq­uid JATO in the United States. This led to orders by the army air forces (AAF) with the newly formed Aerojet Engineering Corpora­tion, which Malina and von Karman had helped to found.3

Aerojet did considerable business with the AAF and navy for JATO units during the war and had become by 1950 the largest rocket – engine manufacturer in the world, as well as a leader in research and development of rocket technology. Until Aerojet’s acquisition by General Tire in 1944-45, the rocket firm and the GALCIT rocket project maintained close technical relations. Although GALCIT/JPL was involved essentially with JATO work from 1939 to 1944, in the summer of 1942 the project began designing pumps to deliver liquid propellants to a combustion chamber instead of feeding the propel­lants by gas under pressure. By the fall of that year, project engineers were working on using the propellants to cool the combustion cham­ber of a 200-pound-thrust engine.4

Meanwhile, as a result of Stiff’s discovery, Truax’s group at Annapolis began using 1,500-pound JATOs burning nitric acid and aniline on navy PBY aircraft in 1943. Both Truax and Stiff subse­quently got orders to work at Aerojet, where Stiff devoted his efforts to a droppable JATO using storable, hypergolic propellants. Aerojet produced about 100 of these units, and some came to be used by the U. S. Coast Guard. In these ways, Aerojet became familiar with use of storable propellants, and Stiff joined the firm after completing his obligatory service with the navy.5

The next important development involving engines with stor­able propellants was the WAC Corporal sounding rocket (the term WAC standing for Women’s Auxiliary Corps or Without Attitude Control, depending upon the source consulted). The Army Ord­nance Corps had requested that Malina’s project investigate the fea­sibility of developing a rocket carrying meteorological equipment that could reach a minimum altitude of 100,000 feet. The JPL team redesigned an Aerojet motor that used monoethylene as a fuel and nitric acid mixed with oleum as an oxidizer. The original motor was regeneratively cooled by the monoethylene. JPL adapted the motor to use RFNA containing 6.5 percent nitrogen dioxide as oxidizer and aniline containing 20 percent furfuryl alcohol as a fuel, thereby increasing the exhaust velocity from 5,600 to 6,200 feet per second but leaving the thrust at 1,500 pounds for 45 seconds. According to one source, the specific impulse was 200 lbf-sec/lbm (slightly lower than the V-2).6

Besides exceeding the requirements of the army, the small, liquid-propellant rocket also functioned as a smaller test version of the Corporal E research vehicle, providing valuable experience in the development of that larger unit. During the testing, the pro­gram decided to modify the WAC Corporal to attain higher alti­tudes. A substantial modification of the engine reduced its weight from 50 to 12 pounds. The WAC A initial version of the rocket had a comparatively thin, cylindrical inner shell of steel for the combus­tion chamber, with an outer shell that fit tightly around it but was equipped with a joint to permit expansion. Helical coils (ones that spiraled around the outside of the combustion chamber like a screw thread) provided regenerative cooling, with a shower-type injector in which eight fuel streams impinged on eight oxidizer streams. For the modified WAC B engine, designers reduced the combustion chamber in length from 73 to 61 inches and made minor modifica­tions to the injector. It had an inner shell spot-welded to the outer shell, still with helical cooling passages. The injector remained a showerhead with eight pairs of impinging jets.7

In a series of flight tests at White Sands Proving Ground, New Mexico, in December 1946, none of the WAC Corporal B vehicles rose more than 175,000 feet in altitude. Apparently the test team suspected cavitation (gas bubbles) in the injector system as the cause of the less-than-optimal performance, since team members 148 constructed three more B-model vehicles with orifice inserts that Chapter 4 were screwed in, rather than drilled as before, to achieve cavitation – free injection of the propellants into the combustion chamber. In three February-March 1947 tests, one WAC Corporal B reached an altitude of 240,000 feet. Overall, the WAC Corporal demonstrated that the propulsion system was sound and the nitric acid-aniline – furfuryl alcohol propellant combination was viable.8

The WAC Corporal led directly to the successful Aerobee sound­ing rocket built by Aerojet, which was used by the Applied Physics Laboratory of Johns Hopkins University for research in the upper atmosphere. Then, in the Bumper-WAC project, the WAC Corporal B flew as a second stage on V-2 missiles. The reported altitude of 244 miles and maximum speed of 7,553 feet per second (reached on February 24, 1949) were records. This highly successful launch demonstrated that a rocket’s velocity could be increased with a sec­ond stage and that ignition of a rocket engine could occur at high altitudes.9

In addition, the engine for the WAC Corporal contributed to the Corporal missile’s propulsion system. As first conceived, Corporal E was a research vehicle for the study of guidance, aerodynamic, and

propulsion problems of long-range rockets. In 1944, von Karman estimated that a rocket with a range of 30 to 40 miles would be necessary to serve as a prototype for a later missile. He thought such a vehicle would need an engine with 20,000 pounds of thrust and 60 seconds of burning time. Experience at JPL to that point had indicated that the only already-developed rocket type meeting von Karman’s specifications would be a liquid-propellant vehicle burn­ing red fuming nitric acid and aniline. Early plans called for use of centrifugal, turbine-driven pumps to feed the propellants. Since Aerojet had a turborocket under development, JPL thought it could draw on the nearby rocket firm’s experience to provide a pump for the Corporal. This design became the never-completed Corporal F. Corporal E used air pressurization, as had the WAC Corporal.10

Scaling the WAC Corporal engine up to a larger size proved chal­lenging. The first major design for a Corporal E engine involved a 650-pound, mild-steel version with helical cooling passages. Such a heavy propulsion device resulted from four unsuccessful attempts to scale up the WAC Corporal B engine to 200 pounds. None of them passed their proof testing. In the 650-pound engine, the cool­ing passages were machined to a heavy outer shell that formed a sort of hourglass shape around the throat of the nozzle. The injector consisted of 80 pairs of impinging jets that dispersed the oxidizer (fuming nitric acid) onto the fuel. The direction, velocity, and di­ameter of the streams were similar to those employed in the WAC Corporal A. The injector face was a showerhead type with orifices more or less uniformly distributed over it. It mixed the propellants in a ratio of 2.65 parts of oxidizer to 1 of fuel. The outer shell of the combustion chamber was attached to an inner shell by silver solder. When several of these heavyweight engines underwent proof test­ing, they cracked and nozzle throats eroded as the burning propel­lant exhausted out the rear of the engine. But three engines with the inner and outer shells welded together proved suitable for flight testing.11

