Spacecraft Reference Engines

(Last Updated 10 March 2014)

A Note Regarding References:

My stance on this has changed regarding the catastrophe that befell the NASA Technical Reports Server (NTRS) in May 2013 when at the behest of Congressman Frank Wolf (Idiot-VA), the entire archive was taken offline for “technical review” following a furor over the case of Bo Jiang, a Chinese national who did contract work for NASA, and was suspected of spying on America.

Hilariously enough, when Bo’s laptop was inspected after it was seized, inspectors did not find TOP SEKRIT information, but...pornography, lots of it, along with pirated software.

Before this incident, I had been more than happy to provide a name for the reference; e.g.

Boeing Integrated Manned Interplanetary Spacecraft Concept Definition – Volume IV (January 1968)

So that people could then look it up on their own via NTRS, as opposed to hosting dozens of 50+ MB and above files on my own web server.

Unfortunately, the Boeing IMIS studies disappeared in the Great Frank Wolf Debacle of 2013; and as of March 2014, are still not back on NTRS, despite claims of “75% of material restored”.

I suspect this is because the Boeing IMIS studies used NERVA II nuclear propulsion modules, and anything with “nuclear” attached to it is classified automatically in the post 9/11 paranoia environment.

Fortunately, I had the Boeing IMIS studies saved to my hard drive; so I will upload them at some future date as they are an important series of historical documents.

As a result of the Frank Wolf (Idiot-VA) debacle, I will now in the future provide full PDFs or excerpted PDFs so that people can see what exactly I’m referencing in each entry and I will work to update older entries in this document.

Multiple Engine References

(Saturn Engine Program Office Planned/Actual Deliveries: October 1963 – December 1966) (730+ kb GIF)
(Engine Costs for Development and Production over Time – 2001)
(F-1 / J-2 Engine Cost History FY59–70)
(F-1 Engine Production Cost History)
(F-1 Engine Project Plan FY59-75 – 22 Sep 1967)
Liquid Rocket Engine Testing by Dr. Shamin Rahman, Stennis, 2005 (5.68 MB PDF)

Miscellaneous Engine Notes

A 200:1 Expansion Ratio provides the optimum nozzle thrust coefficient for straight LH2 in a vacuum. [1]

References:
[1] Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

Early Rocketry up to the Second World War

Every major and minor industrialized country before World War II had at least one (or two) hobbyists who investigated rocketry, whether it was liquid propellant or solid propellant. In many cases, each hobbyist operated on their own and never truly published their work.

The most notable case of the “independent hobbyist” was Robert Goddard in the United States. Despite coming up with many innovations which would later be repeated by everyone who went into the liquid propellant rocket field, Goddard’s work was ultimately a ‘dead end’, because he was intensely secretive and did not publish his work and kept his patents hidden.

Thus, all modern rocketry stems from three sources essentially:

  1. Military-backed attempts to develop rocket engines for aircraft to boost their speeds and/or reduce their takeoff lengths. Aerojet in the United States was founded for JATO work along with the Jet Propulsion Laboratory (JPL) Their work involved various aircraft (A-20A, USN Flying Boats, P-51D, etc). In the Soviet Union, Sergei Korolev’s first major rocket propulsion project was the BI-1 rocket plane, followed by the usual conventional aircraft boosted by rockets (Pe-2RD, La-7R1, La-7R2 and Yak-3RD). These engines generally used hypergolic propellants and were “low” performance, with thrust levels measured in a few thousands of pounds of thrust. The engineers working on these programs encountered all the problems that that hobbyists like Goddard did, but unlike the hobbyists’ work, their ‘corporate knowledge’ was retained; providing a foundational base for post-war rocket research.

  2. Military-backed attempts to develop compact (and somewhat safe) solid propellant motors for a variety of tasks, such as air launched ground attack rockets (HVAR, Tiny Tim, Holy Moses) or ground launched artillery rockets (Katyusha). The ‘corporate knowledge’ involved in developing these motors was (at least in the United States) later used for the first upper stages of space launch vehicles (Explorer I was placed into orbit via a cluster of solid motors).

