Encyclopedia Astronautica
MA-5A



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MA-5A
Credit: Boeing / Rocketdyne
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MA-5A Hot Fire Test
MA-5A engine in hot firing test.
Rocketdyne Lox/Kerosene rocket engine. 2100 kN. Out of Production. Isp=296s. Atlas Engine System, an updated version of the MA-5, included replacments with selected RS-27 components for sea-level Isp increase of 4 secs. First flight 1991.

MA - 5A/RS - 56 Atlas Engine System. An updated version of the Atlas MA - 5 system. Upgrades included replacments with selected RS-27 components. This provided a sea-level Isp increase of 4 secs. In addition to this, the side-mounted verniers were deleted. The roll control and final adjustment functions were integrated into the vehicle's interstage.. Applications: Atlas 2 primary propulsion system. First Flown: December 7th, 1991. Flown: 7 til the end of 1993 (49 systems contracted). Dry Mass: 1610 kg. Length: 343 cm. Maximum Diameter: 119 cm. Mounting: all thrust chambers are gimballed. Engine Cycle: gas generator. Oxidizer: liquid oxygen pump-fed at 505 kg/sec. Fuel: RP-1 hydrocarbon pump-fed at 224 kg/sec. Mixture Ratio: 2.25:1. Oxidizer Turbopump: each 1903 kW, 6730 rpm, 70 atm discharge. Pressure: Fuel Turbopump: 1362 kW, 75 atm discharge Pressure. Thrust: total 2100 kN vacuum, 1890 kN sea level. Isp: 295 sec vacuum, 263 sec sea level. Time to Full Thrust: 2.0 sec. Expansion Ratio: 8:1. Thrust Chamber Length: 249 cm. Thrust Chamber Materials: 347 CRES austenitic stainless steel. Thrust Chamber Cooling Method: regenerative, two passes of fuel through 292 chamber tubes. Combustion Chamber Pressure: 48 atm at injector end. Combustion Chamber Temperature: 3316 Celsius. Combustion Chamber Materials: 347 CRES austenitic stainless steel. Combustion Chamber Cooling as thrust chamber. Combustion Chamber Ignition: hypergolic fluid chamber enclosed in burst diaphragms. Burn Time: 167 sec maximum flight duration.

Thrust (sl): 1,865.900 kN (419,471 lbf). Thrust (sl): 190,270 kgf. Engine: 1,610 kg (3,540 lb). Chamber Pressure: 49.00 bar. Area Ratio: 8. Thrust to Weight Ratio: 133.

Status: Out of Production.
Unfuelled mass: 1,610 kg (3,540 lb).
Height: 3.43 m (11.25 ft).
Diameter: 1.19 m (3.90 ft).
Thrust: 2,100.00 kN (472,000 lbf).
Specific impulse: 296 s.
Specific impulse sea level: 263 s.
Burn time: 167 s.

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Associated Propellants
  • Lox/Kerosene Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. In January 1953 Rocketdyne commenced the REAP program to develop a number of improvements to the engines being developed for the Navaho and Atlas missiles. Among these was development of a special grade of kerosene suitable for rocket engines. Prior to that any number of rocket propellants derived from petroleum had been used. Goddard had begun with gasoline, and there were experimental engines powered by kerosene, diesel oil, paint thinner, or jet fuel kerosene JP-4 or JP-5. The wide variance in physical properties among fuels of the same class led to the identification of narrow-range petroleum fractions, embodied in 1954 in the standard US kerosene rocket fuel RP-1, covered by Military Specification MIL-R-25576. In Russia, similar specifications were developed for kerosene under the specifications T-1 and RG-1. The Russians also developed a compound of unknown formulation in the 1980's known as 'Sintin', or synthetic kerosene. More...

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