On May 22, 1947, the first Corporal E with this heavyweight engine launched from the army’s White Sands Proving Ground. Its intended range was 60 miles, and it actually achieved a range of 62.5 (in one account, 64.25) miles. The second launch occurred on July 17, 1947, but the rocket failed to achieve enough thrust to rise significantly until 90 seconds of burning reduced the weight to the point that it flew a very short distance. On November 4, 1947, the third launch was more successful, but its propellants burned for only 43 (instead of 60) seconds before the engine quit. This reduced its range to just over 14 miles. Both it and the “rabbit killer" (the

second vehicle, so-called because it flew along the ground) expe­rienced burnthroughs in the throat area, the helical cooling coils proving inadequate for their purpose.12

Deciding that in addition to these flaws, the engine was too heavy, the Corporal team determined to design a much lighter-weight en­gine. Several engines combining features of the WAC Corporal B and 650-pound Corporal E combustion chambers all suffered burnouts of the throat area during static tests. Finally, a redesigned engine weighing about 125 pounds stemmed in part from an examination of the V-2, revealing that its cooling passages were axial (with no helix angle, i. e., they took the shortest distance around the combustion chamber’s circumference). Analysis showed the advantage of that ar­rangement, so JPL adopted it. The inner shell of the new engine was corrugated, and the outer shell, smooth. The shape of the combus­tion chamber changed from semispherical to essentially cylindrical, with the inside diameter reduced from 23 to 11 inches and the length shortened slightly, contributing to the much lighter weight.

It took two designs to achieve a satisfactory injector, the first having burned through on its initial static test. The second injector had 52 pairs of impinging jets angled about 2.5 degrees in the direc­tion of (but located well away from) the chamber wall. Initially, the Corporal team retained the mixture ratio of 2.65:1. But static tests of the axially cooled engine in November 1948 at the Ordnance – 150 California Institute of Technology (ORDCIT) Test Station in Muroc, Chapter 4 California (in the Mojave Desert above the San Gabriel Mountains and well north of JPL), showed that lower mixture ratios yielded higher characteristic velocities and specific impulses, as well as smoother operation. Thus, the mixture ratio was first reduced to 2.45 and then 2.2. Later still, the propellant was changed to stabi­lized fuming nitric acid (including a very small amount of hydrogen fluoride) as the oxidizer and aniline-furfuryl alcohol-hydrazine (in the percentages of 46.5, 46.5, and 7.0, respectively) as the fuel. With this propellant, the mixture ratio shifted further downward to 2.13 because of changes in the densities of the propellants. The resul­tant engine, made of mild steel, provided high reliability. Its suc­cess rested primarily upon its “unique configuration, wherein the cool, uncorrugated outer shell carrie[d] the chamber pressure loads, and the thin inner shell, corrugated to form forty-four axial cooling passages, [wa]s copper-brazed to the outer shell." Finally, the inside of the inner shell (the combustion chamber inside face) was plated with chrome to resist corrosion from the propellants.13

The sixth Corporal E launch took place on November 2, 1950. The missile experienced multiple failures. It landed 35.9 miles

downrange, about 35 miles short of projections. Later static tests revealed problems with a propellant regulator that had caused over­rich mixture ratios on both the fifth and sixth launches. Failure of a coupling had resulted in loss of air pressure. The radar beacon to provide overriding guidance in azimuth operated satisfactorily until failure of a flight-beacon transmitter some 36 seconds into the flight. The Doppler beacon never went into operation to cut off propellant flow at the proper moment because the missile failed to achieve the velocity prescribed, but also because the Doppler bea­con itself failed at 24 seconds after liftoff. As a final blow, all elec­tronic equipment failed, apparently from extreme vibration.14

On launch seven of the Corporal E in January 1951, the vehicle landed downrange at 63.85 miles, 5 miles short of the targeted im­pact point. This was the first flight to demonstrate propellant shut­off and also the first to use a new multicell air tank and a new air – disconnect coupling. These two design changes to fix some of the problems on launch six increased the reliability of the propulsion system significantly. However, although the Corporal performed even better on launch eight (March 22, 1951), hitting about 4 miles short of the target, on launch nine (July 12, 1951) the missile landed 20 miles beyond the target because of failure of the Doppler tran­sponder and the propellant cutoff system. The final “round" of Cor­poral E never flew. But the Corporal team had learned from the first nine rounds how little it understood about the flight environment of the vehicle, especially vibrations that occurred when it was op­erating. The team began to use vibration test tables to make the design better able to function and to test individual components before installation. This testing resulted in changes of suppliers and individual parts as well as to repairs before launch (or redesigns in the case of multiple failures of a given component.)15

The next 20 Corporals, with the airframes built by Douglas Air­craft (like the Corporal Es), received the designation Corporal I. Its first flight occurred on October 10, 1951. But the frequency regulator for the central power supply failed on takeoff, causing the missile to follow nearly a vertical trajectory. Range safety cut its flight short so that it would impact between White Sands and the city of Las Cruces, New Mexico. Flight 11 (referred to as round 12, counting the last Corporal E, which never launched) occurred on December 6, 1951. Before the launch, the army invited several companies to bid on production contracts as prime contractors. Ryan Aeronautical Company of San Diego manufactured the engines for both the air­frames built by Douglas and those from the new prime contractor, Firestone Tire and Rubber Company of Los Angeles. JPL received

the Firestone missiles and disassembled them for inspection. It then rebuilt them and performed preflight testing before sending them to White Sands for the actual flight tests. Then it sent comments to the manufacturer to help improve factory production. According to Clayton Koppes, however, the two major contractors and JPL failed to work together effectively. Meanwhile, between January and De­cember 1952, JPL launched 26 Corporals, including the first 10 of the Firestone lot as well as 16 produced by Douglas.16

Because of problems with the missile’s guidance system and engi­neering changes to correct them, a second production order to Fire­stone for Corporal missiles in late 1954 resulted in a redesignation of the missile as Corporal II. JPL retained technical control of the Corporal program throughout 1955, relinquishing it in 1956 while continuing to provide technical assistance to the army’s contrac­tors, including Firestone. Corporal II continued to have problems with its guidance/control system but also with propellant shutoff during firings of the missile by army field forces. Fact-finding in­vestigations and informal discussions on the parts of contractors, the field forces, Army Ordnance Corps, and JPL led to greater care by field forces personnel in following operational procedures. These eliminated shutoff problems when not violated. The army declared the Corporal to be operational in 1954, and in January 1955 the Corporal I deployed to Europe. Eight Corporal II battalions replaced 152 it during 1956 and the first half of 1957.17

