Spacecraft Reference Engines

(Last Updated 11 October 2012)

Notes:

Flight Certification/Man Rating: For engines where I have been unable to find a clear date listed in the documentation – I use their first unmanned launch as the flight certification; and for man-rating, I use MSFC’s traditional “two launches to manrate” criteria; though there are exceptions to this (the LEM and Space Shuttle).

Rocketdyne LR89-NA-7

Propellants: LOX/RP-1
O/F Ratio: 2.28
Thrust (sl): 330,00 lbf at 254 ISP
Chamber Pressure: 578 psia
Expansion Ratio (Ae/At): 8
Notes: Used on Atlas MA-5.
References:
NASA SP-8120 Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

Aerojet LR87-AJ-3 (Titan I Stage I Engine)

Thrust (sl): 300,000 lbf
References:
Titan II: A History of a Cold War Missile Program by David K. Stumpf

Aerojet LR91-AJ-3 (Titan I Stage II Engine)

Thrust (vac): 80,000 lbf
References:
Titan II: A History of a Cold War Missile Program by David K. Stumpf

Aerojet LR87-AJ-5 (Titan II Stage I Engine)

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.93
Thrust (sl): 430,000 lbf at 263 ISP
Thrust (vac): 473,800 lbf at 289 ISP
Chamber Pressure: 783 psia
Expansion Ratio (Ae/At): 8
References:
NASA SP-8120 Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)
Titan II: A History of a Cold War Missile Program by David K. Stumpf

Aerojet LR91-AJ-5 (Titan II Stage II Engine)

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.80
Thrust (sl): 100,000 lbf at 321 ISP
Chamber Pressure: 827 psia
Expansion Ratio (Ae/At): 49.2
Notes: Largely identical to the LR87-AJ-5 with the following distinguishing features:
References:
NASA SP-8120 Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)
Gemini Design Certification Report: GLV Section (19 February 1965)
Titan II: A History of a Cold War Missile Program by David K. Stumpf

Aerojet LR87-AJ-7 (Titan II GLV Stage I Engine)

Notes: Uses the basic Titan II (-5) engine as a building block for the special GLV version with added safety features and qualifications.
References:
Gemini Design Certification Report: GLV Section (19 February 1965)

Aerojet LR91-AJ-7 (Titan II GLV Stage II Engine)

Notes: Uses the basic Titan II (-5) engine as a building block for the special GLV version with added safety features and qualifications.
References:
Gemini Design Certification Report: GLV Section (19 February 1965)

Aerojet LR87-AJ-11 (Titan IIIE Stage I Engine)

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.915
Rated Thrust (vac): 520,000 lbf at 301.1 ISP
Expansion Ratio (Ae/At): 15
Operating Cycle: 165 seconds
Notes: Air starts after SRM burnout, hence why it has vacuum thrust values given.
References:
Titan IIIE/Centaur D-1T Systems Summary (September 1973)

Aerojet LR91-AJ-11 (Titan IIIE Stage II Engine)

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.79
Rated Thrust (vac): 101,000 lbf at 318.7 ISP
Expansion Ratio (Ae/At): 49.2
Operating Cycle: 225 seconds
References:
Titan IIIE/Centaur D-1T Systems Summary (September 1973)

Rocketdyne LR79-NA-11

Propellants: LOX/RP-1
O/F Ratio: 2.15
Thrust (sl): 169,500~ lbf at 256 ISP
Dry Weight: 2,166 lbs
T/W Ratio (sl): 78.25
Chamber Pressure: 588 psia
Expansion Ratio (Ae/At): 8
Diameter: 49 inches
Notes: Used on Thor missile.
References:
Design Information Report for the Thor YLR-79-NA-13 Main Engine and LR101-NA-11 Vernier Engines (9~ MB PDF)
NASA SP-8120 Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

UA-1205 SRM

Diameter: 120 inches
Segments:
5
Propellants:
UTP 3001 PBAN
Casing Material: D6AC Steel (195 ksi)
Rated Thrust: 1,200,000 lb at 266 ISP
Burn Time: 117~ seconds
Forward Closure Segment: 95” long with 38,150 lbs of PBAN
5 x Center Segments: 10 foot segment with 73,250 lbs of PBAN each
Aft Closure Segment: 64” long with 19,917 lbs of PBAN
Total Propellant: 424,317 lbs
Burnout Mass: 71,535 lbs
Notes: Figures are for a single SRM. Has a liquid injected thrust-vectoring system. Originally had a thrust termination system provided for in the original designs, but this was deleted from the unmanned designs.
(UA-1205 Thrust/Time Curve)
(UA-1205 General Arrangement Drawing)
(UA-1205 Forward Closure Segment Drawing)
(UA-1205 Aft Closure Segment Drawing)
(UA-1205 SRM Nozzle Drawing)
(UA-1205/1207 Components)
(UA-1205 General Arrangement)
(UA-1205/1207 General Arrangement)
(UA-1205/1207 Pressure/Time Curves)
(UA-1205/1207 Thrust/Time Curves)
(UA-1205/1207 Detailed Weight Statement)
(UA-1205 Forward Closure – Stock)
(UA-1205 Forward Closure – Thrust Termination Ports Added Back)
(UA-1205 Center Segment)
References:
Titan IIIE/Centaur D-1T Systems Summary (September 1973)
A Study of Performance and Cost Improvement Potential of the 120-In.- (3.05M) Diameter Solid Rocket Motor Volume II (December 1971)

UA-1207 SRM

Diameter: 120 inches
Segments:
7
Propellants:
UTP 3001B PBAN
Casing Material: D6AC Steel (195 ksi)
Rated Thrust: 1,460,300 lb at 269.5 ISP
Burn Time: 125~ seconds
Forward Closure Segment: 135” long with 60,932 lbs of PBAN
7 x Center Segments: 10 foot segment with 73,155 lbs of PBAN each
Aft Closure Segment: 64” long with 19,840 lbs of PBAN
Total Propellant: 592,857 lbs
Burnout Mass: 89,910 lbs
Notes: Due to this being developed for the Gemini B/MOL program, the forward closure segment contains two 33” thrust termination ports.
(UA-1207 Thrust/Time Curve)
(UA-1205/1207 Components)
(UA-1207 General Arrangement)
(UA-1205/1207 General Arrangement)
(UA-1205/1207 Pressure/Time Curves)
(UA-1205/1207 Thrust/Time Curves)
(UA-1205/1207 Detailed Weight Statement)
(UA-1207 Center Segment)
(UA-1207 Aft Closure)
References:
Titan IIIE/Centaur D-1T Systems Summary (September 1973)
A Study of Performance and Cost Improvement Potential of the 120-In.- (3.05M) Diameter Solid Rocket Motor Volume II (December 1971)

