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MA-5A -

Credit: Boeing / Rocketdyne. 15,106 bytes. 220 x 239 pixels.

Designer: Rocketdyne. Propellants: Lox/Kerosene Thrust(vac): 214,144 kgf. Thrust(vac): 2,100.00 kN. Isp: 296 sec. Isp (sea level): 263 sec. Burn time: 167 sec. Mass Engine: 1,610 kg. Diameter: 1.2 m. Length: 3.4 m. Chambers: 1. Chamber Pressure: 48.00 bar. Area Ratio: 8.00. Country: USA. Status: In Production.

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.

MA-5A Hot Fire TestMA-5A Hot Fire Test - MA-5A engine in hot firing test.

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Last update 12 March 2001.
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