Chapter 4 Although the Corporal was less powerful and had a shorter range than the V-2, the U. S. missile’s propulsion system had a higher spe­cific impulse (about 220 lbf-sec/lbm as compared with 210 for the V-2). In some respects, such as the axial nature of the cooling sys­tem and the use of Doppler radar for propellant cutoff, the Corporal had borrowed from the V-2. In most respects, however, the Ameri­can missile was an independent development, in some cases one that separately adopted features developed at Peenemunde after it was too late to incorporate them into the V-2. These included a showerhead injector and the use of hypergolic propellants. Both had been developed for the Wasserfall antiaircraft rocket, and a single injector plate later became a standard element in the construction of the rockets designed in Huntsville.18

Among the achievements of the Corporal was testing the effects of vibration on electronic equipment. The vibration tables used for this purpose may have been the first effective simulators of the flight environment in that area. Subsequently, both testing for the effects of vibrations and analysis of components and systems for reliability became standard practice in missile development.19

The engine itself was also a notable achievement. Although the idea for the axial direction of cooling flow came from the V-2, the overall engine was certainly original. It was both light and efficient, and even though there seems to be no evidence that its design influ­enced subsequent engines, it seems likely that propulsion engineers learned something of their art from it. Moreover, the early work of JPL in hypergolics transferred to Aerojet, later the contractor for the Titan II, which used storable liquid propellants that ignited on contact. This technology was also used in the Titan III and Titan IV liquid rockets, which employed direct descendants of the early hy – pergolic propellants Malina learned about in part from the navy in Annapolis. This was a significant contribution from both indige­nous U. S. research efforts during World War II. It illustrates one of the ways that technology transferred from one program to another in American rocketry. The borrowings from the V-2 exemplify a dif­ferent pattern of information flow.

Meanwhile, the first major application of the technologies devel­oped for the RV-A-10 was the Sergeant missile, for which JPL began planning in 1953 under its ORDCIT contract with Army Ordnance. JPL submitted a proposal for a Sergeant missile in April 1954, and on June 11, 1954, the army’s chief of ordnance programmed $100,000 for it. At the same time, he transferred control of the effort to the com­manding general of Redstone Arsenal. Using lessons learned from the liquid-propellant Corporal missile, JPL proposed a co-contractor for the development and ultimate manufacture of the missile. In February 1956, a Sergeant Contractor Selection Committee unani­mously chose Sperry Gyroscope Company for this role, based on JPL’s recommendation and Sperry’s capabilities and experience with other missiles, including the Sparrow I air-to-air missile system for the navy. In April 1954, the Redstone Arsenal had reached an agreement with the Redstone Division of Thiokol to work on the solid-propellant motor for the Sergeant, with the overall program to develop Sergeant beginning in 1955.30

There is no need to provide a detailed history of the Sergeant missile here. It took longer to develop than originally planned and was not operational until 1962. By then the navy had completed the far more significant Polaris A1, and the air force was close to field­ing the much more important Minuteman I. The Sergeant did meet a slipped ordnance support readiness date of June 1962 and became a limited-production weapons system until June 1968. It did equal its predecessor, Corporal, in range and firepower in a package only

FIG. 6.3

Technical drawing of the Jupiter C (actually,

Juno I, including scaled-down Sergeant upper stages) with America’s first satellite, Explorer I, showing the latter’s characteristics. (Photo courtesy of NASA)

half as large and requiring less than a third as much ground-support equipment. Its solid-propellant motor could also be readied for fir­ing much more quickly than the liquid-propellant Corporal.31

The Sergeant motor was a modification or direct descendant of the RV-A-10’s motor. The latter (using the TRX-110A propulsion formulation) employed 63 percent ammonium perchlorate as an oxidizer, whereas the TP-E8057 propellant for the Sergeant motor (designated JPL 500) had 63.3 percent of that oxidizer and 33.2 per­cent LP-33 liquid polymer in addition to small percentages of a curing agent, two reinforcing agents, and a curing accelerator. At a nozzle expansion ratio of 5.39, its specific impulse was about 185 lbf-sec/lbm, considerably lower than the performance of Polaris A1. It employed a five-point-star grain configuration, used a case of 4130 steel at a nominal thickness of 0.109 inch (almost half that of the RV-A-10 case), and a nozzle (like that of the RV-A-10) using 1020 steel with a graphite nozzle-throat insert.32 Ironically, perhaps, the main contributions the Sergeant made to launch-vehicle tech­nology were through a scaled-down version of the missile used for testing. These smaller versions became the basis for upper stages in reentry test vehicles for the Jupiter missile and in the launch vehicles for Explorer and Pioneer satellites.33

The development of missiles and rockets for DoD needs arguably contributed to national defense and, through deterrence, kept the

cold war from becoming hotter than it actually got in Korea, Viet­nam, and Afghanistan, among other places. For the purposes of this book, however, the importance of the missiles and rockets discussed in this chapter lay in the technology that could transfer to launch – vehicle uses. In many cases, actual missiles, with some adaptations, became either launch vehicles or stages in larger combinations of rockets used to place satellites or spacecraft on their trajectories. Without the perceived urgency created by cold-war concerns and without the heterogeneous engineering of missile proponents, it conceivably would have taken much longer for launch vehicles to develop, although many satellites themselves were high on the DoD’s priority lists.

Quite apart from their contributions to launch-vehicle technol­ogy, the missiles and rockets discussed in this chapter also illustrate many of the themes that will be further explored in subsequent chapters. Missiles such as the Titan II and Minuteman showed the ways in which technology for earlier missiles contributed to their successors. Although this chapter provides only an overview of mis­sile development, it shows several examples of trial-and-error engi­neering that was necessary to overcome often unforeseen problems. Clearly, the missiles discussed here required a wide range of talents and a huge number of different organizations to design and develop them. Also important was a considerable sharing of information, even between competing organizations and firms. Finally, manage­ment systems such as the one Schriever adopted at WDD (and a similar system called Program Evaluation and Review Technique [PERT] adopted by Raborn for the Polaris program) enabled very complicated missiles and launch vehicles to be developed reason­ably on time and in such a way that all component systems (such as propulsion, structures, guidance and control) worked together effectively.

LAUNCH VEHICLES FREQUENTLY USED MIS­siles as first stages, but these required many modi­fications, particularly when they had to boost hu­mans into space. Even for satellite and spacecraft launches, technology for the booster stages fre­quently represented modification of technologies missiles needed for their ballistic paths from one part of Earth to another. Thus, the history of the Thor-Delta, Atlas, Scout, Saturn, Titan, and Space Shuttle launch vehicles differed from, but remained

dependent on, the earlier development of the missiles discussed in chapter 1. Missiles and launch vehicles represented a continuum, with many of the same people contributing to both. But they re­mained different enough from one another to require separate treat­ment in this chapter.