Pratt & Whitney RL10A-1

Ready (Flight Certified): November 1961
Propellants: LOX/LH2
Thrust (Vac): 15,000 lbf at 422 ISP
Chamber Pressure: 300 psia
Expansion Ratio: 40
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Rocketdyne H-1 (SA-201 through SA-205)

Ready (Man Rated): April 1962 (SA-2)
Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.23
Thrust (sl): 200,000 lbf at 260.5 ISP
Dry Weight (Inboard Engines on Saturn I): 1,830 lbs
Dry Weight (Outboard Engines on Saturn I): 2,100 lbs
T/W Ratio (sl): 95.23 to 109.28
Expansion Ratio: 8
Length: 8.8 Feet (105.6”)
Diameter:
4.9 Feet (58.8”)
Notes: The H-1C was the inboard engine variant, while the H-1D was the outboard engine variant.
References:
(H-1 Specifications)

Rocketdyne H-1 (SA-206 and Subsequent)

Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.23
Thrust (sl): 205,000 lbf at 261 ISP
Dry Weight (Inboard/Outboard Engines): 2,100 lbs
T/W Ratio (sl): 97.61
Expansion Ratio: 8
Length: 8.8 Feet (105.6”)
Diameter:
4.9 Feet (58.8”)
Notes: The H-1C was the inboard engine variant, while the H-1D was the outboard engine variant.
References:
(H-1 Specifications)

Pratt & Whitney RL10A-3

Ready (Flight Certified): June 1962
Propellants: LOX/LH2
Thrust (Vac): 15,000 lbf at 427 ISP
Chamber Pressure: 300 psia
Expansion Ratio: 40
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Pratt & Whitney RL10A-3-1

Ready (Flight Certified): September 1964
Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (Vac): 15,000 lbf at 433 ISP
Weight: 306 lbs
T/W Ratio: 49.01
Chamber Pressure: 300 psia
Expansion Ratio: 40
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
(RL-10A-3-1 Specifications from ‘Design Report for RL10-A-3-1’ – 2~ MB PDF Extract)
(RL-10A-3-1 Dimensioned Drawing – 780 KB GIF)

Rocketdyne J-2

Ready (Man Rated): July 1966 (AS-203)
Propellants: LOX/LH2
O/F Ratio: 5.5
Thrust (vac): 230,000 lbf at 425 ISP
Thrust (sl): Unknown at 294 ISP
Weight: 3,492 lbs
T/W Ratio (vac): 65.86
Chamber Pressure: 717 psia
Expansion Ratio: 27.5
References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

Rocketdyne SE-7-1 (S-IVB Auxiliary Propulsion System Ullage Control Engine)

Ready (Man Rated): July 1966 (AS-203)
Propellants:
NTO (N2O4) / MMH
Thrust (vac):
72 lbf at 274 ISP
Notes: Pressure fed for simplicity.
References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)

Aerojet AJ10-137 (Apollo Service Propulsion System)

Ready (Man Rated): August 1966 (AS-202)
Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.6
Thrust (vac): 20,000 lbf at ISP
Weight: 650~ lbs (approximately) [2]
T/W Ratio (vac): 30.76
Chamber Pressure: 100 psia
Expansion Ratio: 62.5
Notes: Was designed for a service lifetime of 750 seconds with 50 restarts.
References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
[2]
Apollo Operations Handbook: Block II Spacecraft (SM2A-03-Block II) (15 April 1969)
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

Aerojet AJ10-138 (Titan IIIA/IIIC Transstage Engine)

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 2.0
Thrust (vac): 8,000 lbf at 302 ISP
Chamber Pressure: 105 psia
Expansion Ratio (Ae/At): 40
References:
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

Pratt & Whitney RL10A-3-3

Ready (Flight Certified): October 1966
Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (Vac): 15,000 lbf at 444 ISP
Weight: 301.19 lbs
T/W Ratio: 49.8
Chamber Pressure: 400 psia
Expansion Ratio: 57
Notes: Capable of three starts during a single mission, with each firing time being as long as 450 seconds. Engine life of 4,000 seconds.
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
(RL-10A-3-3 Specifications from ‘Design Report for RL10-A-3-3’ – 600~ kb PDF Extract)
(RL-10A-3-3 Dimensioned Drawing – 135~ kb GIF)

LMDE / Apollo Lunar Module Descent Engine)

Ready (Man Rated): January 1968 (Apollo 5) ← Cleared after only one flight (!)
Propellants: NTO (N2O4) / Aerozine-50
O/F Ratio: 1.6
Thrust (vac): 10,500 lbf, at 301 ISP (See Notes on Throttling)
Weight: 394 lbs (179~ kg)
T/W Ratio: 25~
Chamber Pressure: 116 psig design requirement, 103.4 psia at FTP.
Nozzle Exit Diameter:
63 inches
Expansion Ratio:
46
Throttling Notes: The LMDE operates in two regimes: Full Throttle Power (FTP) is approximately 94.2% of rated thrust (9,900 lbs), and the Throttling Regime goes from 12.2% of rated thrust to 65%~ rated thrust (1,280 lbs to 6,825 lbs). Due to severe throat erosion problems in the range between 65% and 92.5% of rated thrust, any throttle command above 65% is automatically commanded to full throttle power.
Notes: Rated for a service lifetime of 1,000 seconds, which is 90 seconds of acceptance testing and a 910 second duty cycle. In extreme cases, it can go to a service lifetime of 1,100 seconds, but this is dangerous as this will severely char the combustion chamber, and probably result in burn-through.
Notes II: Further development of this by TRW led to the TR-201 engine.
References:
Mechanical Design of the Lunar Module Descent Engine by Jack M. Cherne for TRW (LINK to webpage with it)
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
Apollo Operations Handbook: Lunar Module LM 10 and Subsequent: Volume I Subsystems Data (LMA790-3-LM 10 and Subsequent)