  3. Post-war analysis of captured V-2 rockets and interrogations of captured German rocket engineers. At the time, the V-2 was the largest rocket ever manufactured, with an unprecedented 25 tonnes of thrust. The V-2 itself was a very unoptimized system, being a very hasty kludge to get a totally new weapons concept into service in a compressed time frame. This meant that post war, it was not hard for everyone to significantly improve on V-2 performance.

    The V-2’s biggest contribution to aerospace was turning rocketry into a serious science/engineering discipline accepted by everyone. It’s hard to believe, but before WWII and the V-2, there was a serious ‘giggle factor’ attached to the term ‘rocket’ – thus why JPL was called that instead of RPL; and why so many early RATO (Rocket Assisted Takeoff) concepts were called JATO, despite the name(s) being technically inaccurate.

Peenemünde Army Research Center (HVP)

V-2 (A-4 Series B) Twenty-Five Tonne Engine

Propellants: LOX / B-Stoff (75% Alcohol, 25% Water)
O/F Ratio: 1.24
Thrust (sl): 25,000 kgf (55,115 lbf) at 200 ISP
Thrust (vac): 29,601 kgf (65,260 lbf) at 236 ISP (calculated)
Chamber Pressure: 217.55 psi (15 bar)
Dry Weight: 2,484 lb
T/W (sl): 22.5
T/W (vac): 26.27
Propellant Flow: 125 kg/sec
Combustion Chamber Area: 0.69m2 (36.9 inch diameter)
Nozzle Throat Area: 0.13m2 (16.01 inch diameter)
Nozzle Exit Area:
0.42m2 (28.79 inch diameter)
Expansion Area Ratio (ε = Ae/At):
3.23

Notes: The B series engine featured a series of eighteen separate “pot” injectors which were the results of pressure to have a working 25-tonne thrust engine as fast as possible, via clustering smaller stable combustion chambers together. Once the A-4 went into production, pressure from Hitler and other officials for maximum rocket production kept the B series engine ‘frozen’, despite the complex plumbing arrangement needed to feed the eighteen pots not being optimal from a weight and mass production standpoint.

The engine turbopump was driven by the decomposition of T-Stoff (80% Hydrogen Peroxide) and Z-Stoff (1/3rd Sodium Permanganate and 2/3 Water). The V-2 carried 175 kg of T-Stoff and 22 kg of Z-Stoff for this. Originally, von Braun's development team wanted a 725.18 psi (50 bar) combustion chamber pressure; but had to settle for the much lower 217~ psi level due to chamber burnthroughs at the higher pressures and temperatures.

Notes II: Exit diameter accuracy was confirmed via contacting the US Space and Rocket Center in Huntsville, AL and having a staff member measure the inner diameter of their V-2 engine's exit nozzle. According to them, it was about 29 1/8 inches when the engine was brand new [their engine was slightly warped].

Reference:
Modern Engineering for Design of Liquid Propellant Rocket Engines by Harry Arbit (PNG Image)
Physics of Continuous Matter, 2E: Exotic and Everyday Phenomena by B. Lautrup (Page 239)
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

V-2 (A-4 Series C) Twenty-Five Tonne Engine

Notes: This version replaced the complex series of eighteen “pot” injectors on the Series B engine with a single centrally mounted injector plate 350mm in diameter with orifice holes for propellant injection arranged in a complex series of radial, parallel and circular patterns. It never went into production, but helped form inspiration for post-war engines.

(A-4 Series C Engine Drawing)

Reference:
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

Misc Cancelled Programs

STME

STBE

LOCUS (Low Cost Upper Stage) by Aerojet

ARRE by Aerojet (Advanced Peroxide Upper Stage)

RBCC

TBCC by GE

ARC

ARC / Aerojet Agena-2000

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Aerojet GenCorp

Aerojet Transtar Upper Stage Engine

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Aerojet 25AL-1000 (A-20A / Misc JATO)

Propellant: RFNA / Aniline, Pressure Fed
Thrust: 1,000 lbf for 25 seconds

Notes: By 1944, 64 had been produced for the USAAF at a cost of $3450 each.

