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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

There were other developments relating to kerosene-based engines for Thor and Delta, among other vehicles, but the huge Saturn en­gines marked the most important step forward in the use of RP-1 for launch-vehicle propulsion. The H-1 engine for the Saturn I’s first stage resulted from the work of the Rocketdyne Experimental Engines Group on the X-1. Under a contract to the Army Ballistic

H-1 engines, which were used in a cluster of eight to power the first stage of both the Saturn I and the Saturn IB. (Photo courtesy of NASA)

passed its preliminary flight-rating tests, leading to the first flight test on October 27, 1961.47

Meanwhile, Rocketdyne had begun uprating the H-1 to 188,000 pounds of thrust, apparently by adjusting the injectors and increas­ing the fuel and oxidizer flow rates. Although the uprated engine was ready for its preliminary flight-rating tests on September 28, 1962, its uprating created problems with combustion instability that en­gineers had not solved by that time but did fix without detriment to the schedule. The first launch of a Saturn I with the 188,000-pound engine took place successfully on January 29, 1964.48

Development had not been unproblematic. Testing for combus­tion instability (induced by setting off small bombs in the combus­tion chamber beginning in 1963) showed that the injectors inher­ited from the Thor and Atlas could not recover and restore stable combustion once an instability occurred. So Rocketdyne engineers rearranged the injector orifices and added baffles to the injector face. These modifications solved the problem. Cracks in liquid-oxygen domes and splits in regenerative-cooling tubes also required rede­sign. Embrittlement by sulfur from the RP-1 in the hotter environ­ment of the 188,000-pound engine required a change of materials in the tubular walls of the combustion chamber from nickel alloy to stainless steel. There were other problems, but the Saturn person­nel resolved them in the course of the launches of Saturn I and IB from late 1961 to early 1968.49

Because the H-1s would be clustered in two groups of four each for the Saturn I first stage, there were two types of engines. H-1Cs used for the four inboard engines were incapable of gimballing to steer the first stage. The four outboard H-1D engines did the gim – 132 balling. Both versions used bell-shaped nozzles, but the outboard Chapter 3 H-1Ds used a collector or aspirator to channel the turbopump exhaust gases, which were rich in unburned RP-1 fuel, and deposit them in the exhaust plume from the engines to prevent the still-combustible materials from collecting in the first stage’s boat tail.50

The first successful launch of Saturn I did not mean that devel­opers had solved all problems with the H-1 powerplant. On May 28, 1964, Saturn I flight SA-6 unexpectedly confirmed that the first stage of the launch vehicle could perform its function with an en­gine out, a capability already demonstrated intentionally on flight SA-4 exactly 14 months earlier. An H-1 engine on SA-6 ceased to function 117.3 seconds into the 149-second stage-one burn. Telem­etry showed that the turbopump had ceased to supply propellants. Analysis of the data suggested that the problem was stripped gears in the turbopump gearbox. Previous ground testing had revealed to

Rocketdyne and Marshall technicians that there was need for rede­sign of the gear’s teeth to increase their width. Already programmed to fly on SA-7, the redesigned gearbox did not delay flight testing, and there were no further problems with H-1 engines in flight.51

None of the sources for this history explain exactly how Rock­etdyne increased the thrust of each of the eight H-1 engines from 188,000 to 200,000 pounds for the first five Saturn IBs (SA-201 through SA-205) and then to 205,000 pounds for the remaining ve­hicles. It would appear, as with the uprated Saturn I engines, that the key lay in the flow rates of the propellants into the combustion chambers, resulting in increased chamber pressure. After increasing with the shift from the 165,000- to the 188,000-pound H-1s, these flow rates increased again for the 200,000-pound and once more for the 205,000-pound H-1s.52