Rocketdyne RS-18 (LMAE / Apollo Lunar Module Ascent Engine)

Ready (Man Rated): January 1968 (Apollo 5) ← Cleared after only one flight (!)
Propellants: NTO (N2O4) / Aerozine-50
O/F Ratio: 1.6
Thrust (vac): 3,500 lbf at 310 ISP
Weight: 180~ pounds
T/W Ratio (vac): 19.44
Chamber Pressure: 120 psia
Length: 47 inches
Nozzle Exit Diameter: 34 inches
Expansion Ratio:
45.6
Notes: Pressure fed for simplicity.
References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
Apollo Operations Handbook: Lunar Module LM 10 and Subsequent: Volume I Subsystems Data (LMA790-3-LM 10 and Subsequent)
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

Aerojet R-4D (Apollo Lunar Module and Service Module RCS)

Propellants: NTO (N2O4) / Aerozine-50
O/F Ratio:
2.03
Thrust (sl): 60 lbf at 168 ISP
Thrust (vac): 100 lbf at 280 ISP
Weight:
4.99 lbs
T/W Ratio (sl): 12.02
T/W Ratio (vac): 20.04
Chamber Pressure: 96.5 psia
Engine Length: 13.4 Inches
Nozzle Exit Diameter: 5.6 inches
Expansion Ratio:
40
Notes: 650 production engines made, of which 469 flew. Designed for a 1,000 second service life and 10,000 operational cycles. The maximum burn possible was 750 seconds.
References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
Apollo Operations Handbook: Block II Spacecraft (SM2A-03-Block II) (15 April 1969)
Apollo Operations Handbook: Lunar Module LM 10 and Subsequent: Volume I Subsystems Data (LMA790-3-LM 10 and Subsequent)
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

Aerojet R-4D-10

Aerojet R-4D-11

Aerojet R-4D-12

Aerojet R-4D-15 (HiPAT)

O/F Ratio: 1
Thrust (vac): 100 lbf at 329 ISP
Chamber Pressure: 137 psia
Expansion Ratio:
375
References:
Advanced Chemical Propulsion for Science Missions (NASA TM-2008-215069)

Rocketdyne SE-8 (Apollo Command Module RCS)

Propellants: NTO (N2O4) / MMH
O/F Ratio: 2.0
Thrust (vac):
93 lbf at 274 ISP
Chamber Pressure: 137 psia
Expansion Ratio (Ac/At):
9
Notes: Pressure fed for simplicity.
References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

Rocketdyne F-1

Ready (Flight Certified): 9 May 1967
Ready (Man Rated): April 1968 (Apollo 6)
Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.27
Thrust (sl): 1,522,000 lbf at 265.4 ISP
Thrust (vac): 1,748,200 lbf at 304.1 ISP
Weight: 18,616 lbs
T/W Ratio (sl): 81.75
T/W Ratio (vac): 93.90
Chamber Pressure: 982 pisa
Engine Length: 220.4 inches
Expansion Ratio: 16
Notes: Designed for a mission duration of 165 seconds, with a minimum acceptance firing of 495 seconds at Stennis. It was qualified however for 20 starts and 2,250 seconds of lifetime. 98 production engines were delivered, of which 65 were expended on Saturn V flights.
Cost: $2.08 million in FY64 (30 Mar 1964 contract for 76 x F-1 engines at $158.4 million.)
References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
(F-1/F-1A Comparison Sheet from Rocketdyne Circa 1993 for SEI Restart Studies – 473 kb GIF)

Rocketdyne F-1A

Ready: 1970s
Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.27
Thrust (sl): 1,800,000 lbf at 269.7 ISP
Thrust (vac): 2,020,700 lbf at 303.1 ISP
Weight: 19,000 lbs
T/W Ratio (sl): 94.73
T/W Ratio (vac): 106.35
Chamber Pressure: 1,161 psia
Engine Length:
220.4 inches
Expansion Ratio:
16
Cost: In the 1990s, Rocketdyne estimated that a F-1A Restart program would cost $315 million in FY92 dollars in non-recurring costs to restart production and re-certify the engine. Recurring costs would have been $1,080 million in FY92 dollars for 72 engines at an average cost of $15m FY92 dollars per engine, with deliveries over a five year period. Deliveries would have commenced four years after authority to proceed, with a peak delivery rate of 16 engines per year.
Resources:
(F-1/F-1A Comparison Sheet from Rocketdyne Circa 1993 for SEI Restart Studies – 473 kb GIF)
(F-1A Restart Costing and Schedule – 175 kb GIF)
Advanced Transportation System Studies: Technical Area 3: Alternate Propulsion Subsystem Concepts (RI/RD 93-123-3) (April 1993)
Advanced Transportation System Studies – Technical Area 2: Heavy Lift Launch Vehicle Development Contract NAS8-39208 DR 4 Final Report – July 1995 (LINK to image of table found on page 3-3, LINK to image of table on page 2-6)

Rocketdyne J-2S (J-2 Simplified)

Ready: 1970s
Propellants: LOX/LH2
O/F Ratio: 5.5 (can operate at 5.0 and 4.5 upon command for optimum propellant utilization)
Thrust (vac): 265,000 lbf at 436 ISP
Thrust (sl): 197,000 lbf at 320 ISP
Weight: 3,800 lbs
T/W Ratio (vac): 69.73
T/W Ratio (sl): 51.84
Chamber Pressure: 1,200 PSIA
Expansion Ratio:
40
Notes: It was intended to provide performance upgrades for the J-2 and to also simplify the production and operation of the engine. Much of the original J-2 design team worked on the J-2S effort. It was designed with all engine interfaces as such so that it could be a direct “drop-in” replacement for the J-2.
The engine cycle was changed to a tap-off cycle to eliminate the gas generator. Throttling capability was added as an option for applications other than the Saturn Program. The engine also included a feature for low thrust operation known as “Idle Mode” which was to be used for propellant tank settling, on orbit maneuvering, and rapid engine chill-down prior to firing. Two versions were developed; one for the S-II with single-start capability, and one for the S-IVB with three-start capability.
This engine system was validated with 6 flight configuration engines in 273 tests for a total operating experience of 30,858 seconds. Upon the termination of the J-2S program, the engine was ready to go into certification for flight operations.
The technology developed for the J-2S did not go to waste, however, as the Mk-29 turbopump developed for it was later used by Rocketdyne on the Linear Aerospike Engine Program.
Cost: It was estimated in the 1990s by Rocketdyne that it would require $245 million in FY92 dollars in non-recurring costs to restart production and re-certify the engine. Average engine costs were undetermined, as it depended on production rate. (see Costing image below).
Resources:
(J-2S Sheet from Rocketdyne Circa 1993 for SEI Restart Studies – 579 kb GIF)
(J-2S Restart Costing and Schedule – 175 kb GIF)
Advanced Transportation System Studies: Technical Area 3: Alternate Propulsion Subsystem Concepts (RI/RD 93-123-3) (April 1993)
Advanced Transportation System Studies: Technical Area 3: Alternate Propulsion Subsystem Concepts (RD00-164-1) (April 2000)
Advanced Transportation System Studies – Technical Area 2: Heavy Lift Launch Vehicle Development Contract NAS8-39208 DR 4 Final Report – July 1995 (LINK to image of table found on page 3-3, LINK to image of table on page 2-6)