(Photo of 25AL-1000)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet 38ALDW-1500 (PB2Y-2 JATO)

Propellant: RFNA / Aniline, Pressure Fed
Thrust: 1,500 lbf for 38 seconds

Notes: About 100 were delivered to the USN for the PB2Y-2 Flying Boat.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet 35ALDW-6000 (PB2Y-2 JATO)

Propellant: Concentrated Nitric Acid / Aniline, Pressure Fed
Thrust: 6,000 lbf

Notes: Advanced version developed for USN PB2Y-2. Had three thrust chambers.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet ALD-3000 JATO (B-29 JATO)

Thrust: 3,000 lbf

Notes: Not put into production due to contract cancelation.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet X60ALD-4000 (B-45 JATO)

Thrust: 4,000 lbf

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet LR13-AL-3 (B-29 JATO)

Propellant: Concentrated Nitric Acid / 65% xylidine and 35% Gasoline
Thrust:
4,000 lbf for 40 seconds

Notes: Due to the success of the ALD-4000 tests, the USAF contracted with Aerojet to develop a cold-weather JATO for the B-29 fleet. Xylidine was used as a propellant because the USAF had it stockpiled. Each B-29 would carry four of these disposable units for a total takeoff boost of 16,000 lbf. Approximately 150 units were delivered. The extra thrust from the JATOs enabled the B-29s to carry enough fuel to reach deep into the USSR and then return to a friendly base.

(Photo of LR13-AL-3)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet YLR63-AJ-1 (F-84 JATO)

Propellants: WFNA / Jet Fuel

Notes: Was cancelled because the USAF chose to switch over to a different engine for the F-84 which didn’t need JATO assistance.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet XCALR-2000A-1 “Rotojet” aka Aerotojet

Propellant: RFNA / Anilines

Notes: This was a completely insane concept in which there would be two stationary thrust chambers of 750 lbf each surrounded by two rotating thrust chambers of 200-300 lbf each. The rotating thrust chambers would drive the turbopumps through rotational motion. This would be the main engine for the Northrop XP-79 and the prototype MX-324.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton
Rocketing Into the Future: The History and Technology of Rocket Planes by Michel van Pelt

Aerojet 1.3K Superperformance Rocket Engine (P-51 Rocket Boost)

Propellant: RFNA / 65% aniline and 35% furfuryl alcohol
Thrust: 1,300 lbf for 1 minute
Weight: 1,090 lbs (fully loaded of entire installation)

Notes: Developmental work started in late 1944. One P-51D, tail number 44-73099 was modified to incorporate the Aerojet engine. The installation consisted of two pressurized 75 gallon non-jettisonable tanks under each wing, and a simple ON/OFF switch in the cockpit. It was first flown on 23 April 1945 when NAA Test Pilot Bob Chilton activated it at 21,000 feet. Speed was boosted by approximately 100 MPH for 1 minute, but due to the still-significant drag of the under-wing tanks (despite being much smaller than the standard P-51 drop tanks), the full performance that the rocket engine could impart to the Mustang airframe wasn’t achieved. When the war ended, the program was canceled after completing several flight tests.

(Photo of P-51 with Rocket being test fired on ground)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton
http://aviationtrivia.blogspot.com/2010/12/rocket-boosted-p-51-mustang.html
http://retromechanix.com/rocket-boosted-north-american-p-51d-mustang-1945/ (A fuller more complete version, including reports)

Aerojet 1KS-23,800 Escape SRM

Thrust (sl): 23,800 lbf at 245 ISP
Chamber Pressure: 1,130 psia
Burning Time: 1 second
Propellant Mass: 102 lbs
Total Mass: 126.4 lbs
Case Material: High Strength Nickel Steel (25% Ni) (260,000 PSI Yield, 285,000 PSI Ultimate)
Case Diameter: 8 inches
Case Thickness: 0.030 inches
Nozzle Exit Diameter: 11.95 inches
Expansion Area Ratio (ε = Ae/At): 9