LAUNCH VEHICLE ENGINES

H-l ON SATURN IB FIRST STAGE

F-l ON SATURN V FIRST STAGE

J-2 ON SATURN IB

SECOND STAGE

J-2 ON SATURN V SECOND STAGE

* ALSO SATURN V THIRD STAGE

FIG. 3.10 Diagram of the engines used on the Saturn IB and Saturn V. The Saturn IB used eight H-1s on its first stage and a single J-2 on its second stage; the Saturn V relied on five F-1 engines for thrust in its first stage, with five J-2s in the second and one J-2 in the third stage. (Photo courtesy of NASA)

For the Saturn V that launched the astronauts and their spacecraft into their trajectory toward the Moon for the six Apollo Moon land­ings, the first-stage engines had to provide much more thrust than the eight clustered H-1s could supply. Development of the larger F-1 engine by Rocketdyne originated with an air force request in 1955. NASA inherited the reports and other data from the early de – 134 velopment, and when Rocketdyne won the NASA contract to build Chapter 3 the engine in 1959, it was, “in effect, a follow-on" effort. Since this agreement preceded a clear conception of the vehicle into which the F-1 would fit and the precise mission it would perform, designers had to operate in a bit of a cognitive vacuum. They had to make early assumptions, followed by reengineering to fit the engines into the actual first stage of the Saturn V, which itself still lacked a firm configuration in December 1961 when NASA selected Boeing to build the S-IC (the Saturn V first stage). Another factor in the design of the F-1 resulted from a decision “made early in the program. . . [to make] the fullest possible use of components and techniques proven in the Saturn I program."

In 1955, the goal had been an engine with a million pounds of thrust, and by 1957 Rocketdyne was well along in developing it. That year and the next, the division of NAA had even test-fired

SATURN IB LAUNCH VEHICLE

PROPOSED MISSIONS

CHARACTERISTICS

• APOLLO SPACECRAFT DEVELOP­MENT AND ORBITAL MANEUVERS • APOLLO CREW TRAINING IN LM RENDEZVOUS AND DOCKING •ADVANCE LARGE BOOSTER TECHNOLOGY •ORBIT LARGE SCIENTIFIC PAYLOADS

MS’C-M-INO HAM

such an engine, with much of the testing done at Edwards AFB’s rocket site, where full-scale testing continued while Rocketdyne did the basic research, development, and production at its plant in Canoga Park. It conducted tests of components at nearby Santa Su – sana Field Laboratory. At Edwards the future air force Rocket Pro­pulsion Laboratory (so named in 1963) had three test stands (1-A, 1-B, and 2-A) set aside for the huge engine. The 1959 contract with NASA called for 1.5 million pounds of thrust, and by April 6, 1961, Rocketdyne was able to static-fire a prototype engine at Edwards whose thrust peaked at 1.64 million pounds.53

Burning RP-1 as its fuel with liquid oxygen as the oxidizer, the F-1 did not break new ground in its basic technology. But its huge thrust level required so much scaling up that, as an MSFC publica­tion said, “An enlargement of this magnitude is in itself an inno­vation." For instance, the very size of the combustion chamber— 40 inches in diameter (20.56 inches for the H-1) with a chamber area almost 4 times that of the H-1 (1,257 to 332 square inches)—required new techniques to braze together the regenerative cooling tubes. Also because of the engine’s size, Rocketdyne adopted a gas-cooled, removable nozzle extension to make the F-1 easier to transport.54

The engine was bell shaped and had an expansion ratio of 16:1 with the nozzle extension attached. Its turbopump consisted of a single, axial-flow turbine mounted directly on the thrust chamber with separate centrifugal pumps for the oxidizer and fuel that were driven at the same speed by the turbine shaft. This eliminated the

The huge first stage of the Saturn V launch vehicle being hoisted by crane from a barge onto the B-2 test stand at the Mississippi Test Facility (later the Stennis Space Center) on January 1, 1967. Nozzles for the F-1 engines show at the bottom of the stage. (Photo courtesy of NASA)

