Titan II Engines

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

FIG. 4.1 Technical drawing with description of the Agena upper stage as of 1968. (Photo courtesy of NASA)

NASA C-1968-538

AGENA PROJECT

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

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

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

National Aeronautics and Space Administration Lewis Research Center

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Chapter 4