Notes: Designed for General Electric’s Apollo Proposal’s Launch Abort Escape System.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (1.92 MB Excerpt)

Aerojet 2KS-23,800 Escape SRM

Thrust (sl): 23,800 lbf at 245 ISP
Chamber Pressure: 1,130 psia
Burning Time: 2 seconds
Propellant Mass: 210 lbs
Total Mass: 242.9 lbs
Case Material: High Strength Nickel Steel (25% Ni) (260,000 PSI Yield, 285,000 PSI Ultimate)
Case Thickness: 0.041
Nozzle Exit Diameter: 11.95 inches
Expansion Area Ratio (ε = Ae/At): 9

Notes: Designed for General Electric’s Apollo Proposal’s Launch Abort Escape System. Provides thrust along the axis of the spacecraft after the burnout of the 1KS-23,800 motors.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (1.92 MB Excerpt)

Aerojet 5KS-560 Separation SRM

Thrust (vac): 560 lbf at 285 ISP
Chamber Pressure: 500 psia
Burning Time: 2.5 seconds
Propellant Mass: 5.1 lbs
Total Mass: 6.42 lbs
Expansion Area Ratio (ε = Ae/At): 24

Notes: Designed for General Electric’s Apollo Proposal – it fills the function of a separation motor to separate the re-entry vehicle from the spacecraft.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (1.92 MB Excerpt)

Aerojet 9KS-18,100 Escape SRM

Thrust (sl): 18,100 lbf at 245 ISP
Chamber Pressure: 1,100 psia
Burning Time: 1.9 seconds
Propellant Mass: 156 lbs
Total Mass: 182.47 lbs
Expansion Area Ratio (ε = Ae/At): 9

Notes: Designed for General Electric’s Apollo Proposal – it fills the function of an escape motor to separate the glider from the spacecraft for a pad abort.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (1.92 MB Excerpt)

Aerojet LR73

Thrust (sl): 5,000 to 10,000 lbf at 225 ISP
Firing Duration: 300 Seconds

Notes: Terminated in 1955 by the US Government.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (Miscellaneous Aerojet Engines, PDF Excerpt)

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

Thrust (sl): 300,000 lbf at 278 ISP
Thrust (vac): 328,000 lbf
Chamber Pressure: 766 psia

References:
Titan II: A History of a Cold War Missile Program by David K. Stumpf
Aerojet Final Report 4008-F-1: High Chamber Pressure Operating for Launch Vehicle Engines (15 April 1963); Figure III-2
History of Liquid Propellant Rocket Engines by George Paul Sutton

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

Thrust (vac): 80,000 lbf
Chamber Pressure: 819 psia

References:
Titan II: A History of a Cold War Missile Program by David K. Stumpf
Aerojet Final Report 4008-F-1: High Chamber Pressure Operating for Launch Vehicle Engines (15 April 1963); Figure III-2
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet LR87-AJ-5 (Titan II Stage I Engine) / AJ23-132

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 Area Ratio (ε = Ae/At): 8

(Perspective Drawing of LR87-AJ-5)

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
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet LR91-AJ-5 (Titan II Stage II Engine) / AJ23-133

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.80
Thrust (sl): 100,000 lbf at 321 ISP
Chamber Pressure: 827 psia
Height: 110.6 inches
Nozzle Exit Diameter: 64 inches
Expansion Area Ratio (ε = Ae/At): 49.2

Notes: Largely identical to the LR87-AJ-5 with the following distinguishing features:

(Perspective Drawing of LR91-AJ-5)

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
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet LR87-AJ-7 (Titan II GLV Stage I Engine) / AJ23-134(?)