136

Chapter 3

need for a gearbox, which had been a problematic feature of many earlier engines. A fuel-rich gas generator burning the engine pro­pellants powered the turbine. The initial F-1 had the prescribed 1.5 million pounds of thrust, but starting with vehicle 504, Rock – etdyne uprated the engine to 1.522 million pounds. It did so by in­creasing the chamber pressure through greater output from the tur­bine, which in turn required strengthening components (at some expense in engine weight). There were five F-1s clustered in the S-IC stage, four outboard and one in the center. All but the center engine gimballed to provide steering. As with the H-1, there was a hypergolic ignition system.55

Perhaps the most intricate design feature of the F-1 was the in­jection system. As two Rocketdyne engineers wrote in 1989, the

injector “might well be considered the heart of a rocket engine, since virtually no other single component has such a major impact on overall engine performance." The injector not only inserted the propellants into the combustion chamber but mixed them in a pro­portion designed to produce optimal thrust and performance. “As is the case with the design of nearly all complex, high technology hardware," the two engineers added, “the design of a liquid rocket injector is not an exact science, although it is becoming more so as analytical tools are continuously improved. This is because the basic physics associated with all of the complex, combustion pro­cesses that are affected by the design of the injector are only partly known." A portion of the problem lay with the atomization of the propellants and the distribution of the fine droplets to ensure proper mixing. Even as late as 1989, “the atomization process [wa]s one of the most complex and least understood phenomena, and reli­able information [wa]s difficult to obtain." One result of less-than – optimal injector design was combustion instability, whose causes and mechanisms still in 1989 were “at best, only poorly known and understood." Even in 2006, “a clear set of generalized validated design rules for preventing combustion instabilities in new TCs [thrust chambers] ha[d] not yet been identified. Also a good uni­versal mathematical three-dimensional simulation of the complex nonlinear combustion process ha[d] not yet been developed."56

This was still more the case in the early 1960s, and it caused huge problems for development of the F-1 injector. Designers at Rocket – dyne knew from experience with the H-1 and earlier engines that injector design and combustion instability would be problems. They began with three injector designs, all based essentially on that for the H-1. Water-flow tests provided information on the spacing and shape of orifices in the injector, followed by hot-fire tests in 1960 and early 1961. But as Leonard Bostwick, the F-1 engine manager at Marshall, reported, “None of the F-1 injectors exhibited dynamic stability." Designers tried a variety of flat-faced and baffled injec­tors without success, leading to the conclusion that it would not be possible simply to scale up the H-1 injector to the size needed for the F-1. Engineers working on the program did borrow from the H-1 effort the use of bombs to initiate combustion instability, saving a lot of testing to await its spontaneous appearance. But on June 28, 1962, during an F-1 hot-engine test in one of the test stands built for the purpose at the rocket site on Edwards AFB, combustion instabil­ity caused the engine to melt.57

Marshall appointed Jerry Thomson, chief of the MSFC Liquid Fuel Engines Systems Branch, to chair an ad hoc committee to ana-

lyze the problem. Thomson had earned a degree in mechanical en­gineering at Auburn University following service in World War II. Turning over the running of his branch to his deputy, he moved to Canoga Park where respected propulsion engineer Paul Castenholz and a mechanical engineer named Dan Klute, who also “had a spe­cial talent for the half-science, half-art of combustion chamber de­sign," joined him on the committee from positions as Rocketdyne managers. Although Marshall’s committee was not that large, at Rocketdyne there were some 50 engineers and technicians assigned to a combustion devices team, supplemented by people from uni­versities, NASA, and the air force. Using essentially cut-and-try methods, they initially had little success. The instability showed no consistency and set in “for reasons we never quite understood," as Thomson confessed.58

Using high-speed instrumentation and trying perhaps 40-50 de­sign modifications, eventually the engineers found a combination of baffles, enlarged fuel-injection orifices, and changed impinge­ment angles that worked. By late 1964, even following explosions of the bombs, the combustion chamber regained its stability. The en­gineers (always wondering if the problem would recur) rated the F-1 injector as flight-ready in January 1965. However, there were other problems with the injector. Testing revealed difficulties with fuel and oxidizer rings containing multiple orifices for the propellants. Steel rings called lands held copper rings through which the propel­lants flowed. Brazed joints held the copper rings in the lands, and these joints were failing. Engineers gold-plated the lands to create a better bonding surface. Developed and tested only in mid-1965, the new injector rings required retrofitting in engines Rocketdyne had already delivered.59

