![]() ![]() Cryogenic supercritical helium was loaded and stored at 3500 psi. To keep the DPS as simple, lightweight, and reliable as possible, the propellants were pressure-fed with helium gas instead of using heavy, complicated, and failure-prone turbopumps. Gilruth convened his own five-member board, also chaired by Faget, which reversed Grumman's decision on January 18 and awarded the contract to STL. ![]() ![]() Still not satisfied, MSC Director Robert R. Grumman chose Rocketdyne on January 5, 1965. Apollo Spacecraft Program Office manager Joseph Shea formed a committee of NASA, Grumman and Air Force propulsion experts, chaired by American spacecraft designer Maxime Faget, in November 1964 to recommend a choice, but their results were inconclusive. NASA planners expected one of the two drastically different designs would emerge the clear winner, but this did not happen throughout 1964. The first full-throttle firing of Space Technology Laboratories' LM descent engine was carried out in early 1964. STL proposed an engine that was gimbaled as well as throttleable, using flow control valves and a variable-area pintle injector, in much the same manner as does a shower head, to regulate pressure, rate of propellant flow, and the pattern of fuel mixture in the combustion chamber. In May, STL was selected as the competitor to Rocketdyne's concept. Grumman held a bidders' conference on March 14, 1963, attended by Aerojet General, Reaction Motors Division of Thiokol, United Technology Center Division of United Aircraft, and Space Technology Laboratories, Inc. (In fact, accidental ingestion of helium pressurant proved to be a problem on AS-201, the first flight of the Apollo Service Module engine in February 1966.) Therefore, MSC directed Grumman to conduct a parallel development program of competing designs. While NASA's Manned Spacecraft Center (MSC) judged this approach to be plausible, it represented a considerable advance in the state of the art. Rocketdyne proposed a pressure-fed engine using the injection of inert helium gas into the propellant flow to achieve thrust reduction at a constant propellant flow rate. Very little advanced research had been done in variable-thrust rocket engines up to that point. Development Īccording to NASA history publication Chariots for Apollo, "The lunar module descent engine probably was the biggest challenge and the most outstanding technical development of Apollo." A requirement for a throttleable engine was new for crewed spacecraft. The exhaust nozzle extension was designed to crush without damaging the LM if it struck the surface, which happened on Apollo 15. A lightweight cryogenic helium pressurization system was also used. To accomplish these maneuvers, a propulsion system was developed that used hypergolic propellants and a gimballed pressure-fed ablative cooled engine that was capable of being throttled. The propulsion system for the descent stage of the lunar module was designed to transfer the vehicle, containing two crewmen, from a 60-nautical-mile (110 km) circular lunar parking orbit to an elliptical descent orbit with a pericynthion of 50,000 feet (15,000 m), then provide a powered descent to the lunar surface, with hover time above the lunar surface to select the exact landing site. This engine used a pintle injector, which paved the way for other engines to use similar designs. It used Aerozine 50 fuel and dinitrogen tetroxide ( NĤ) oxidizer. and developed by Space Technology Laboratories (TRW) for use in the Apollo Lunar Module descent stage. The descent propulsion system (DPS - pronounced 'dips') or lunar module descent engine (LMDE), internal designation VTR-10, is a variable- throttle hypergolic rocket engine invented by Gerard W. Apollo Lunar Module rocket engine Descent propulsion system (DPS) Country of originġ0,500 lbf (47 kN) maximum, throttleable betweenġ,050 lbf (4.7 kN) and 6,825 lbf (30.36 kN) ![]()
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