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RL-10 development

The Saturn program's association with the development of liquid hydrogen-oxygen engines officially commenced on 10 August 1960, when MSFC signed a contract with Pratt & Whitney for the development and production of at, engine, known as the LR-119, to be used in the S-IV and S-V stages of the C-1 vehicle envisioned in the Silverstein report. Designed to give 66 700 newtons (17 500 pounds) of thrust, the LR-119 was an uprated version of an early Centaur engine concept, the LR-115. Problems with the development of this new version led to the reconsideration of the original Centaur propulsion system, and in March 1961, the management of MSFC recommended the design of a liquid-hydrogen S-IV stage using the original LR-115 hardware. To compensate for the loss of thrust, MSFC decided to cluster six engines instead of four. On 29 March 1961, NASA Headquarters concurred, and the new six-engine cluster became the official configuration. In the course of development, Pratt & Whitney assigned various designations to the basic liquid hydrogen-oxygen engine. The final design, RL-10-A-1, replaced both the LR-115 and 119, and the RL-10 configuration became standard for both the Centaur and S-IV vehicles by 1961. An early version of the RL-10 design went through its first successful firing in August 1959, and by the winter of 1961, technicians finished the last of the RL-10-A-1 preflight rating tests. The engine's 66 700 newtons (15 000 pounds) of thrust performed 30 percent better than similar designs using hydrocarbon fuels. The A-1 designation identified a test article; on 9 June 1962, Pratt & Whitney finished the preliminary flight rating tests on the RL-10-A-3, intended for installation in operational flight versions of the second stage of the C-1 launch vehicle.16 The nation's first operational liquid hydrogen-oxygen engine was cleared for production.

THE RL-10 PROPULSION SYSTEM

Pratt & Whitney engine design unquestionably benefited from the work at Lewis during 1953-1957, especially the virtues of regenerative cooling with liquid hydrogen.17 Pratt & Whitney added other innovative features. The Saturn program's RL-10 engines were mounted on the S-IV booster manufactured by Douglas as the second stage for the Saturn I. In physical terms, the RL-10 was about as tall as an average man. Its major components included the thrust chamber, fuel and oxidizer [138] turbopump assembly, liquid oxygen flow control valve, spark ignition subsystem, thrust control assembly, and miscellaneous control valves.

The contours of the nozzle configuration owed much to the influence of applied mathematics. Pratt & Whitney wanted a nozzle designed for optimum size and weight in relation to performance, but liquid hydrogen technology was so new that few ground rules were available. Applied math bypassed a lot of costly hardware experimentation, and Pratt & Whitney claimed that the procedures established during the effort became widely used within the rocket propulsion industry.18

The injector, part of the thrust chamber assembly, featured a porous injector face, which was an important innovation. The RL-10 injector strongly resembled a large dish with a shallow, concave surface. Fabricated from material that looked like a heavy screen, the injector's propellant orifices poked through the mesh in concentric rings. The porous injector face did, in fact, consist of layers of stainless steel mesh, produced by a carefully controlled sintering procedure that caused the layers of mesh to become a coherent structure without melting. A controlled flow of gaseous hydrogen filtered through, cooling the injector face and reducing thermal stresses. The material, called Rigi-Mesh by its supplier (the Pall Corporation), apparently originated as a filter used in nuclear research. The product had been extensively used in hydraulic and pneumatic filters in aircraft and jet engines, where extreme vibration environments, high temperatures, and other operational requirements discouraged the use of nonmetallic filters. How Rigi-Mesh was first suggested for use in rocket thrust chambers is unclear. In any case, the Pratt & Whitney injector approach, using the porous mesh face, was a distinct improvement over conventional, flat-face injectors that Lewis Research Center had used.19

The fuel and oxidizer pumps were driven in a "boot strap" arrangement from a turbine assembly rated at 479 to 513 kilowatts. The propellant pumps consisted of a two-stage centrifugal fuel pump and a single-stage centrifugal oxidizer pump. General Dynamics/ Astronautics described the engine's turbopump as the key to operating the RL-10 production version, in which the "boot strap" sequence used gaseous hydrogen. At the start, liquid hydrogen trickled through the turbopump and down through the thrust chamber tubes of the regeneratively cooled engine. Even before the ignition sequence and main stage operation, the flowing liquid hydrogen became gaseous, and could be forced back through the turbopump with enough pressure to start it. This pressure set the hydrogen fuel pump in motion, and a gear train from the hydrogen turbine's main shaft began to drive the liquid oxygen pump-the "boot strap" sequence. After the start of combustion, the heat produced enough gas in the chamber walls to drive the high-speed turbine and also to maintain the combustion level.20

[140] This design offered two main advantages. First, the engine did not require a third propellant or a bipropellant to service a gas generator system (at a weight penalty) for the turbopump. Second, the designers obtained an efficient performance advantage because the hydrogen gases, after driving the turbine, were exhausted into the combustion chamber. All propellants, then, contributed directly to maximum thrust and highest specific impulse. The operation of the turbomachinery incorporated another interesting design feature. The RL-10 was the first production engine to use liquid hydrogen in place of conventional lubrication systems.21

During the test program, NASA and contractor personnel pushed the design to extremes to verify the engine's capability. Designed for a total firing time of 470 seconds, test engineers piled more than 3.5 times that duration onto one engine, running it for a total of 1680 seconds. Some of the test engines successfully operated through 5 to 70 separate firings with no maintenance or replacement of parts, equivalent in some instances to 10 round trips to the moon. "This philosophy of `limits' testing has proven successful in developing an engine with a high reliability and a high degree of confidence," explained key personnel in MSFC's engine program office. They characterized the pioneering RL-10 as a system of notable sophistication and versatility.22