Despite the differences, launch-vehicle development exhibited many of the same themes that characterized missiles. It featured the same engineering culture that relied heavily on extensive test­ing on the ground. But this did not always succeed in revealing all problems that occurred in flight. When unexpected problems oc­curred, it was not always possible for engineers to understand the exact causes. But they were able to arrive at fixes that worked. There continued to be a wide range of organizations and disciplines that contributed to launch-vehicle development, including the solution of unanticipated problems. Also characteristic of launch vehicles was a competitive environment that nevertheless featured sharing of information among organizations involved in development. In part, this sharing occurred through the movement of knowledge­able engineers from one organization to another. More often, the information sharing (plus its recording and validation) occurred through professional societies, papers delivered at their meetings, and publication of reports in professional journals.1 Finally, mis­siles and launch vehicles shared the use of management systems that tracked development of components to ensure that all of them occurred on schedule and that they all worked together effectively.

Soon after the army deployed Corporal I, Aerojet had occasion to develop its storable-propellant technology further with stage two of the Vanguard launch vehicle for the navy. The firm’s Aerobee sounding rockets, building on the WAC Corporal engine technology, had led to the Aerobee-Hi sounding rocket that provided the basis for the projected stage two. As requirements for that stage became more stringent, though, Aerobee-Hi proved deficient, and Aerojet had to return to the drawing board. The firm charged with designing Vanguard, the Martin Company, contracted with Aerojet on Novem­ber 14, 1955, to develop the second-stage engine. Martin had deter­mined that the second stage needed a thrust of 7,500 pounds and a specific impulse at altitude of 278 lbf-sec/lbm to provide the required velocity to lift the estimated weight of the Vanguard satellite.20

The development of an engine to meet these specifications proved to be difficult. Martin’s calculated thrust and specific im­pulse would not meet the vehicle’s velocity requirements without severe weight limitations. Aerojet engineers selected unsymmetri­cal dimethyl hydrazine (UDMH) and inhibited white fuming ni­tric acid (IWFNA) as the propellants because they were hypergolic (eliminating problems with ignition), had a high loading density (reducing the size, hence weight, of propellant tanks), and delivered the requisite performance. Another advantage of hydrazine and acid was a comparative lack of problems with combustion instability in experimental research.21

The history of the evolution from the aniline-nitric acid propel­lants used in the WAC Corporal (specifically, red fuming nitric acid with 6.5 percent nitrogen dioxide plus aniline with the addition of 20 percent furfuryl alcohol) and in the first Aerobee sounding rocket (35 instead of 20 percent furfuryl alcohol) to the UDMH and IWFNA used in Vanguard is complicated. But it illustrates much about pro­pellant chemistry and the number of institutions contributing to it. The basic aniline-RFNA combination worked as a self-igniting pro­pellant combination. But it had numerous disadvantages. Aniline is highly toxic and rapidly absorbed via the skin. A person who came into contact with a significant amount of it was likely to die rapidly from cyanosis. Moreover, aniline has a high freezing point, so it can be used only in moderate temperatures. RFNA is highly corrosive to propellant tanks, so it has to be loaded into a missile or rocket just before firing, and when poured, it gives off dense concentra­tions of nitrogen dioxide, which is also poisonous. The acid itself burns the skin, as well. Two chemists at JPL had discovered as early as 1946 that white fuming nitric acid (WFNA) and furfuryl alcohol with aniline were just as poisonous and corrosive but did not pro­duce nitrogen dioxide.

But WFNA turned out to be inherently unstable over time. A complicated substance, it was hard for propellant chemists to ana­lyze in the early 1950s. By 1954, however, researchers at the Na – 154 val Ordnance Test Station and at JPL had thoroughly investigated Chapter 4 nitrogen tetroxide and nitric acid and come up with conclusions that were to be used in the Titan II. Meanwhile, chemists at the Naval Air Rocket Test Station, Lake Denmark, New Jersey; JPL; the NACA’s Lewis Flight Propulsion Laboratory; the air force’s Wright Air Development Center in Dayton, Ohio; and Ohio State Univer­sity, among other places, had reached a fundamental understanding of nitric acid by 1951 and published the information by 1955. In the process, the Naval Air Rocket Test Station was apparently the first to discover that small percentages of hydrofluoric acid both reduced the freezing point of RFNA/WFNA and inhibited corrosion with many metals. Thus were born inhibited RFNA and WFNA, for which the services and industrial representatives under air force sponsorship drew up military specifications in 1954. In this way, the services, the NACA, one university, and the competing indus­tries cooperated to solve a common problem.

During the same period, chemists sought either replacements for aniline or chemicals to mix with it and make it less problematic. Hydrazine seemed a promising candidate, and in 1951 the Rocket Branch of the navy’s Bureau of Aeronautics, issued contracts to

Metallectro Company and Aerojet to see if any hydrazine deriva­tives were suitable as rocket propellants. They found that UDMH rapidly self-ignited with nitric acid, leading to a military specifica­tion for UDMH in 1955.22

Despite the severe weight limitations on the second-stage en­gine, the Vanguard project engineers had decided to use a pressure – fed (rather than a pump-fed) propellant-delivery system. The pumps produced angular momentum as they rotated, and for stage two, this would be hard for the roll-control system to overcome, especially after engine cutoff. Concerns about reliability led to a decision to use heated helium gas as the pressurant in the feed system. Aerojet convinced the Martin Company and the navy to use stainless-steel instead of aluminum propellant tanks. Because steel had a better strength-to-weight ratio than aluminum, Aerojet argued that the lighter metal would, paradoxically, have had to weigh 30 pounds more than the steel to handle the pressure.