Overall, from October 1962 to September 1966, there were 1,332 full-scale tests with 108 injectors during the preliminary, the flight­rating, and flight-qualification testing of the F-1 to qualify the en­gine for use. According to one expert, this was “probably the most intensive (and expensive) program ever devoted primarily to solving a problem of combustion instabilities."60

The resultant injector contained 6,300 holes—3,700 for RP-1 and 2,600 for liquid oxygen. Radial and circumferential baffles divided the flat-faced portion of the injector face into 13 compartments, with the holes or orifices arranged so that most of them were in groups of five. Two of the five injected the RP-1 so that the two streams impinged to produce atomization, while the other three in­serted liquid oxygen, which formed a fan-shaped spray that mixed with the RP-1 to combust evenly and smoothly. Driven by the

52,900-horsepower turbine, the propellant pumps delivered 15,471 gallons of RP-1 and 24,811 gallons of liquid oxygen per minute to the combustion chamber via the injector.61

Despite all the effort that went into the injector design, the turbo­pump required even more design effort and time. Engineers experi­enced 11 failures of the system during development. Two of these involved the liquid-oxygen pump’s impeller, which required use of stronger components. The other 9 failures involved explosions. Causes varied. The high acceleration of the shaft on the turbopump constituted one problem. Others included friction between mov­ing parts and metal fatigue. All 11 failures necessitated redesign or change in procedures. For instance, Rocketdyne made the turbine manifold out of a nickel-based alloy manufactured by GE, Rene 41, which had only recently joined the materials used for rocket en­gines. Unfamiliarity with its welding techniques led to cracking near the welds. It required time-consuming research and training to teach welders proper procedures for using the alloy, which could withstand not only high temperatures but the large temperature dif­ferential resulting from burning the cryogenic liquid oxygen. The final version of this turbopump provided the speed and high vol­umes needed for a 1.5-million-pound-thrust engine and did so with minimal parts and high ultimate reliability.62

Once designed and delivered, the F-1 engines required further testing at Marshall and NASA’s Mississippi Test Facility. At the lat­ter, contractors had built an S-IC stand after 1961 on the mud of a swamp along the Pearl River near the Louisiana border and the Gulf of Mexico. Mosquito ridden and snake infested, this area served as home to wild pigs, alligators, and panthers. Construction workers faced 110 bites a minute from salt marsh mosquitoes, against which nets, gloves, repellent, and long-sleeved shirts afforded little protec­tion. Spraying special chemicals from two C-123 aircraft did reduce the number of bites to 10 per minute, but working conditions re­mained challenging. Nevertheless, the stand was ready for use in March 1967, more than a year after the first static test at Marshall. But thereafter, the 410-foot S-IC stand, the tallest structure in Mis­sissippi, became the focus of testing for the first-stage engines.63

Despite static testing, the real proof of successful design came only in actual flight. For AS-501, the first Saturn V vehicle, the flight on November 9, 1967, largely succeeded. The giant launch vehicle lifted the instrument unit, command and service modules, and a boilerplate lunar module to a peak altitude of 11,240 miles. The third stage then separated and the service module’s propulsion sys­tem accelerated the command module to a speed of 36,537 feet per second (about 24,628 miles per hour), comparable to lunar reentry speed. It landed in the Pacific Ocean 9 miles from its aiming point, where the USS Bennington recovered it. The first-stage engines did experience longitudinal oscillations (known as the pogo effect), but these were comparatively minor.64