Kuznetsov NK-33

Ready: 1970s
Burn Time: 600 seconds
Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.8
Thrust (sl): 339,506 lbf ( kgf) at 297 ISP
Thrust (vac): 368,237 lbf ( kgf) at 331 ISP
Weight: 2,694 lbs (1,222 kg)
T/W Ratio (sl): 126.02
T/W Ratio (vac): 136.68
Diameter: 1.5 meters
Expansion Ratio: 27
Chamber Pressure: 2,113.19 PSI (145.70 bar)
Notes: Developed for the N-1F Stage 1 (Blok A).

Kuznetsov NK-43

Ready: 1970s
Burn Time: 600 seconds
Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.8
Thrust (sl): 280,517 lbf ( kgf) at 246 ISP
Thrust (vac): 394,539 lbf ( kgf) at 346 ISP
Weight: 3,077 lbs (1,396 kg)
Chamber Pressure: 2,113.19 PSI (145.70 bar)
T/W Ratio (sl): 91.16
T/W Ratio (vac): 128.21
Diameter: 2.5 meters
Expansion Ratio: 70
Notes: Developed for the N-1F Stage 2 (Blok B).

Aerojet-General M-1

Ready: 1970s
Propellants: LOX/LH2
O/F Ratio: 6
Thrust (sl): 868,851 lbf (394,100 kgf) at 310 ISP
Thrust (vac): 1,199,575 lbf (544,118 kgf) at 428 ISP
Weight: 19,991.5 lbs (9,068 kg)
T/W Ratio (sl): 43.46 units of thrust per unit of mass of engine
T/W Ratio (vac): 60 units of thrust per unit of mass of engine
Notes: Was canceled in 1965, due to the general draw down of NASA and the elimination of post-Apollo missions and hardware development.

NERVA I (1972 Baseline)

Ready: Late 1970s
Operational Time: 600 minute minimum operational time at rated chamber temperature accumulated in up to 60 cycles.
Propellants: Nuclear/LH2 (1,570~ MWt core)
Thrust (vac): 75,000 lbf at 825 ISP
Weight: 27,728 lbm [1] or 25,803 lbm [2]
T/W Ratio (vac): 2.7 [1] or 2.9 [2]
Chamber Pressure: 450 psia
Chamber Temperature: 4,250 deg R
Engine Length: 398 inches
Nozzle Diameter: 116 inches
Notes: The design requirements for a flight rated NERVA engine were for it to be able to be stored for 5 years on the ground, 6 months on the pad, and 3 years in space.
(NERVA I Dimensional Envelope)
[1] SD 71-466-4: Nuclear Flight System Definition Study. Phase III Final Report, Volume II, Part C. (April 1971)
[2] NASA Technical Memorandum 105252: An Overview of Tested and Analyzed NTP Concepts (September 1991)

NERVA II (1968 Baseline)

Ready: Very Late 1970s, very early 1980s
Propellants: Nuclear/LH2 (4,000 MWt core)
Thrust (vac): 195,000 lbf at 850 ISP and a flow rate of 224 lb/sec.
Weight: 28,530 lbs (25,540 lbm for bare engine, 1,940 lbm for light shield, and 1,050 lbm for thrust structure)
T/W Ratio (vac): 6.83
Chamber Pressure: 625 psia
Chamber Temperature: 4,500 deg deg R
Engine Length: 481.18 inches
Nozzle Diameter: 162.8 inches
Notes: This engine essentially died out around 1966-1968, as technological advancements raised the possibility of an advanced NERVA I with a 2,000+ MWt core with ISPs around 900, allowing the smaller NERVA I to do virtually all the tasks that the larger NERVA II was needed for. All the data in this entry is derived from D2-113544-4 “Integrated Manned Interplanetary Spacecraft Concept Definition – Volume IV” (January 1968).
(NERVA II Dimensional Envelope)
(NERVA II Start-up/Shut-down Graphs)
(NERVA II Artwork with projected range of specifications – November 1968)

Rocketdyne RS-25 Space Shuttle Main Engine (SSME)