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/M/F Stage I Engine)

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.915
Rated Thrust (vac): 520,000 lbf at 301.1 ISP
Rated Thrust (sl): 437,000~ lbf
Chamber Pressure: 823 psia
Expansion Area Ratio (ε = Ae/At): 15
Operating Cycle: 165 seconds

Notes: Air starts after SRM burnout, hence why it has vacuum thrust values given. Development was begun as part of the Titan IIIM program, using the LR87-AJ-9 as a baseline. Changes involved were redesigning the thrust chamber to operate at a higher chamber pressure and expansion ratio (from 8:1 to 15:1) in order to meet the 520 klbf at altitude thrust requirement. After the cancellation of the Titan IIIM/F program, the engine was then reused in the Titan IIIE program.

References:
Titan IIIE/Centaur D-1T Systems Summary (September 1973)
Program Titan IIIM Standard Space Launch Vehicle Development Report for Stage I Engine Demonstration Testing: Report 9180-941-DR-9 (31 March 1969) (Link to DTIC PDF)
History of Liquid Propellant Rocket Engines by George Paul Sutton

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

Propellants: NTO (N2O4)/Aerozine-50
O/F Ratio: 1.79
Rated Thrust (vac): 101,000 lbf at 318.7 ISP
Chamber Pressure: 853 psia
Expansion Area Ratio (ε = Ae/At): 49.2
Operating Cycle: 225 seconds

References:
Titan IIIE/Centaur D-1T Systems Summary (September 1973)
History of Liquid Propellant Rocket Engines by George Paul Sutton

LR87-AJ-11A (Titan IV Stage I Engine)

O/F Ratio: 1.91
Thrust (sl): 489,000 lbf
Thrust (vac): 550,900 lbf at 303.5 ISP
Chamber Pressure: 854 psia
Expansion Area Ratio (ε = Ae/At): 16

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

LR91-AJ-11A (Titan IV Stage II Engine)

O/F Ratio: 1.775
Thrust (vac): 105,000 lbf at 316.2 ISP
Chamber Pressure: 879 psia
Expansion Area Ratio (ε = Ae/At): 49.2

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet LR113

Propellants: N2O4/UDMH
O/F Ratio:
2.1
Thrust (sl):
47,000 to 175,000 lbf at 245 ISP (maximum thrust)
Chamber Pressure: 236 to 595 psia

Notes: Can start, run, and shut down at any thrust level between 47 and 175 klbf. Part of an effort to develop a continuously variable thrust engine.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (Miscellaneous Aerojet Engines, PDF Excerpt)

Aerojet AJ10-18

Thrust (sl): 160,000 lbf

Notes: Was a contender for the Redstone missile propulsion system. Four chambered engine.

References:
AMC 23M: History of the Redstone Missile System by John W. Bullard

AJ10-37 “Vanguard”

Propellants: WFNA / UDMH
Thrust: 7,500 lbf

References:
The Development of Propulsion Technology for U.S. Space-launch Vehicles: 1926 – 1991 by J. D. Hunley

AJ10-40 “Able”

Propellants: WFNA / UDMH
Thrust: 7,500 lbf

Notes: Modified AJ10-37 Vanguard design.

References:
To Reach the High Frontier: A History of U.S. Launch Vehicles
The Development of Propulsion Technology for U.S. Space-launch Vehicles: 1926 – 1991 by J. D. Hunley

Aerojet AJ10-104 “Able-Star”

Propellants: IFRNA/UDMH
Thrust: 7,575 lbf (without nozzle extension) or 7,890 lbf (with nozzle extension)
Expansion Ratio: 20 (without nozzle extension) or 40 (with nozzle extension)

Notes: Derived from earlier Able engines.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton
http://www.spacelaunchreport.com/thorh2.html

Aerojet AJ10-118 “Able”

IRFNA/UDMH, Pressure Fed.
7,700 lbf

References:
To Reach the High Frontier: A History of U.S. Launch Vehicles

Aerojet AJ10-118A



Delta Engines:

Pressure Fed NTO/Aerozine-50
9,800 lbf (early) 10,200 lbf (late)
125 psia chamber pressure
320 vacuum USP
Expansion Ratio 65

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Aerojet AJ10-118D

Propellants: IRFNA / Aerozine-50

Notes: First used on Delta B in 1962. Replaced vacuum-tube guidance system with transistorized/semiconductor guidance system. Propellant systems lengthened 3 feet.