Euphoria from this success dissipated, however, on April 4, 1968, when AS-502 (Apollo 6) launched. As with AS-501, this vehicle did 140 not carry astronauts onboard, but it was considered “an all-important Chapter 3 dress rehearsal for the first manned flight" planned for AS-503. The initial launch went well, but toward the end of the first-stage burn the pogo effect became much more severe than on AS-501, reaching five cycles per second, which exceeded the spacecraft’s design speci­fications. Despite the oscillations, the vehicle continued its upward course. Stage-two separation occurred, and all five of the engines ignited. Two of them subsequently shut down, but the instrument unit compensated with longer-than-planned burns for the remain­ing three engines and the third-stage propulsion unit, only to have the latter fail to restart in orbit, constituting a technical failure of the mission, although some sources count it a success.65

As Apollo Program Director Samuel Phillips told the Senate Aero­nautical and Space Sciences Committee on April 22, 1968, 18 days

after the flight, pogo was not a new phenomenon, having occurred in the Titan II and come “into general attention in the early days of the Gemini program." Aware of pogo, von Braun’s engineers had tested and analyzed the Saturn V before the AS-501 flight and found “an acceptable margin of stability to indicate" it would not de­velop. The AS-501 flight “tended to confirm these analyses." Each of the five F-1 engines had “small pulsations," but each engine ex­perienced them “at slightly different points in time." Thus, they did not create a problem. But on AS-502, the five 1.5-million-pound engines “came into a phase relationship" so that “the engine pulsa­tion was additive."66

All engines developed a simultaneous vibration of 5.5 hertz (cy­cles per second). The entire vehicle itself developed a bending fre­quency that increased (as it consumed propellants) to 5.25 hertz about 125 seconds into the flight. The engine vibrations traveled longitudinally up the vehicle structure with their peak occurring at the top where the spacecraft was (and the astronauts would be on a flight carrying them). Alone, the vibrations would not have been a problem, but they coupled with the vehicle’s bending frequency, which moved in a lateral direction. When they intersected (with both at about the same frequency), their effects combined and mul­tiplied. In the draft of an article he wrote for the New York Times, Phillips characterized the “complicated coupling" as “analogous to the annoying feedback squeal you encounter when the microphone and loud speaker of a public address system. . . coupled." This cou­pling was significant enough that it might interfere with astronauts’ performance of their duties.67

NASA formed a pogo task force including people from Marshall, other NASA organizations, contractors, and universities. The task force recommended detuning the five engines, changing the fre­quencies of at least two so that they would no longer produce vi­brations at the same time. Engineers did this by inserting liquid helium into a cavity formed in a liquid-oxygen prevalve with a cast­ing that bulged out and encased an oxidizer feed pipe. The bulging portion was only half filled with the liquid oxygen during engine operation. The helium absorbed pressure surges in oxidizer flow and reduced the frequency of the oscillations to 2 hertz, lower than the frequency of the structural oscillations. Engineers eventually applied the solution to all four outboard engines. Technical people contributing to this solution came from Marshall, Boeing, Martin, TRW, Aerospace Corporation, and North American’s Rocketdyne Division.68 This incidence of pogo showed how difficult it was for

FIG. 3.14

Launch of the giant Saturn V on the Apollo 11 mission (July 16, 1969) that carried Neil Armstrong, Edwin Aldrin, and Michael Collins on a trajectory to lunar orbit from which Armstrong and Aldrin descended to walk on the Moon’s surface. (Photo courtesy of NASA)

142

Chapter 3

rocket designers to predict when and how such a phenomenon might occur, even while aware of and actively testing for it. The episode also illustrated the cooperation of large numbers of people from a variety of organizations needed to solve such problems.

The redesign worked. On the next Saturn V mission, Apollo 8 (AS-503, December 21, 1968), the five F-1s performed their mission without pogo (or other) problems. On Christmas Eve the three astro­nauts aboard the spacecraft went into lunar orbit. They completed 10 circuits around the Moon, followed by a burn on Christmas Day to return to Earth, splashing into the Pacific on December 27. For the first time, humans had escaped the confines of Earth’s imme­diate environs and returned from orbiting the Moon. There were no further significant problems with the F-1s on Apollo missions.69

The development of the huge engines had been difficult and unpre­dictable but ultimately successful.

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

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

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

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

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

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

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

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

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.