Ready (Man Rated): April 1981 (at 100% RPL) 1983 (at 104% RPL), 1998 (at 104.5% RPL)
Cycle: Staged Combustion
Propellants: LOX/LH2
O/F Ratio: 6
100% RPL Thrust (sl): 373,500 lbf (169,417~ kgf) at 362 ISP
100% RPL Thrust (vac): 468,400 lbf (212,463~ kgf) at 454 ISP
104% RPL Thrust (sl): 390,000 lbf (176,901 kgf) at 364.8 ISP
104% RPL Thrust (vac): 488,800 lbf (221,716~ kgf) at 452.9 ISP
Weight: 6,990 lbs (3,170.6 kg) in 1995, 7,748 lbs (3,514.4 kg) in 2011’s Block II.
Chamber Pressure: 3,140 PSIA at 104% RPL for Block I SSME; 2,870 PSIA at 104.5% RPL for Block IIA/II SSME.
T/W Ratio (sl):
55.793 lbf per lb for 1995
T/W Ratio (vac): 69.928 lbf per lb for 1995
Engine Length: 14 feet
Engine Diameter: 8 feet
Expansion Ratio: 77.5
Notes: SSME data is from July 1995 tables in Advanced Transportation System Studies – Technical Area 2 – Heavy Lift Launch Vehicle Development.
The following SSME models have been identified along with their first flights:
RS-25 Phase I First Manned Orbital Flight (FMOF) SSME: Used on STS-1, -2, -3, -4, and -5. Certified to 100% RPL.
RS-25 Phase I Full Power Level (FPL) SSME: First Flight STS-6 in April 1983. Tests by 1983 showed that SSME lacked margin to safely operate at full power level (FPL), so development of certified thrust past 104% RPL was halted, though 109% FPL remained available for emergency abort contingencies from this point onwards in the SSME program.
RS-25A Phase II Return-to-Flight SSME: First Flight STS-26R in September 1988. Re-certified to 104% RPL following post-Challenger safety changes.
RS-25B Block I SSME: First Flight STS-70 in July 1995. Redesigned two-duct powerhead, heat exchanger, and HP Oxidizer turbopump (HPOTP). Changes were made to the thermocoupler design after STS-70, and the slightly redesigned engine flew for the first time again on STS-75 in February 1996.
RS-25B Block IA SSME: First flight STS-73 in October 1995. Modifications centered around main injector.
RS-25C Block IIA SSME: First flight STS-89 in January 1998. Last flew on STS-109 in March 2002. Interim configuration used with new Large Throat Main Combustion Chamber (LTMCC) while waiting for the Block II HP Fuel Turbopump (HPFTP) to be certified. LTMCC reduced operating pressures and temperatures across the board for subcomponents by 10%, dramatically improving engine reliability. Other changes included Block II LP Oxidizer and Fuel turbopumps (LPOTP/LPFTP). First engine certified to higher 104.5% RPL to support ISS operations. Minor changes after the first flight were made, which involved opening up BLC holes to minimize faceplate erosion, and the modified design first flew on STS-96 in May 1999.
RS-25D Block II SSME: First flight STS-104 in July 2001. Incorporated Block II HP Fuel Turbopump (HPFTP), which along with LTMCC, made it twice as reliable as prior SSMEs.
RS-25E Minimal Change Expendable SSME: 2005 proposal for simplified SSME. Designation continues to be used as a stand in for expendable SSMEs to differentiate them from existing RS-25Ds.
RS-25F Low Cost Manufacture Expendable SSME: 2005 proposal for a even cheaper expendable SSME which simplifies almost everything.
References:
MSFC and Exploration: Our Path Forward, September 2005 Presentation (LINK to slide showing SSME variant designations)
AIAA 2002-3758: Space Shuttle Main Engine (SSME) Options for the Future Shuttle (LINK to image of table listing SSME variants)
Space Shuttle Main Engine: Relentless Pursuit of Improvement by Katherine Van Hooser; 27 January 2011 presentation, in particular the following slides:
            SSME Final Specs as of January 2011
            SSME Timeline
            SSME Phase II Specs
            SSME Block I Specs
            SSME Block IIA Specs
            SSME Block II Specs
            SSME AHMS Information
            SSME Failure Probabilities
Space Shuttle Main Engine – Thirty Years Of Innovation by Fred H. Jue (LINK)
Atlantis STS-104 SSME Flight Readiness Review, 28 June 2001 (2.1 MB PDF)

Aerojet AJ10-190 (Shuttle OMS System)

Ready (Man Rated): April 1981
Propellants: NTO (N2O4) / MMH
O/F Ratio: 1.65
Thrust (vac): 6,000 lbf at 313.2 ISP
Weight: 297 lbs
T/W Ratio (vac): 20.2
Chamber Pressure: 125 psia
Engine Length: 77 inches
Engine Nozzle Exit Diameter: 46 inches
Expansion Ratio: 55
Notes: Can be used in 100 missions. To this extent, it can start 1,000 times and run for a total of 15 hours of firing time before replacement. Maximum allowed firing length was 1,250 seconds against a typical deorbit burn of 150 to 250 seconds.
References:
Space Shuttle Propulsion Systems
by Russell Bardos, Office of Space Flight (26 June 1990) (PDF)

Marquardt / Aerojet R-40A (Shuttle Reaction Control System)

Ready (Man Rated): April 1981
Propellants: NTO (N2O4) / MMH
O/F Ratio: 1.6
Thrust (vac): 870 lbf at 280 ISP (at 22 Expansion Ratio)
Weight: 16 lbs
T/W Ratio (vac): 54.375
Chamber Pressure: 152 psia
Expansion Ratio: 22 to 30 (depending on whether it’s a long, short or no scarf configuration).
Notes: Designed for a lifetime of 100 missions or 20,000 cycles and a total firing duration of 12,800 seconds.
References:
Space Shuttle Propulsion Systems
by Russell Bardos, Office of Space Flight (26 June 1990) (PDF)

Marquardt / Aerojet R-1E (Shuttle Vernier Reaction Control System)

Ready (Man Rated): April 1981
Propellants: NTO (N2O4) / MMH
O/F Ratio: 1.65
Thrust (vac): 24 lbf at 265 ISP
Weight: 9.4 lbs
T/W Ratio (vac): 2.55
Chamber Pressure: 110 psia
Expansion Ratio: 20.7
Notes: Mission limited by how long the Columbium/Titanium chamber lasts. Officially it’s rated for 330,000 cycles and a total firing duration of 125,000 seconds.
References:
Space Shuttle Propulsion Systems
by Russell Bardos, Office of Space Flight (26 June 1990) (PDF)

Pratt & Whitney RL10A-3-3A

Ready (Flight Certified): November 1981
Propellants: LOX/LH2
O/F Ratio: 5.5
Thrust (Vac): 16,500 lbf at 444.4 ISP
Chamber Pressure: 475 psia
Expansion Ratio: 61
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

RD-0120 (РД-0120, Ракетный Двигатель-0120)

Ready (Flight Certified): May 1987 (Energia First Flight)
Propellants
: LOX/LH2
O/F Ratio: 6
Thrust (sl): 352,746 lbf (160,000 kgf) at 364 ISP
Thrust (vac): 440,925 lbf (200,000 kgf) at 455 ISP
Weight: 7,607 lbs (3,450~ kg)
Chamber Pressure: 3,000 psia
Engine Length:
14.93 feet
Engine Diameter:
7.94 feet
Expansion Ratio:
85.7
References:
Advanced Transportation System Studies – Technical Area 2: Heavy Lift Launch Vehicle Development
Contract NAS8-39208 DR 4 Final Report – July 1995 (LINK to image of table found on page 3-3, LINK to image of table on page 2-6)

