The J-2, commonly known as Rocketdyne J-2, was a liquid-fuel cryogenic rocket engine used on NASA's Saturn IB and Saturn V launch vehicles. Built in the United States by Rocketdyne, the J-2 burned cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants, with each engine producing of thrust in vacuum. The engine's preliminary design dates back to recommendations of the 1959 Silverstein Committee. Rocketdyne won approval to develop the J-2 in June 1960 and the first flight, AS-201, occurred on 26 February 1966. The J-2 underwent several minor upgrades over its operational history to improve the engine's performance, with two major upgrade programs, the de Laval nozzle-type J-2S and aerospike-type J-2T, which were cancelled after the conclusion of the Apollo program.
The engine produced a specific impulse (Isp) of in a vacuum (or at sea level) and had a mass of approximately . Five J-2 engines were used on the Saturn V's S-II second stage, and one J-2 was used on the S-IVB upper stage used on both the Saturn IB and Saturn V. Proposals also existed to use various numbers of J-2 engines in the upper stages of an even larger rocket, the planned Nova. The J-2 was America's largest production LH2-fuelled rocket engine before the RS-25. A modernized version of the engine, the J-2X, was considered for use on the Earth Departure Stage of NASA's Space Shuttle replacement, the Space Launch System.
Unlike most liquid-fueled rocket engines in service at the time, the J-2 was designed to be restarted once after shutdown when flown on the Saturn V S-IVB third stage. The first burn, lasting about two minutes, placed the Apollo spacecraft into a low Earth parking orbit. After the crew verified that the spacecraft was operating nominally, the J-2 was re-ignited for translunar injection, a 6.5 minute burn which accelerated the vehicle to a course for the Moon.
Components
Thrust chamber and gimbal system The J-2's thrust chamber assembly served as a mount for all engine components, and was composed of the thrust chamber body, injector and dome assembly, gimbal bearing assembly, and augmented spark igniter.
The thrust chamber was constructed of thick stainless steel tubes, stacked longitudinally and furnace-brazed to form a single unit. The chamber was bell-shaped with a 27.5:1 expansion area ratio for efficient operation at altitude, and was regeneratively cooled by the fuel. Fuel entered from a manifold, located midway between the thrust chamber throat and the exit, at a pressure of more than . In cooling the chamber, the fuel made a one-half pass downward through 180 tubes and was returned in a full pass up to the thrust chamber injector through 360 tubes. Once propellants passed through the injector, they were ignited by the augmented spark igniter and burned to impart a high velocity to the expelled combustion gases to produce thrust.
The thrust chamber injector received the propellants under pressure from the turbopumps, then mixed them in a manner that produced the most efficient combustion. 614 hollow oxidizer posts were machined to form an integral part of the injector, with fuel nozzles (each swaged to the face of the injector) threaded through and installed over the oxidizer posts in concentric rings. The injector face was porous, being formed from layers of stainless steel wire mesh, and was welded at its periphery to the injector body. The injector received LOX through the dome manifold and injected it through the oxidizer posts into the combustion area of the thrust chamber, while fuel was received from the upper fuel manifold in the thrust chamber and injected through the fuel orifices which were concentric with the oxidizer orifices. The propellants were injected uniformly to ensure satisfactory combustion. The injector and oxidizer dome assembly was located at the top of the thrust chamber. The dome provided a manifold for the distribution of the LOX to the injector and served as a mount for the gimbal bearing and the augmented spark igniter.
The augmented spark igniter (ASI) was mounted to the injector face and provided the flame to ignite the propellants in the combustion chamber. When engine start was initiated, the control system started the initial flow of oxidizer and fuel to the spark igniter. Simultaneously, the spark exciters energized two spark plugs mounted in the side of the ASI combustion chamber. As the oxidizer and fuel entered the combustion chamber of the ASI, they mixed and were ignited, with proper ignition being monitored by an ignition monitor mounted in the ASI. The ASI operated continuously during entire engine firing, was uncooled, and was capable of multiple re-ignitions under all environmental conditions. On Apollo 6, the hydrogen line to one of the J-2 engines' ASI broke, initiating a chain of events that shut down two of the five S-II engines prematurely.
Thrust was transmitted through the gimbal (mounted to the injector and oxidizer dome assembly and the vehicle's thrust structure), which consisted of a compact, highly loaded () universal joint consisting of a spherical, socket-type bearing. This was covered with a Teflon/fiberglass coating that provided a dry, low-friction bearing surface. The gimbal included a lateral adjustment device for aligning the combustion chamber with the vehicle, so that, in addition to transmitting the thrust from the injector assembly to the vehicle thrust structure, the gimbal also provided a pivot bearing for deflection of the thrust vector, thus providing flight attitude control of the vehicle.