Moreover, a “unique design for the tankage" placed the sphere containing the helium pressure tank between the two propellant tanks, serving as a dividing bulkhead and saving the weight of a sepa­rate bulkhead. A solid-propellant gas generator augmented the pres­sure of the helium and added its own chemical energy to the system at a low cost in weight. Initially, Aerojet had built the combustion chamber of steel. It accumulated 600 seconds of burning without corrosion, but it was too heavy. So engineers developed a lightweight chamber made up of aluminum regenerative-cooling, spaghetti-type tubes wrapped in stainless steel. It weighed 20 pounds less than the steel version, apparently the first such chamber built of aluminum tubes for use with nitric acid and UDMH.23

During 1956 there were problems with welding the stainless – steel tanks despite Aerojet’s experience in this area. Martin recom­mended a different method of inspection and improvements in tool­ing, which resolved these problems. The California firm also had to try several types of injector before finding the right combination of features. One with 72 pairs of impinging jets did not deliver suffi­cient exhaust velocity, so Aerojet engineers added 24 nonimpinging orifices for fuel in the center portion of the injector. This raised the exhaust velocity above the specifications but suggested the empiri­cal nature of the design process, with engineers having to test one design before discovering that it would not deliver the desired per­formance. They then had to use their accumulated knowledge and insights to figure out what modification might work.24

The development of the combustion chamber and related equip­ment illustrated the same process. Despite the use of inhibited

white fuming nitric acid, the lightweight aluminum combustion chamber—which could be lifted with one hand—gradually eroded. It took engineers “weeks of experimenting" to find out that a coating of tungsten carbide substantially improved the life of the combus­tion chamber. There also were problems with the design of valves for flow control, requiring significant modifications.25

A final problem lay in testing an engine for start at altitude. At the beginning of the project, there was no vacuum chamber large enough to test the engine, but according to NRL propulsion engineer Kurt Stehling, “Several tests were [eventually?] made at Aerojet with engine starts in a vacuum chamber." In any event, to preclude prob­lems with near-vacuum pressure at altitude, the engineers sealed the chamber with a “nozzle closure" that kept pressure in the cham­ber until exhaust from ignition blew it out.26

The original Vanguard schedule as of November 1955 called for six test vehicles to be launched between September 1956 and Au­gust 1957, with the first satellite-launching vehicle to lift off in Oc­tober 1957.27 It was not until March 1958, however, that the second stage could be fired in an actual launch—that of Test Vehicle (TV) 4. TV-4 contained modifications introduced into the stage-one engine following the failure of TV-3 (when stage one exploded), but it did not yet incorporate the tungsten-carbide coating in the aluminum combustion chamber of the stage-two engine. And it was still a test 156 vehicle. On March 17, 1958, the slender Vanguard launch vehicle Chapter 4 lifted off. It performed well enough (despite a rough start) to place the small 3.4-pound Vanguard I satellite in an orbit originally esti­mated to last for 2,000 (but later revised to 240) years.28

On TV-5, launched April 28, 1958, the second stage provided less-than-normal thrust, but the first stage had performed better than normal, compensating in advance for the subpar second stage. Then an electrical problem prevented ignition of the third stage, pre­cluding orbit. On the first nontest Vanguard, Space Launch Vehicle (SLV) 1, apparent malfunction of a pressure switch also prevented orbiting a 21.5-pound satellite on May 27, 1958. Here, the second stage performed normally through cutoff of ignition. On SLV-2, June 26, 1958, the second-stage engine cut off after eight seconds, probably due to clogged filters in the inhibited white fuming nitric acid lines from corrosion of the oxidizer tank. The Vanguard team flushed the oxidizer tanks and launched SLV-3 on September 26, 1958, with a 23.3-pound satellite. Despite the flushing, second – stage performance was below normal, causing the satellite to miss orbital speed by a narrow margin. This time, the problem seemed to be a clogged fuel (rather than oxidizer) filter.

On February 17, 1959, however, all systems worked, and SLV-4 placed the 23.3-pound Vanguard II satellite in a precise orbit ex­pected to last for 200 years or more. This did not mean that Aerojet had gotten all of the kinks out of the troublesome second stage. SLV-5 on April 13, 1959, experienced a flame oscillation during sec­ond-stage ignition, apparently producing a violent yaw that caused the second and third stages with the satellite to tumble and fall into the ocean. Engineers made changes in the second-stage engine’s hy­draulic system and programmed an earlier separation of the stage, but on SLV-6 (June 22, 1959), a previously reliable regulating valve ceased to function after second-stage ignition. This caused helium pressure to mount (since it could not vent), resulting in an explo­sion that sent the vehicle into the Atlantic about 300 miles down – range. At least the problem-plagued Vanguard program ended on a happy note. On September 18, 1959, a test vehicle backup (TV-4BU) version of the launch vehicle placed a 52.25-pound X-ray and envi­ronmental satellite into orbit.29

Meanwhile, the navy’s Polaris missile had made more far-reaching contributions. Until Polaris A1 became operational in 1960, all U. S.

long-range missiles had used liquid propellants. These had obvious advantages in their performance, but their extensive plumbing and large propellant tanks made protecting them in silos difficult and costly. Such factors also made them impractical for use on ships. After the operational date of Minuteman I in 1962, the Department of Defense began phasing out liquid-propellant strategic missiles.34

Meanwhile, given the advantages that liquid propellants en­joyed in terms of performance, their head start within the defense establishment, and the disinclination of most defenders of liquids to entertain the possibility that solid propellants could satisfy the demanding requirements of the strategic mission, how did this solid-propellant breakthrough occur? The answer is complicated and technical. But fundamentally, it happened because a number of heterogeneous engineers promoted solids; a variety of partners in their development brought about significant technical innovations; and although interservice rivalries encouraged the three services to development separate missiles, interservice cooperation ironically helped them do so. Despite such cooperation and the accumulat­ing knowledge about rocket technology, however, missile designers still could not foresee all the problems that their vehicles would develop during ground and flight testing. Thus, when problems did occur, rocket engineers still had to gather information about what had caused problems and exercise their ingenuity to develop solu­tions that would cope with the unexpected.

By the time that Polaris got under way in 1956 and Minuteman in 1958, solid-propellant rocketry had already made the tremen­dous strides forward discussed previously. But there were still enor­mous technical hurdles to overcome. The problems remaining to be solved included higher performance; unstable combustion; the inadequate durability of existing nozzle materials under conditions of heat and exposure to corrosive chemicals from the exhaust of the burning propellants; a lack of materials and technology to provide light but large combustion chambers so the burning propellants had to overcome less mass during launch; and ways to terminate com­bustion of the propellant immediately after the desired velocity had been achieved (for purposes of accuracy) and to control the direction of the thrust (for steering).35

Once the navy had overcome the bureaucratic obstacles to devel­oping its own, solid-propellant missile, the Special Project Office (SPO) under Adm. William F. Raborn and Capt. Levering Smith achieved breakthroughs in a number of these technical areas. In early January 1956, the navy had sought the assistance of the Lock­heed Missile and Space Division and the Aerojet General Corpora-

tion in developing a solid-propellant ballistic missile. The initial missile the two contractors and the SPO conceived was the Jupiter S (for “solid"). It had enough thrust to carry an atomic warhead the re­quired distance, a feat it would achieve by clustering six solid rock­ets in a first stage and adding one for the second stage. The problem was that Jupiter S would be about 44 feet long and 10 feet in diam­eter. An 8,500-ton vessel could carry only 4 of them but could carry 16 of the later Polaris missiles. With Polaris not yet developed, the navy and contractors still were dissatisfied with Jupiter S and con­tinued to seek an alternative.36