References:
To Reach the High Frontier: A History of U.S. Launch Vehicles

Aerojet AJ10-118E

Propellants: IRFNA / Aerozine-50

Notes: Propellant tank diameters widened from 33" to 55", doubling propellant capacity. Also, additional helium tanks were added, giving the 118 effectively unlimited in-space restarts.

References:
To Reach the High Frontier: A History of U.S. Launch Vehicles

Aerojet AJ10-118F

Thrust: 9,235 to 9,606 lbf at 290 ISP
Propellants:
NTO/Aerozine-50
Dry Weight: 1,204 lbs

Notes: Very similar to the Transstage (AJ10-138) engine, but not identical. Had more than 1,400 lbf of thrust increase over the AJ10-118E version. Preliminary design completed 1970, didn’t fly until July 1972.

References:
To Reach the High Frontier: A History of U.S. Launch Vehicles
The Development of Propulsion Technology for U.S. Space-launch Vehicles: 1926 – 1991 by J. D. Hunley

Aerojet AJ10-118K

Notes: Introduced 1982.

References:
To Reach the High Frontier: A History of U.S. Launch Vehicles

Aerojet AJ10-133

Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (vac): 24,000 lbf at 430 ISP
Weight: 675~ lbs
T/W Ratio (vac): 35.55
Chamber Pressure: 65 psia
Expansion Area Ratio (ε = Ae/At): 35
Rated Burn Time:
137~ seconds at full thrust (can burn for 546 seconds on a single chamber)

Notes: Pressure-fed four chambered LH2 engine proposed by Aerojet for General Electric’s Apollo proposal. The system would have been insulated with SI-4 insulation, allowing it to operate for fourteen days as a sealed unit with minimal boil-off.

(AJ10-133 Engine Layout)
(AJ10-133 Nozzle Diagram)

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System – Final Report: Volume IV: On-Board Propulsion, Book 2 – Appendix P-A (15 May 1961) by General Electric (1.15 MB PDF excerpt dealing with AJ10-133)

Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System – Final Report: Volume IV: On-Board Propulsion, Book 1 – Text and Appendix P-C (15 May 1961) by General Electric (3.64 MB PDF excerpt dealing with AJ10-133)

Aerojet AJ10-137 (Apollo Service Propulsion System [SPS])

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 Area Ratio (ε = Ae/At): 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 Area Ratio (ε = Ae/At): 40

References:
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)

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 Outside Diameter: 46 inches
Engine Nozzle Inside Diameter: 43.09 inches
Expansion Area Ratio (ε = Ae/At): 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.

(Drawing of AJ10-190 Thrust Chamber/Nozzle)

References:
Space Shuttle Propulsion Systems
by Russell Bardos, Office of Space Flight (26 June 1990) (PDF)

Aerojet AJ23-25

Propellants: HTP/JP-5
Thrust (sl): 3,000 to 10,000 lbf
Rated Starts: Unlimited

Notes: Terminated in 1959 by the US Government.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (Miscellaneous Aerojet Engines, PDF Excerpt)

Aerojet AJ23-153 Transtar

Propellants: N2O4/MMH
Type: Gas Generator
O/F Ratio: 1.8
Thrust (vac):
3,750 lbf at 328 ISP
Engine Weight: 128 pounds
T/W Ratio (vac): 29.29
Chamber Pressure: 350 psia
Expansion Ratio (Ac/At):
136

References:
Orbital Transfer Vehicle: Concept Definition and System Analysis Study NAS8-36108 First Quarterly Review Briefing (30 Oct 1984) (476~ kb PDF excerpt)

Aerojet AJ23-151 Pump Fed OMS

Propellants: N2O4/MMH
Type: Gas Generator
O/F Ratio: 1.93
Thrust (vac):
6,000 lbf at 334 ISP
Engine Weight: 322 pounds
T/W Ratio (vac): 18.63
Chamber Pressure: 350 psia
Expansion Ratio (Ac/At):
154