RD-170 (РД-170, Ракетный Двигатель-170)

Ready (Flight Certified): May 1987 (Energia First Flight)
Propellants
: LOX/Syn10
O/F Ratio: 2.6
Thrust (sl): 1,632,000 lbf (740,262 kgf) at 309 ISP
Thrust (vac): 1,777,000 lbf (806,033 kgf) at 337 ISP
Weight: 21,510 lbs (9,756.7~ kg)
Chamber Pressure: 3,560 psia
T/W Ratio (sl): 75.87
T/W Ratio (vac): 82.61
Engine Length:
13.12 feet
Engine Diameter:
12.2 feet
Expansion Ratio:
36.87
References:
Advanced Transportation System Studies – Technical Area 2: Heavy Lift Launch Vehicle Development
Contract NAS8-39208 DR 4 Final Report – July 1995 (LINK to image of table found on page 3-3, LINK to image of table on page 2-6)

Thiokol/ATK Space Shuttle Solid Rocket Motor (SRM)

Ready (Man Rated): April 1981 (STS-1)
Case Material: D6AC Steel
Propellant: TP-H1148 PBAN
Notes: Original SRB design. Flown on STS-1 (1981) through STS-7 (1983).
(1977 Diagram Showing Split between SRM and SRB – 92~ KB GIF)
References:
Reusable Solid Rocket Motor—Accomplishments, Lessons, and a Culture of Success by Dennis R. Moore and Willie J. Phelps
NASA TN D-8511 Materials and Processes for Shuttle Engine, External Tank, and Solid Rocket Booster (June 1977)

Thiokol/ATK Space Shuttle High-Performance Solid Rocket Motor (HP-SRM aka HPM)

Ready (Man Rated): August 1983 (STS-8)
Propellant:
TP-H1148 PBAN
Thrust at 1.2 seconds (vac): 3,159,000~ lbf
Thrust at 20 seconds (vac): 3,330,000~ lbf (roughly the highest obtained)
Length:
126 ft
Diameter:
146 inches
Loaded Mass:
1,255,750 lbs
Propellant Mass:
1,110,000 lbs
Empty Mass:
145,750 lbs
Notes: Performance upgraded SRM, which had increased motor chamber pressure, reduced nozzle throat diameter, increased nozzle expansion ratio and changes to the propellant grain pattern to modify the thrust/time history. These changes resulted in a 3-second increase in ISP, and an additional 3,000 lbs of payload. Catastrophically failed on Challenger. Detailed thrust/pressure levels extracted from TM-86561 data. Raw data available in Excel format (HERE) and in graph format (HERE).
References:
Reusable Solid Rocket Motor—Accomplishments, Lessons, and a Culture of Success by Dennis R. Moore and Willie J. Phelps
Block II SRM Conceptual Design Studies Final Report: Conceptual Design Package: Volume I, Book 1 – Morton Thiokol (19 Dec 1986)
NASA TM-86561 Shifts in Shuttle SRM Performance because of Ammonium Perchlorate Chrystal Shape on Missions 51-I/J and 61-A/B. (August 1986)

Thiokol/ATK Space Shuttle Filament Wound Case Solid Rocket Motor (FWC-SRM)

Ready: July 1986~ (Cancelled due to Challenger Disaster, see Notes)
Length:
1,513.38 inches (motor itself)
Diameter: 150 inches
Notes: Design work began May 1982 as part of long range performance improvement plans for the Shuttle. It would have been used in place of the HP-SRM for special missions requiring increased performance. In place of conventional steel casings, graphite/epoxy filament wound casings would be used, resulting in a mass reduction of 25,000~ lb of the motor’s dry weight, increasing shuttle payload by about 4,600~ pounds. The full scale FWC qualification motor (QM-5) was assembled and ready for firing when Challenger exploded, and the first flight motors were stacked at Vandenberg AFB, for the July 1986 flight of STS-62A.
References:
(FWC-SRM Diagram and Perspective Cutaway Drawing – 209~ kb GIF)
(Image of DM-6 or QM-4 FWC-SRM before test firing – 432~ kb JPG)
Space Shuttle Filament Wound Case Solid Rocket Motor Static Test Results (DM-6) by C.A. Saderholm
Reusable Solid Rocket Motor—Accomplishments, Lessons, and a Culture of Success by Dennis R. Moore and Willie J. Phelps

Thiokol/ATK Space Shuttle Redesigned Solid Rocket Motor (RSRM)

Preliminary Requirements Review: July-August 1986
Preliminary Design Review:
September 1986
Critical Design Review:
October 1987
Ready (Man Rated): September 1988 (STS-26R)
Average Thrust (vac): 2,590,000 lbf at 267.9 ISP
      Thrust at ½ seconds (vac): 2,760,000~ lbf
      Thrust at 1 second (vac): 3,220,000~ lbf
      Thrust at 20 seconds (vac): 3,370,000~ lbf (roughly the highest obtained)
Chamber Pressure (Average) 625 psia
Chamber Pressure (Highest): 925~ psia at around 0.8 to 1 seconds after ignition
Motor Weight (Loaded): 1,255,978 lb
Booster Weight (Loaded): 1,300,000~ lbs [1]
Booster Weight (Burnout): 192,000 lbs [1]
Propellant Weight: 1,107,169 lb
Propellant Mass Fraction: 0.882
Total Inert Weight: 148,809 lbs
      Case Weight: 98,740 lbs in four D6AC steel-cased segments using Abestos/NBR insulation
      Nozzle Weight: 23,965 lbs
Length (Motor): 1,513.382 inches
Length (Booster): 1,800 inches
Diameter: 146 inches
Expansion Ratio: 7.72
Notes: Is sometimes called Reusable Solid Rocket Motor in an attempt to pretend Challenger didn’t happen. Interestingly, the manufacture of RSRM test/flight hardware began before the Preliminary Design Review; proving that NASA can move fast when it needs to. Detailed thrust/pressure levels extracted from STS-35 data. Raw data available in Excel format (HERE) and in graph format (HERE).
References:
[1] Solid Rocket Booster (SRB) - Evolution and Lessons Learned During the Shuttle Program
National Space Transportation System: Overview (September 1988) (NASA)
From Earth to Orbit: An Assessment of Transportation Options (1992) (National Research Council)
Reusable Solid Rocket Motor—Accomplishments, Lessons, and a Culture of Success by Dennis R. Moore and Willie J. Phelps
Space Shuttle Propulsion Systems by Russell Bardos, Office of Space Flight (26 June 1990) (PDF)
RSRM-11 (360W011) Final Report: Ballistics Mass Properties (STS-35) (21 January 1991)

