Rocket Engine Specifications
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Aerojet Rocket Engines

Credits for providing information to Jim Glass
Boeing North American, and Jane's Space

AeroJet Propulsion Division

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)

Control Thrusters

Orbital Maneuvering System (OMS)

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

Titan Space Launcher Engines

A-J-5 Titan 2 Engines

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 Ratio1.80


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)
Length3.84 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

Aerojet Satellite and Attitude Engines

The company also produces four bipropellant thrusters of varying power settings.

Principle Characteristics

Dry Mass(kg)0.270.571.131.86
Exp. Ratio15015075150
Mixture Ratio(O/F)1.651.601.651.65

Rocket Engine Specifications

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