Home - Search - Browse - Alphabetic Index: 0- 1- 2- 3- 4- 5- 6- 7- 8- 9
A- B- C- D- E- F- G- H- I- J- K- L- M- N- O- P- Q- R- S- T- U- V- W- X- Y- Z

---

RS-27A

---

RS-27A Credit: Boeing / Rocketdyne

Status: First flight 1989. Date: 1987. Number: 20 . Thrust: 1,054.20 kN (236,994 lbf). Unfuelled mass: 1,091 kg (2,405 lb). Specific impulse: 302 s. Specific impulse sea level: 255 s. Burn time: 274 s. Height: 3.78 m (12.40 ft). Diameter: 1.07 m (3.51 ft).

The RS - 27A powerplant comprises an RS2701B main engine, and twin LR101 - NA - 11 verniers. Introduced in 1990 on the McDonnell Douglas' Delta 7000 series launcher it replaced the RS-27 as the main system for that launcher. It continues in service as part of the Atlas MA- 5A powerplant. Flown: 14 Delta plus 8 Atlas to the end of 1993. Mounting: gimbaled-mounted for pitch/yaw control with gimbaled verniers for roll control. Engine Cycle: gas generator. Oxidizer: liquid oxygen at 250 kg/sec. Fuel: RP-1 hydrocarbon at 111 kg/sec. Mixture Ratio: 2.245:1. Oxidizer Turbopump: 1900 kW, 6784 rpm (7085 rpm at altitude), 70 atm discharge Pressure: Fuel Turbopump: 1289 kW, 70 atm discharge Pressure. Thrust: 890 kN sea level/1054.2 kN vacuum. Thrust Chamber Length: 234 cm. Thrust Chamber Materials: 347 CRES austenitic stainless steel. Thrust Chamber Cooling: regenerative, two passes of fuel through 292 tubes. Combustion Chamber Pressure: 48 atm at injector end. Combustion Chamber Temperature: 3315 Celsius. Combustion Chamber Materials: 347 CRES austenitic stainless steel. Combustion Chamber Cooling: same as thrust chamber. Combustion Chamber Ignition: hypergolic fluid cartridge enclosed in burst diaphragms. Burn Time: 274 sec. Verniers: each LR101-NA-11 at 21.8 kg mass, 4.63/5.30 kN sea level/vac thrust, 209/246 sec sea level/vac Isp, 1.8 mixture ratio, 5.6 expansion ratio(9.8 cm exit diameter), 283 sec burn time. Designed for booster applications. Gas generator, pump-fed. Two vernier engines provide roll control.

Thrust (sl): 890.100 kN (200,102 lbf). Thrust (sl): 90,770 kgf. Engine: 1,091 kg (2,405 lb). Chamber Pressure: 49.00 bar. Area Ratio: 12. Propellant Formulation: Lox/RP-1 Lox/RP-1. Thrust to Weight Ratio: 102.468075150265. Oxidizer to Fuel Ratio: 2.245. Coefficient of Thrust vacuum: 1.80185738739855. Coefficient of Thrust sea level: 1.60185738739855.