Thiokol/ATK Space Shuttle Advanced Solid Rocket Motor (ASRM)

Ready: Late 1990s
Propellant: HTPB
Average Thrust (vac): 2,636,600 lbf at 269.18 ISP
Motor Weight (loaded): 1,350,381 lb
Propellant Mass Fraction: 0.895
Total Inert Weight: 142,313 lbs
      Case Weight: 99,442 lbs in three Ultra-High Strength steel-cased segments using Kevlar/Glass/EPDM insulation.
      Nozzle Weight: 18,217 lbs
Length: 1,513 inches
Diameter: 150 inches
Expansion Ratio: 7.54
Notes: Program ran from 1986 to 1993 and consumed about $2.2 billion before being cancelled. The cases were made out of 9Ni-4 Co-0.3C steel.
References:
From Earth to Orbit: An Assessment of Transportation Options (1992) (National Research Council)
Space Shuttle Propulsion Systems by Russell Bardos, Office of Space Flight (26 June 1990) (PDF)

Space Transportation Main Engine (STME)

Ready: 1990s
Cycle: Gas Generator
Propellants: LOX/LH2
O/F Ratio: 6
Thrust (sl): 551,430 lbf at 364 ISP
Thrust (vac): 650,000 lbf at 428.5 ISP
Weight: 9,974 lbs
T/W Ratio (sl): 55.28
T/W Ratio (vac): 65.16
Chamber Pressure: 2,250 PSIA
Engine Length:
161 inches
Nozzle Diameter: 96 inches
Expansion Ratio:
45
Notes: Was designed as a stripped down SSME-heritage expendable engine to support the propulsion requirements of the National Launch System (NLS) by a consortium of Aerojet, Pratt & Whitney, and Rocketdyne. It’s 500th unit cost goal was $5.3 million per engine [1].
References:
[1] From Earth to Orbit: An Assessment of Transportation Options (1992) (National Research Council)
Advanced Transportation System Studies – Technical Area 2 – Heavy Lift Launch Vehicle Development Contract NAS8-39208 DR 4 Final Report – July 1995 (LINK to image of table found on page 3-3, LINK to image of table on page 2-6)
(STME Data Excerpt from Alternate Propulsion Subsystem Concept Database v1.3 from 5 April 1993 – 441 kb PDF)

Aerojet-General M-1A

Ready: 1990s?
Propellants: LOX/LH2
O/F Ratio: 5
Thrust (sl): 1,300,000 lbf at 334.5 ISP
Thrust (vac): 1,562,000 lbf at 414 ISP
Weight: 20,200 lbs (9,162.56~ kg)
T/W Ratio (sl): 64.35
T/W Ratio (vac): 77.32
Chamber Pressure: 1,000 PSIA
Engine Length:
19.08 feet
Engine Diameter:
12.58 feet
Expansion Ratio: 20
Notes: Proposed uprated version of M-1 studied during the SEI phase of the 1990s.
References:
Advanced Transportation System Studies – Technical Area 2 – Heavy Lift Launch Vehicle Development Contract NAS8-39208 DR 4 Final Report – July 1995 (LINK to image of table found on page 3-3, LINK to image of table on page 2-6)

Pratt & Whitney XNR2000 CERMET NTR Family

Ready: Late 1990s to Early 2000s
Burn Time: 270> minutes of total burn time at rated thrust, single burn duration of 60 minutes maximum.
Propellants: Nuclear/LH2


Thrust Level (klbf):

25

50

75


Weight (lbm)

4,752

7,586

9,518


T/W Ratio:

5.26

6.59

7.88


ISP (Sec)

900

901

897


Chamber Pressure (psia):

766

735

836


Nozzle Exit Diameter (ft):

5.8

8.3

9.5


Stowed Engine Length (ft):

11

12.4

12


Deployed Engine Length (ft):

15.3

20.3

22.7


Number of Fuel Elements:

151

313

379


Fuel Element Temp. (K):

2,880

2,880

2,880

Notes: The XNR2000 family was designed around the expander cycle. Throttling could be done down to 25% of rated thrust.
References
Advanced Propulsion Engine Assessment based on a Cermet Reactor by Pratt & Whitney (October 1992) (1.5~ MB PDF)

Pratt & Whitney RL10A-4

Ready (Flight Certified): December 1990
Propellants: LOX/LH2
O/F Ratio: 5.5
Thrust (Vac): 20,800 lbf at 449 ISP
Chamber Pressure: 578 psia
Expansion Ratio: 84
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Pratt & Whitney RL10A-5

Ready (Flight Certified): August 1992
Propellants: LOX/LH2
Thrust (Vac): 14,560 lbf at 368 ISP
Chamber Pressure: 485 psia
Expansion Ratio: 43
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Pratt & Whitney RL10A-4-1

Ready (Flight Certified): February 1994
Propellants: LOX/LH2
Thrust (Vac): 22,300 lbf at 451 ISP
Chamber Pressure: 610 psia
Expansion Ratio: 84
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Pratt & Whitney RL10B-2

Ready (Flight Certified): May 1998
Propellants: LOX/LH2
O/F Ratio: 5.88
Thrust (Vac): 24,750 lbf at 465.5 ISP
Weight: 664 lbs
T/W Ratio (vac): 32.27
Chamber Pressure: 644 psia
Engine Length: 86.5” (Stowed), 163.5” (Deployed)
Nozzle Diameter: 84.5 inches
Expansion Ratio: 285
References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
(RL-10B-2 Cutaway image showing the engine with it’s Nozzle extension stowed)
(RL-10B-2 Dimensioned image by Russians showing it’s different expansion ratios)

