AeroJet Propulsion Division
P.O.Box 13222,
Highway 50 & Aerojet Rd.
Telephone (916) 355-1000
Fax | (916) 355-2448
A pressure-fed engine, optimized for altitude operation.
| Application | Delta Stage 2 |
| First Flown | August 1982, Delta 164 |
| Dry Mass | not available, but older model at 109.7 kg |
| Length | not available |
| Mounting | fixed |
| Engine Cycle | pressure-fed |
| Oxidizer | nitrogen tetroxide at 9.1 kg/sec |
| Fuel | Aerozine-50 at 4.76 kg/sec |
| Mixture Ratio | 1.9:1 |
| Thrust | 43.38 kN vacuum |
| Isp | 320.5 sec vacuum |
| Expansion Rate | 65:1 |
| Combustion Chamber Pressure | 8.84 atm |
| Cooling Method | fuel regenerative chamber, radiative
skirt |
| Burn Time | qualified up to 500 sec (unlimited
starts) |
The NASA Space Shuttle Orbiter carries two OMS pods (name coined by
Aerojet), each housing a single Aerojet OMS engine for orbit insertion,
maneuvering, and re-entry initiation. They are capable of 100 missions and
500 starts in space.
| Appilcations | Space Shuttle
orbit/de-orbit insertion, circularization |
| First Flown | April 12th, 1981, on the Orbiter Columbia |
| Number Flown | 14, to end of 1993 |
| Dry Mass | 118 kg |
| Length | 195.6 cm |
| Maximum Diameter | 116.8 cm |
| Mounting | gimballed ( 7 degrees yaw ( 6 pitch by two
electromechanical actuators for thrust vector control |
| Engine Cycle | pressure-fed (improvement underway for
pump-fed) |
| Oxidizer | 6743 kg nitrogen tetroxide in each pod (pods can
be cross-linked) |
| Fuel | 4087 kg of monomethyl hydrazine in each pod (pods can
be cross-linked) |
| Mixture Ratio | 1.65:1 |
| Thrust | 26.7 kN vacuum |
| Isp | 316 sec vacuum |
| Expansion Ratio | 55:1 |
| Combustion Chamber Pressure | 8.62 atm |
| Cooling Method | fuel regenerative for chamber, radiative
for nozzle |
| Burn Time | qualified for 500 starts, 15 hr/100 mission
life, longest firing 1250 sec, de-orbit burn typically 150-250
secs |
Refurbished for space launcher versions from overhaul between 1974 -
1982.
| Designation | Aerojet LR-87-AJ-5 |
| Configuration | twin fixed motors with individual
turbo-pumped assemblies |
| Application | Titan 2 Stage 1 |
| First Flown | 1962 ICBM; Sept. 1988 orbital |
| Dry Mass | 739 kg |
| Length | 3.13 m |
| Maximum Diameter | 1.14 m |
| Engine Cycle | Gas generator |
| Propellants | hypergolic nitrogen tetroxide and
Aerozine-50, delivered at 750 kg/sec |
| Mixture Ratio | 1.93:1 |
| Thrust | 1913 kN sea level |
| Isp | 259 sec at sea level |
| Expansion Ratio | 8:1 |
| Combustion Chamber Pressure | 53.3 atm |
| Burn Time | 158 sec |
| Designation | Aerojet LR-91-AJ-5 |
| Configuration | scaled down version of stage 1 engine
featuring fixed single chamber |
| Applications | Titan 2 stage 2 |
| First Flown | as stage 1 engine |
| Dry Mass | 500 kg |
| Length | 2.80 m |
| Maximum Diameter | 1.68 m |
| Engine Cycle | gas generator |
| Propellants | as stage 1, at 163 kg/sec |
| Thrust | 444.8 kN vacuum |
| Isp | 315 sec vacuum |
| Expansion Ratio | 49.2:1 |
| Combustion Chamber Pressure | 56.2 atm |
| Burn Time | ( 175 secs |
| Mixture Ratio | 1.80 |
Both the current Titan 3 & 4 first stages are powered by these engines.
Replacing the 9 model, this is the only engine - together with its stage 2
derivative - to be operated on storable LOX/RP. It is also the only one
to be tested on LOX/LH2.
| Applications | Titan 3 & 4 stage 1 |
| First Flown | 1968 Titan 3, 1989 Titan 4 |
| Dry Mass | 2281 kg (paired), 758 kg (single) |
| Length | 3.84 m to top of thrust structure, 3.13 m to top of
turbopump assembly |
| Maximum Diameter | 1.14 m |
| Mounting | fixed pair |
| Engine Cycle | gas generator |
| Oxidizer | nitrogen tetroxide at 540.7 kg/sec |
| Fuel | Aerozine-50 at 284 kg/sec |
| Mixture Ratio (O/F) | 1.91:1 |
| Thrust | 2437.5 kN vacuum paired |
| Isp | 301 sec vacuum |
| Expansion Ratio | 15:1 |
| Combustion Chamber Pressure | 58.3 atm |
| Cooling Method | fuel regenerative & ablative skirt |
| Burn Time | ( 200 secs |
| Applications | Titan 3 & 4 stage 2 |
| First Flown | late 1968 Titan 3, 1989 Titan 4 |
| Dry Mass | 589 kg |
| Length | 281 cm |
| Maximum Diameter | 163 cm (skirt outer diameter) |
| Mounting | fixed, but turbine exhaust utilized for roll
control |
| Engine Cycle | gas generator |
| Oxidizer | nitrogen tetroxide at 97 kg/sec |
| Fuel | Aerozine-50 at 54.7 kg/sec |
| Mixture Ratio | 1.86:1 |
| Thrust | 467 kN vacuum |
| Isp | 316 sec vacuum |
| Expansion Ratio | 49.2:1 |
| Combustion Chamber Pressure | 58.5 atm |
| Cooling Method | fuel regenerative thrust chamber, with
separate ablative skirt |
| Burn Time | ( 247 secs |
The Transtar system (developed by Aerojet) was an upper stage engine
using injector(s), chamber, and nozzle derived from the OMS system. These
propellants are pump-fed which increase chamber pressure and Isp. They
also permit the use of low-pressure lightweight tanks.
| Applications | upper stage |
| First Flight | not flown |
| Dry Mass | 76 kg |
| Length | 127 cm |
| Mounting | gimballed ( 10 degrees by two electromechanical
actuators |
| Propellants | nitrogen tetroxide and MMH |
| Mixture Ratio | 1.8:1 |
| Thrust | 16.68 kN vacuum |
| Isp | 328 secs vacuum |
| Expansion Ratio | 132:1 |
| Combustion Chamber Pressure | 23.8 atm |
| Cooling Method | fuel regenerative for chamber, radiative
for extension |
| Burn Time | not available, 15 starts |
The company also produces four bipropellant thrusters of varying power
settings.
Principle Characteristics
| Thrust | 2.00 | 21.35 | 62 | 445 |
| Isp | 265 | 285 | 287 | 309 |
| Dry
Mass(kg) | 0.27 | 0.57 | 1.13 | 1.86 |
| Exp. Ratio | 150 | 150 | 75 | 150 |
| Mixture
Ratio(O/F) | 1.65 | 1.60 | 1.65 | 1.65 |
ASI W9800149r1.1.
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