One contribution to a better solution came from Atlantic Re­search Corporation (ARC). Keith Rumbel and Charles B. Hender­son, chemical engineers with degrees from MIT who were working 240 for ARC, had begun theoretical studies in 1954 of how to increase Chapter 6 solid-propellant performance. They learned that other engineers, in­cluding some from Aerojet, had calculated an increase in specific impulse from adding aluminum powder to existing ingredients. But these calculations had indicated that once aluminum exceeded 5 per­cent of propellant mass, performance would again decline. Hence, basing their calculations on contemporary theory and doing the cumbersome mathematics without the aid of computers, the other researchers abandoned aluminum as an additive except for damping combustion instability. Refusing to be deterred by theory, Rumbel and Henderson tested polyvinyl chloride with much more alumi­num in it. They found that with additional oxygen in the propellant and a flame temperature of at least 2,310 kelvin, a large percent­age of aluminum by weight yielded a specific impulse significantly higher than that of previous composite propellants.37

ARC’s polyvinyl chloride, however, did not serve as the binder for Polaris. Instead, the binder used was a polyurethane material developed by Aerojet in conjunction with a small nitropolymer pro­gram funded by the Office of Naval Research about 1947 to seek high-energy binders for solid propellants. A few Aerojet chemists synthesized a number of high-energy compounds, but the process required levels of heating that were unsafe with potentially explo­sive compounds. Then one of the chemists, Rodney Fischer, found “an obscure reference in a German patent" suggesting “that iron chelate compounds would catalyze the reaction of alcohols and iso­cyanates to make urethanes at essentially room temperature." This discovery started the development of polyurethane propellants in many places besides Aerojet.

In the meantime, in 1949 Karl Klager, then working for the Of­fice of Naval Research in Pasadena, suggested to Aerojet’s parent

firm, General Tire, that it begin work on foamed polyurethane, leading to two patents held by Klager, along with Dick Geckler and R. Parette of Aerojet. In 1950, Klager began working for Aerojet. By 1954, he headed the rocket firm’s solid-propellant development group. Once the Polaris program began in December 1956, Klager’s group decided to reduce the percentage of solid oxidizer as a compo­nent of the propellant by including oxidizing capacity in the binder, using a nitromonomer as a reagent to produce the polyurethane plus some inert polynitro compounds as softening agents. In April 1955, the Aerojet group found out about the work of Rumbel and Hender­son. Overcoming explosions due to cracks in the grain and profiting from other developments from multiple contributors, they discov­ered successful propellants for both stages of Polaris A1.

These consisted of a cast, case-bonded polyurethane composition including different percentages of ammonium perchlorate and alu­minum for stages one and two, both of them featuring a six-point, internal-burning, star configuration. With four nozzles for each stage, this propellant yielded a specific impulse of almost 230 lbf – sec/lbm for stage one and nearly 255 lbf-sec/lbm for stage two. The latter specific impulse was higher in part because of the reduced atmospheric pressure at the high altitudes where it was fired, com­pared with stage one, which was fired at sea level.38

The addition of aluminum to Aerojet’s binder essentially solved the problem of performance for Polaris. Other innovations in the areas of warhead size plus guidance and control were necessary to make Polaris possible, but taken together with those for the propel­lants, they enabled Polaris A1 to be only 28.6 feet long and 4.5 feet in diameter (as compared with Jupiter S’s 44 feet and 10 feet, respec­tively). The weight reduction was from 162,000 pounds for Jupiter S to less than 29,000 pounds for Polaris. The cases for both stages of Polaris A1 consisted of rolled and welded steel. This had been modified according to Aerojet specifications resulting from exten­sive metallurgical investigations.

Each of the four nozzles for stage one (and evidently, stage two as well) consisted of a steel shell, a single-piece throat of molybdenum, and an exit-cone liner made of “molded silica phenolic between steel and molybdenum." A zirconium oxide coating protected the steel portion. The missile’s steering came from jetavators designed by Willy Fiedler of Lockheed, a German who had worked on the V-1 program during World War II and had developed the concept for the device while employed by the U. S. Navy at Point Mugu Naval Air Missile Test Center, California. He had patented the idea and then adapted it for Polaris. Jetavators for stage one were molybdenum

FIG. 6.4

An unidentified solid-rocket motor being tested in an altitude wind tunnel at NASA’s Lewis Research Center (later Glenn Research Center) in 1959, one kind of test done for Polaris. (Photo courtesy of NASA)

rings with spherical inside surfaces that rotated into the exhaust stream of the four nozzles and deflected the flow to provide pitch, yaw, and roll control. The jetavators for stage two were similar.39

Besides requiring steering, the missile needed precise thrust ter­mination when it reached the correct point in its trajectory. This could be achieved on liquid-propellant missiles simply by stopping the flow of propellants. For solids, the task was more difficult. The Polaris team used pyrotechnics to blow out plugs in six ports in front of the second stage at the proper moment in the trajectory. This per­mitted exhaust gases to escape and halt the acceleration so that the warhead would travel on a ballistic path to the target area.40

Flight testing of Polaris revealed, among other problems, a loss of control due to electrical-wiring failure at the base of stage one. This resulted from aerodynamic heating and a backflow of hot exhaust gases. To diagnose and solve the problem, engineers in the program obtained the help of “every laboratory and expert," using data from four flights, wind tunnels, sled tests, static firings, “and a tremen­dous analytic effort by numerous laboratories." The solution placed fiberglass flame shielding supplemented by silicone rubber over the affected area to shield it from hot gases and flame.41

Another problem for Polaris to overcome was combustion in­stability. Although this phenomenon is still not fully understood,
gradually it has yielded to research in a huge number of institutions, including universities and government labs, supported by funding by the three services, the Advanced Research Projects Agency, and NASA. Levering Smith credited Edward W. Price in the Research Department at NOTS with helping to understand the phenomenon. By this time, Price had earned a B. S. in physics and math at UCLA. In February 1960, he completed a major (then-classified) paper on combustion instability, which stated, “This phenomenon results from a self-amplifying oscillatory interaction between combustion of the propellant and disturbances of the gas flow in the combustion chamber." It could cause erratic performance, even destruction of motor components. Short of this, it could produce vibrations that would interfere with the guidance/control system. To date, “only marginal success" had been achieved in understanding the phenom­enon, and “trial-and-error development continues to be necessary." But empirical methods gradually were yielding information, for example, that energy fluxes could amplify pressure disturbances, which had caused them in the first place.42 The subsequent success of Polaris showed that enough progress had been made by this time that unstable combustion would not be a major problem for the missile.