Boeing NSTAR Ion Thruster

Ready (Flight Certified): October 1998 (Deep Space 1 Probe)
Size
: 30 cm diameter
Weight: 8 kilograms
Propellant: Electric/Xenon
Power Needed: 2.3 kilowatts at full thrust
Thrust (vac): 20 to 92 millinewtons at 3,100 ISP

ATK Space Shuttle 5-Segment Booster (FSB)

Ready: 2000s
Length: 2,120 inches
Weight: 5,964,430 lbs at ignition (roughly)
Thrust (sl): 3,799,000 lbf
T/W Ratio (sl): 1.57
Chamber Pressure (Average): 639 psia
ISP (vac): 264.7
Burn Time: 129.6 seconds, for 368 mlbf-sec total impulse
Expansion Ratio: 6.55
Notes: An additional center segment was added to the current 4-segment RSRB, along with a new nozzle to handle the increased mass flow rate. The nozzle had many features from the cancelled ASRM program.
References:
AIAA 2000-5070 Shuttle Upgrade Using 5-Segment Booster (FSB) – (19-21 September 2000)

ATK GEM-60 Strap-On SRM

Diameter: 60 inches
Length: 53 feet
Thrust (sl): 280,000 lbf
Propellant Mass: 65,472 lb
Notes: Developed to boost payload-to-orbit of the Delta IVM+ family. The first two motor configuration flew in November 2002 and the first four motor configuration flew in 2009.
References:
GEM-60 Press Brochure (400 kb PDF)

Rocketdyne XRS-2200 Linear Aerospike

Ready: 2000s
Propellants:
LOX/LH2
O/F Ratio:
5.5
Thrust (sl):
204,420 lbf (92,723~ kgf) at 339 ISP
Thrust (vac): 266,230 lbf (120,759~ kgf) at 436.5 ISP
Chamber Pressure: 854 PSIA
Expansion Ratio: 58
Notes: Three engines were built during the X-33 program and were tested at Stennis in single engine layouts.

RD-180 (РД-180, Ракетный Двигатель-180)

Ready (Flight Certified): May 2000 (Atlas III First Flight)
Propellants: LOX/RP-1
Thrust (sl): 860,568 lbf (390,347 kgf) at 311 ISP
Thrust (vac): 933,400 lbf (423,383 kgf) at 338 ISP
Weight: 12,081 lb (5,480 kg)
Chamber Pressure: 3,868 psia (26.7 MPa)
Notes: Two-Chamber Staged combustion engine. Used in U.S. EELV program, though all engines so far have been built in Russia.

Rocketdyne RS-68

Ready (Flight Certified): November 2002 (Unmanned Delta IV Launch)
Propellants: LOX/LH2
O/F Ratio: 6
Thrust (sl): 663,000 lbf (300,732 kgf) at 359 ISP
Thrust (vac): 758,000 lbf (343,823 kgf) at 409 ISP
Weight: 14,876 lbs (6,748 kg)
T/W Ratio (sl):
44.57 lbf per lb
T/W Ratio (vac): 50.95 lbf per lb
Notes: Developed as a cheap expendable engine building off the failed STME program. Has 80% fewer parts and requires 92% less labor than than the RS-25 SSME. It uses an ablative engine nozzle, which while being heavier than a regeneratively cooled nozzle, is much cheaper and faster to construct than a regenerative nozzle with it's hordes of tubing.

Rocketdyne RS-84

Ready: 2000-2010s
Propellants: LOX/RP-1
O/F Ratio: 2.7
Thrust (sl): 1,064,000 lbf at 304 ISP
Thrust (vac): 1,130,000 lbf at 324 ISP
Weight: 17,919 lbs
T/W Ratio (sl):
59.38 lbf per lb
T/W Ratio (vac): 63.06 lbf per lb
Notes: Developed as a reusable LOX/RP-1 engine with a lifetime of about 100 flights. Cancelled by NASA in 2005.

NASA Evolutionary Xenon Thruster (NEXT)

Ready: 2010s
Size: 40 cm diameter
Propellants: Electric/Xenon
Power Needed: 6.9 kilowatts
Thrust (vac): 0.05 lbf (0.24 newtons / 237 millinewtons) at 4,100 ISP
Fuel Density: 3,100 kg/m3 plus unknown amount of electric power.
Notes: Part of a project to develop high powered electric propulsion capabilities.

NASA High Power Electric Propulsion (HiPEP)

Ready: 2010s
Propellants: Electric/Xenon
Power Needed: 39.3 kilowatts
Thrust (vac): 0.15 lbf (0.67 newtons) at 9,620 ISP
Fuel Consumption: 7 milligrams/second at full power.
Fuel Density: 3,100 kg/m3
Notes: Was intended for use on the JIMO probe, which was cancelled in 2005.

Aerojet BPT-4000 Hall Effect Thruster

Ready (Flight Certified): August 2010 (USA-214 AEHF Military Satellite)
Burn Time: 6,000+ Hours
Propellants: Electric/Xenon
Thrust (vac): 0.06 lbf (0.27 newtons) at 1,950 ISP
Weight: 16.5 lbs
Fuel Density: 3,100 kg/m3 plus 4500 W at 350 V
T/W Ratio (vac): 0.0036 lbf per lb

Pratt & Whitney Rocketdyne J-2X

Ready: 2012-2020s
Propellants: LOX/LH2
O/F Ratio: 5.5
Thrust (vac): 294,000 lbf at 448 ISP
Weight: 5,450 lbs
T/W Ratio (vac): 53.94
Chamber Pressure: 1,337 psia
Expansion Ratio: 92
Engine Length: 185 inches (4.699 m)
Nozzle Exit Diameter: 120 inches (3.048 m)
Note: Had to meet much more stringent requirements, including a restart in space after 90~ days cold. Can restart eight times during a mission, for a total firing time of 2,600 seconds.
References:
Presentation by Tracy Lamm of P&W Rocketdyne from 26 Feb 2007 (1.8 MB PDF)

Ad Astra VX-200 VASMIR Engine Prototype

Ready: 2015-2020?
Propellants: Electric/Argon
Thrust (vac): 1.28 lbf (5.7 newtons) at 5,000 ISP and 120 mg/sec of Argon.
Fuel Density: 1,430 kg/m3 plus 200~ kW of DC electric power
Notes: A flight version, designated the VF-200 will fly on the ISS at some point in the future.
References:
AIAA-2011-1071:
Performance studies of the VASIMR® VX-200