Propellant feed system The propellant feed system consists of separate fuel and oxidizer turbopumps (the bearings of which were lubricated by the fluid being pumped because the extremely low operating temperature of the engine precluded use of lubricants or other fluids), several valves (including the main fuel valve, main oxidizer valve, propellant utilization valve and fuel and oxidizer bleed valves), fuel and oxidizer flowmeters, and interconnecting lines.
Fuel turbopump The fuel turbopump, mounted on the thrust chamber, was a turbine-driven, axial flow pumping unit consisting of an inducer, a seven-stage rotor, and a stator assembly. It was a high-speed pump operating at 27,000 rpm, and was designed to increase hydrogen pressure from (absolute) through high-pressure ducting at a flowrate which develops . Power for operating the turbopump was provided by a high-speed, two-stage turbine. Hot gas from the gas generator was routed to the turbine inlet manifold which distributed the gas to the inlet nozzles where it was expanded and directed at a high velocity into the first stage turbine wheel. After passing through the first stage turbine wheel, the gas was redirected through a ring of stator blades and enters the second stage turbine wheel. The gas left the turbine through the exhaust ducting. Three dynamic seals in series prevented the pump fluid and turbine gas from mixing. Power from the turbine was transmitted to the pump by means of a one-piece shaft.
Oxidizer turbopump The oxidizer turbopump was mounted on the thrust chamber diametrically opposite the fuel turbopump. It was a single-stage centrifugal pump with direct turbine drive. The oxidizer turbopump increases the pressure of the LOX and pumps it through high-pressure ducts to the thrust chamber. The pump operated at 8,600 rpm at a discharge pressure of (absolute) and developed . The pump and its two turbine wheels are mounted on a common shaft. Power for operating the oxidizer turbopump was provided by a high-speed, two-stage turbine which was driven by the exhaust gases from the gas generator. The turbines of the oxidizer and fuel turbopumps were connected in a series by exhaust ducting that directed the discharged exhaust gas from the fuel turbopump turbine to the inlet of the oxidizer turbopump turbine manifold. One static and two dynamic seals in series prevented the turbopump oxidizer fluid and turbine gas from mixing.
Beginning the turbopump operation, hot gas entered the nozzles and, in turn, the first stage turbine wheel. After passing through the first stage turbine wheel, the gas was redirected by the stator blades and entered the second stage turbine wheel. The gas then left the turbine through exhaust ducting, passed through the heat exchanger, and exhausted into the thrust chamber through a manifold directly above the fuel inlet manifold. Power from the turbine was transmitted by means of a one-piece shaft to the pump. The velocity of the LOX was increased through the inducer and impeller. As the LOX entered the outlet volute, velocity was converted to pressure and the LOX was discharged into the outlet duct at high pressure.
Fuel and oxidizer flowmeters The fuel and oxidizer flowmeters were helical-vaned, rotor-type flowmeters. They were located in the fuel and oxidizer high-pressure ducts. The flowmeters measured propellant flowrates in the high-pressure propellant ducts. The four-vane rotor in the hydrogen system produced four electrical impulses per revolution and turned approximately 3,700 rpm at nominal flow. The six-vane rotor in the LOX system produced six electrical impulses per revolution and turned at approximately 2,600 rpm at nominal flow.
Valves The propellant feed system required a number of valves to control the operation of the engine by changing the flow of propellant through the engine's components: The main fuel valve was a butterfly-type valve, spring-loaded to the closed position, pneumatically operated to the open position, and pneumatically assisted to the closed position. It was mounted between the fuel high-pressure duct from the fuel turbopump and the fuel inlet manifold of the thrust chamber assembly. The main fuel valve controlled the flow of fuel to the thrust chamber. Pressure from the ignition stage control valve on the pneumatic control package opened the valve during engine start and, as the gate started to open, it allowed fuel to flow to the fuel inlet manifold. The main oxidizer valve (MOV) was a butterfly-type valve, spring-loaded to the closed position, pneumatically operated to the open position, and pneumatically assisted to the closed position. It was mounted between the oxidizer high-pressure duct from the oxidizer turbopump and the oxidizer inlet on the thrust chamber assembly. Pneumatic pressure from the normally closed port of the mainstage control solenoid valve was routed to both the first and second stage opening actuators of the main oxidizer valve. Application of opening pressure in this manner, together with controlled venting of the main oxidizer valve closing pressure through a thermal-compensating orifice, provided a controlled ramp opening of the main oxidizer valve through all temperature ranges. A sequence valve, located within the MOV assembly, supplied pneumatic pressure to the opening control part of the gas generator control valve and through an orifice to the closing part of the oxidizer turbine bypass valve. The propellant utilization (PU) valve was an electrically operated, two-phase, motor-driven, oxidizer transfer valve and is located at the oxidizer turbopump outlet volute. The propellant utilization valve ensured the simultaneous exhaustion of the contents of the propellant tanks. During engine operation, propellant level sensing devices in the vehicle propellant tanks controlled the valve gate position for adjusting the oxidizer flow to ensure simultaneous exhaustion of fuel and oxidizer.