RL-10

home
topic index

RL-10
Credit - Lockheed Martin

Engine Model: RL-10. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 66.700 kN (14,995 lbf). Isp: 410 sec. Isp (sea level): 10 sec. Burn time: 482 sec. Mass Engine: 131 kg (288 lb). Diameter: 0.92 m (3.00 ft). Chambers: 1. Chamber Pressure: 24.00 bar. Area Ratio: 40.00. Thrust to Weight Ratio: 44.63. Country: USA. First Flight: 1961. Last Flight: 1965. Flown: 60.

Engine Model: RL-10A-1. Government Designation: LR115. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 66.700 kN (14,995 lbf). Isp: 425 sec. Isp (sea level): 10 sec. Burn time: 430 sec. Mass Engine: 131 kg (288 lb). Diameter: 1.53 m (5.00 ft). Chambers: 1. Chamber Pressure: 24.00 bar. Area Ratio: 47.00. Thrust to Weight Ratio: 51.94. Country: USA. First Flight: 1961. Last Flight: 1967. Flown: 44.

Engine Model: RL-10A-3. Government Designation: LR119. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 65.600 kN (14,747 lbf). Isp: 444 sec. Burn time: 470 sec. Mass Engine: 131 kg (288 lb). Diameter: 1.53 m (5.00 ft). Length: 2.49 m (8.16 ft). Chambers: 1. Chamber Pressure: 28.00 bar. Area Ratio: 57.00. Oxidizer to Fuel Ratio: 5.00. Thrust to Weight Ratio: 51.07. Country: USA. Status: Study 1968. First Flight: 1967. Last Flight: 1983. Flown: 112.

Engine Model: RL-10A-3A. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 73.400 kN (16,501 lbf). Isp: 444 sec. Burn time: 550 sec. Mass Engine: 141 kg (310 lb). Diameter: 1.01 m (3.31 ft). Length: 1.78 m (5.83 ft). Chambers: 1. Chamber Pressure: 32.31 bar. Area Ratio: 61.00. Oxidizer to Fuel Ratio: 5.00. Thrust to Weight Ratio: 53.24. Country: USA. First Flight: 1984. Last Flight: 2005. Flown: 134.

Engine Model: RL-10A-4. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 92.500 kN (20,795 lbf). Isp: 449 sec. Burn time: 392 sec. Mass Engine: 168 kg (370 lb). Diameter: 1.17 m (3.83 ft). Length: 2.29 m (7.51 ft). Chambers: 1. Chamber Pressure: 39.12 bar. Area Ratio: 84.00. Oxidizer to Fuel Ratio: 5.50. Thrust to Weight Ratio: 36.17. Country: USA. Status: Out of production. First Flight: 1992. Last Flight: 2004. Flown: 106.

Engine Model: RL-10A-4-1. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 99.100 kN (22,279 lbf). Isp: 451 sec. Burn time: 740 sec. Mass Engine: 167 kg (368 lb). Diameter: 1.53 m (5.00 ft). Chambers: 1. Chamber Pressure: 39.00 bar. Area Ratio: 84.00. Thrust to Weight Ratio: 60.53. Country: USA. Status: Out of production. First Flight: 2000. Last Flight: 2004. Flown: 2.00.

For Centaur IIIA, one of Centaur IIAS's two RL10A-4 engines is removed. The remaining engine is re-positioned to a center-mount, and electro-mechanical thrust vector control actuators replace the hydraulically actuated system previously in use. Guidance, tank pressurization, and propellant usage controls for both Atlas and Centaur phases are provided by the inertial navigation unit (INU) located on the forward equipment module. The first Centaur burn lasts about nine minutes after which the Centaur and its payload coast in a parking orbit. During the first burn, approximately ten seconds after ignition, the payload fairing is jettisoned. The second Centaur ignition occurs about 23 minutes into the flight, continues for about three minutes, and is followed several minutes later by the separation of the spacecraft from Centaur.

Engine Model: RL-10A-4-2. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 99.100 kN (22,279 lbf). Isp: 451 sec. Burn time: 740 sec. Mass Engine: 167 kg (368 lb). Diameter: 1.53 m (5.00 ft). Chambers: 1. Chamber Pressure: 39.00 bar. Area Ratio: 84.00. Thrust to Weight Ratio: 60.53. Country: USA. Status: In production. First Flight: 2002. Last Flight: 2006. Flown: 16.

Engine Model: RL-10A-5. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 64.700 kN (14,545 lbf). Thrust(sl): 54.800 kN (12,320 lbf). Isp: 373 sec. Isp (sea level): 316 sec. Burn time: 127 sec. Mass Engine: 143 kg (315 lb). Diameter: 1.02 m (3.33 ft). Length: 1.07 m (3.51 ft). Chambers: 1. Chamber Pressure: 39.12 bar. Area Ratio: 4.00. Oxidizer to Fuel Ratio: 6.00. Thrust to Weight Ratio: 46.85. Country: USA. First Flight: 1993. Last Flight: 1996. Flown: 48.

Throttleable to 30% of thrust, sea level version of RL10. Four engines were built and were used on the DC-X and the upgraded DC-XA VTOVL SSTO launch vehicle demonstrators. The DC-Y was a conceptual vehicle design and did not use these engines. The RL-10A-5KA was proposed for use on the Kistler K-1 vehicle but was not selected.

Engine Model: RL-10A-5KA. Designer: Pratt and Whitney. Application: Kistler proposal. Propellants: Lox/LH2. Thrust(vac): 100.488 kN (22,591 lbf). Thrust(sl): 88.926 kN (19,991 lbf). Isp: 398 sec. Isp (sea level): 352 sec. Mass Engine: 145 kg (319 lb). Diameter: 1.02 m (3.33 ft). Length: 1.19 m (3.92 ft). Chambers: 1. Chamber Pressure: 40.81 bar. Area Ratio: 8.20. Oxidizer to Fuel Ratio: 6.00. Thrust to Weight Ratio: 70.66. Country: USA. Status: Design 1992.

Throttleable to 30% of thrust, sea level version of RL10 with extendable nozzle for high altitude operation. The RL-10A-5KA was proposed for use on the Kistler K-1 vehicle but was not selected.

Engine Model: RL-10B-2. Designer: Pratt and Whitney. Propellants: Lox/LH2. Thrust(vac): 110.000 kN (24,720 lbf). Isp: 462 sec. Burn time: 700 sec. Diameter: 2.13 m (6.98 ft). Chambers: 1. Area Ratio: 250.00. Country: USA. Status: In production. First Flight: 1998. Last Flight: 2006. Flown: 10.

The upgraded cryogenic second-stage Pratt & Whitney RL10B-2 engine is based on the 30-year heritage of the reliable RL10 engine. It incorporates an extendable exit cone for increased specific impulse (Isp) and payload capability. The basic engine and turbo pump are unchanged relative to the RL10. The engine gimbal system uses electromechanical actuators that increase reliability while reducing both cost and weight. The propulsion system and attitude control system (ACS) utilize flight-proven off-the-shelf components. The second-stage propulsion system produces a thrust of 24,750 lb with a total propellant load of 37,090 lb, providing a total burn time of approxi-mately 700 sec. Missions requiring more than one restart are accommodated by adding an extra helium bottle for the additional tank repressurization.