Long before Polaris A1 was operational, in April 1958 the DoD had begun efforts to expand the missile’s range from the 1,200- nautical-mile reach of the actual A1 to the 1,500 nautical miles origi­nally planned for it. The longer-range missile, called Polaris A2, was originally slated to achieve the goal through higher-performance propellants and lighter cases and nozzles in both stages. But the navy Special Projects Office decided to confine these improvements to the second stage, where they would have greater effect. (With the second stage already at a high speed and altitude when it began fir­ing, it did not have to overcome the weight of the entire missile and the full effects of the Earth’s gravity at sea level.) Also, in the sec­ond stage, risk of detonation of a high-energy propellant after igni­tion would not endanger the submarine. Hence, the SPO invited the Hercules Powder Company to propose a higher-performance second stage.43

As a result, Aerojet provided the first stage for Polaris A2, and Her­cules, the second. Aerojet’s motor was 157 inches long (compared with 126 inches for Polaris A1; the additional length could be ac­commodated by the submarines’ launch tubes because the navy had them designed with room to spare). It contained basically the same propellant used in both stages of Polaris A1 with the same grain con­figuration. Hercules’ second stage had a filament-wound case and a

cast, double-base grain that contained ammonium perchlorate, ni­trocellulose, nitroglycerin, and aluminum, among other ingredients. The grain configuration consisted of a 12-point, internal-burning star. It yielded a specific impulse of more than 260 lbf-sec/lbm under firing conditions. The motor was 84 inches long and 54 inches in diameter, featuring four swiveling nozzles with exit cones made of steel, asbestos phenolic, and Teflon plus a graphite insert.44

This second-stage motor resulted from an innovation that in­creased performance by adding ammonium perchlorate to the cast, double-base process used in Hercules’ third stage for the Vanguard launch vehicle. Hercules’ ABL developed this new kind of propel­lant, known as composite-modified double base (CMDB), by 1958, evidently with the involvement of John Kincaid and Henry Shuey, 244 developers of the earlier cast, double-base process.45

Chapter 6 Even before 1958, however, Atlantic Research Corporation had developed a laboratory process for preparing CMDB. In its manu­factured state, nitrocellulose is fibrous and unsuitable for use as an unmodified additive to other ingredients being mixed to create a propellant. Arthur Sloan and D. Mann of ARC, however, developed a process that dissolved the nitrocellulose in nitrobenzene and then separated out the nitrocellulose by mixing it with water under high shear (a process known as elutriation). The result was a series of compact, spherical particles of nitrocellulose with small diameters (about 1 to 20 microns). Such particles combined readily with liq­uid nitroplasticizers and crystalline additives in propellant mixers. The result could be cast into cartridge-loaded grains or case-bonded rocket cases and then converted to a solid with the application of moderate heat. Sloan and Mann patented the process and assigned it to ARC. Then, in 1955, Keith Rumbel and Charles Henderson at ARC began scaling the process up to larger grain sizes and devel­oping propellants. They developed two CMDB formulations begin­ning in 1956. When ARC’s pilot plant became too small to support the firm’s needs, production shifted to Indian Head, Maryland. Be­cause the plastisol process they had developed was simpler, safer, and cheaper than other processes then in existence, Henderson said that Hercules and other producers of double-base propellants even­tually adopted his firm’s basic method of production.

Engineers did not use it for upper stages of missiles and launch vehicles until quite a bit later, however, and then only after chem­ists at several different laboratories had learned to make the pro­pellant more rubberlike by extending the chains and cross-linking the molecules to increase the elasticity. Hercules’ John Allabashi at ABL began in the early 1960s to work on chain extenders and

cross-linking, with Ronald L. Simmons at Hercules’ Kenvil, New Jersey, plant continuing this work. By about the late 1960s, chem­ists had mostly abandoned use of plastisol nitrocellulose in favor of dissolving nitrocellulose and polyglycol adipate together, followed by a cross-linking agent such as isocyanate. The result was the type of highly flexible CMDB propellant used on the Trident submarine- launched ballistics missiles beginning in the late 1970s.46

Meanwhile, the rotatable nozzles on the second stage of Polaris A2, which were hydraulically operated, were similar in design to those already being used on the air force’s Minuteman I, and award of the second-stage contract to Hercules reportedly resulted from the performance of the third-stage motor Hercules was developing for Minuteman I, once again illustrating technology transfer be­tween services. (Stage one of the A2 retained the jetavators from A1.) The A2 kept the same basic shape and guidance/control system as the A1, the principal change being more reliable electronics for the guidance/control system. By the time Polaris A2 became opera­tional in June 1962, now-Vice Admiral Raborn had become deputy chief of naval operations for research and development. In February

1962, Rear Adm. I. J. “Pete" Gallantin became director of the Spe­cial Projects Office with Rear Adm. Levering Smith remaining as technical director.47

As a follow-on to Polaris A2, in September 1960, Secretary of De­fense Robert McNamara approved development of a 2,500-nautical – mile version of Polaris that became the A3. To create a missile that would travel an additional 1,000 nautical miles while being launched from the same tubes on the submarine as the A1 required new propellants and a higher mass fraction. The new requirement also resulted in a change from the “bottle shape" of the A1 and A2 to a shape resembling a bullet. In the attempt to help increase the mass fraction, Aerojet, the first-stage manufacturer, acquired the Houze Glass Corporation at Point Marion, Pennsylvania, and moved that firm’s furnaces, patents, technical data, and personnel to the Aerojet Structural Materials Division in Azusa, California, in early

1963. This acquisition gave Aerojet the capability to make filament – wound cases like those used by Hercules on stage two of Polaris A2. The new propellant Aerojet used was a nitroplasticized polyurethane containing ammonium perchlorate and aluminum, configured with an internal-burning, six-point star. This combination raised the spe­cific impulse less than 10 lbf-sec/lbm.48

Unfortunately, the flame temperature of the new propellant was so high that it destroyed the nozzles on the first stage. Aerojet had to reduce the flame temperature from a reported 6,300°F to slightly

less than 6,000°F and to make the nozzles more substantial, using silver-infiltrated tungsten throat inserts to withstand the high tem­perature and chamber pressure. As a result, the weight saving from using the filament-wound case was lost in the additional weight of the nozzle. Hence, the mass fraction for the A3’s first stage was actually slightly lower than that for the A2, meaning that the same quantity of propellant in stage one of the A3 had to lift slightly more weight than did the lower-performing propellant for stage one of the A2.49

For stage two of the A3, Hercules used a propellant containing an energetic high explosive named HMX, a smaller amount of am­monium perchlorate, nitrocellulose, nitroglycerin, and aluminum, among other ingredients. It configured this propellant into an in – 246 ternal-burning cylindrical configuration with many major and mi – Chapter 6 nor slots in the aft end of the stage, creating a cross section that resembled a Christmas-tree ornament. This propellant offered a significantly higher specific impulse than stage two of Polaris A2. Also, the new stage two used a different method of achieving thrust vector control (steering). It injected Freon into the exhaust, creating a shock pattern to deflect the stream and achieve the same results as movable nozzles at a much smaller weight penalty.

