Credit: Boeing / Rocketdyne
AKA: MA. Status: Out of Production. Thrust: 2,100.00 kN (472,000 lbf). Unfuelled mass: 1,610 kg (3,540 lb). Specific impulse: 296 s. Specific impulse sea level: 263 s. Burn time: 167 s. Height: 3.43 m (11.25 ft). Diameter: 1.19 m (3.90 ft).
MA - 5A/RS - 56 Atlas Engine System. An updated version of the Atlas MA - 5 system. Upgrades included replacements 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 up to the end of 1993 (49 systems contracted). Dry Mass: 1610 kg. Length: 343 cm. Maximum Diameter: 119 cm. Mounting: all thrust chambers are gimbaled. 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.
|MA-5A Hot Fire Test|
MA-5A engine in hot firing test.