A further advantage of this system was its lack of sensitivity to the temperature of the propellant flame. The Naval Ordnance Test Station performed the early experimental work on this use of a liq­uid for thrust vector control. Aerojet, Allegany Ballistics Laboratory, and Lockheed did analytical work, determined the ideal locations for the injectors, selected the fluid to be used, and developed the injectors as well as a system for expelling excess fluid. The Polaris A3 team first successfully tested the new technology on the sec­ond stage of flight A1X-50 on September 29, 1961. This and other changes increased the mass fraction for stage two of Polaris A3 to 0.935 (from 0.871 for stage two of the A2), a major improvement that together with the increased performance of the new propellant, contributed substantially to the greater range of the new missile.50

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

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

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

May 5, 1961, from Pad 5 at Cape Canaveral. (Photo courtesy of NASA)

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

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

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

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

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

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

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

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

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

Despite the problems with the second stage of Vanguard, the air force used a modified version on its Thor-Able launch vehicle, showing the transfer of technology from the navy to the air service. The Able was more successful than the Vanguard prototype for two reasons. One was special cleaning and handling techniques for the propel­lant tanks that came into being after Vanguard had taken delivery of many of its tanks. Also, Thor-Able did not need to extract maximum performance from the second stage as Vanguard did, so it did not have to burn the very last dregs of propellant in the tanks. The residue that the air force did not need to burn contained more of the scale from the tanks than did the rest of the propellant. Consequently, the valves could close before most of the scale entered the fuel lines, the evident cause of many of Vanguard’s problems.30

Designated AJ10-40 (in contrast to the Vanguard second stage’s AJ10-37), the Able was a modified Vanguard stage that still used a regeneratively cooled combustion chamber with aluminum-tubular construction. Able kept the propellants (IWFNA and UDMH), tanks, helium pressurization system, and propellant valves from the Van­guard. The engine produced a thrust of about 7,500 pounds for roughly 120 seconds.31

First used in 1958, the Able upper stage remained in operation until January 1960, when Aerojet’s much more capable Able-Star replaced it. The newer stage resulted from an Advanced Research

Projects Agency directive of July 1, 1959. Aerojet could develop the Able-Star engine (AJ10-104) in a matter of months because it was derived from the Able engines and because it was simple. Directed to make the system rugged, with only those subsystems and compo­nents needed to meet requirements for restart, attitude control dur­ing coasting periods, and longer burning time than the Able could provide, Aerojet engineers sought “to achieve maximum flight ca­pability through limited redesign, overall simplification and opti­mum utilization of flight-proven components."32

Aerojet designed and built the combustion chamber to be “prac­tically identical" to the one used on the Able upper stages, so it re­mained an aluminum, regeneratively cooled, pressure-fed device. For unstated reasons that may have involved the air force’s desire to have the same propellants for Agena and Able-Star, the latter stage switched from the IWFNA used in Able to IRFNA (inhibited red fum­ing nitric acid) as the oxidizer, keeping UDMH as the fuel. Helium under pressure continued to feed the propellants to the combustion chamber, where the injector had concentric rings of orifices that mixed the hypergolic IRFNA and UDMH in an impinging-stream pattern. There were three helium containers, made of titanium, to supply the pressurizing gas. Experience had suggested no need for a nozzle closure diaphragm previously used to ensure high-altitude starts. Another change from Able was an optional nozzle extension 158 that allowed an expansion ratio of either 20:1 without it or 40:1 Chapter 4 with it. Rated thrust rose from 7,575 to 7,890 pounds with the noz­zle extended; rated specific impulse climbed proportionally.33

Although Aerojet’s design and development of Able-Star were quick, they were not problem-free. The virtually identical com­bustion chambers for the Able engines required test firings of only 115 seconds in duration. Able-Star’s had to undergo five static test firings of 300 seconds in duration because of the longer tanks and the increased burning time of the newer upper stage. With a new design for the injector manifold apparently resulting from the con­version to IRFNA as the oxidizer and coolant, during November 1959 Aerojet experienced a burnthrough of the injector plate and the cooling tubes in its vicinity. In a further piece of apparent cut – and-try engineering, Aerojet made appropriate (but unspecified) ad­justments to the designs, and two combustion chambers operated successfully for the full 300 seconds later that month.34

The Transit 1B launch by a Thor/Able-Star on April 13, 1960, marked the first programmed restart of a rocket engine in flight. For Transit 2A—launched on June 22, 1960—there was a problem with sloshing of the propellants in the stage-two tanks, which produced

roll forces. This resulted in an imperfect but usable orbit. To limit the sloshing, engineers added anti-slosh baffles to both Able-Star propellant tanks.35

In a successful launch of the Courier 1B satellite (on October 4, 1960), the anti-slosh baffles apparently had worked. The purpose of the satellite was to test the ability of a spacecraft to relieve crowded communications lines via delayed relay of information. The experi­ment was successful, with large amounts of data transmitted be­tween Puerto Rico and New Jersey. However, the delay of up to two hours before the message was repeated (after the satellite completed its orbit) was unsatisfactory for military purposes and telephone transmission, so the future lay with much higher, geosynchronous orbits. There, the satellite remained in a fixed position relative to a given location on the rotating Earth, allowing nearly instantaneous relay of messages. Overall, there were 20 Thor/Able-Star launches, of which 5 were failures, for a 75 percent success rate.36

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

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

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

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

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

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

Long before the final launch of the Thor-Able, nevertheless, the Department of Defense’s Advanced Research Projects Agency had

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

FIG. 2.6

A Delta 1910 vehicle launching Orbiting Solar Observatory 8 on June 21, 1975, showing the Castor II strap-on boosters at the base of the vehicle to add to the thrust. (Photo courtesy of NASA)

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