References:
Orbital Transfer Vehicle: Concept Definition and System Analysis Study NAS8-36108 First Quarterly Review Briefing (30 Oct 1984) (476~ kb PDF excerpt)

Aerojet AJ23-154

Propellants: LOX/LH2
Type: Dual Expander
O/F Ratio: 6.0
Thrust (vac):
3,000 lbf at 483 ISP
Engine Weight: 90 pounds
T/W Ratio (vac): 33.33
Chamber Pressure: 2,000 psia
Expansion Ratio (Ac/At):
1,000

References:
Orbital Transfer Vehicle: Concept Definition and System Analysis Study NAS8-36108 First Quarterly Review Briefing (30 Oct 1984) (476~ kb PDF excerpt)

Aerojet AJ23-156 Transtar III

Propellants: N2O4/MMH
Type: Gas Generator
O/F Ratio: 2.1
Thrust (vac):
3,750 lbf at 343 ISP
Engine Weight: 104 pounds
T/W Ratio (vac): 36.05
Chamber Pressure: 1,430 psia
Expansion Ratio (Ac/At):
400

References:
Orbital Transfer Vehicle: Concept Definition and System Analysis Study NAS8-36108 First Quarterly Review Briefing (30 Oct 1984) (476~ kb PDF excerpt)

Aerojet AJ23-65

Propellants: LOX/LH2

Notes: Thrust chamber was a modified LR87-AJ-1 assembly, and the LOX turbopump was a modified LR87-AJ-1 model. The program objectives were:

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961) (Image of page this is mentioned on)

Aerojet Hylas/Hylas Star

Notes: 6,000 lbf (Hylas) and 13,000 lbf (Hylas Star) LH2 engines designed for upper stage purposes. Hylas Star was formerly called Hydra. Designed to act as a third stage for Titan II, and Titan II/Centaur. Optimum Hylas Star stage weight was determined to be 12,000 lbm for Titan II.

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System, Final Report – Volume IV On-Board Propulsion – Book 2, Appendix P-A. General Electric (15 May 1961)

Aerojet M-1

Ready: 1970s
Propellants: LOX/LH2
O/F Ratio: 5.0
Oxidizer Flow Rate: 2,921 lb/sec
Fuel Flow Rate: 584 lb/sec
Thrust (vac): 1,500,000 lbf at 428 ISP
Weight: 20,000 lbs
T/W Ratio (vac): 75
Chamber Pressure: 1,000 psia
Overall Length: 321 inches
Throat Area: 803 square inches
Nozzle Exit Diameter: 208 inches or 212 inches
Expansion Area Ratio (ε = Ae/At): 40
Rated Lifetime: 500 seconds

Notes: Started on 30 April 1962 under NASA sponsorship by Aerojet as a 1.2 million pound thrust engine to provide second stage power for the NOVA launch vehicle (400 klbf to LEO) with management by Marshall Spaceflight Center. In October 1962, Lewis Research Center obtained management responsibility. At that point, due to NOVA payload increases (now 1 mlbf in LEO), M-1 power had increased to 1.5 million pounds thrust, with design features allowing a future uprating to 1.8 million pounds thrust if necessary.

In order to achieve the 1.8 mlbf future growth goal, special attention was paid to the turbopumps; making sure that they would have the power necessary to achieve the 1,200 PSIA chamber pressure for that growth option. The M-1’s fuel turbopump developed 75,000 shp, compared to the F-1’s 60,000 shp combination fuel/oxidizer turbopump. Used turbine exhaust gas to cool the nozzle extension, similar to the F-1 and J-2.

Due to the death of NOVA, the M-1 was terminated before it’s development was complete in 1965.

(M-1 Side General Outline)
(M-1 Drawing showing engine with thrust chamber only)
(M-1 Drawing showing Initial Design Configuration of 28 June 1962)
(M-1 Drawing showing Configuration B-8 from the 3rd Quarter of 1962)

Reference:
AIAA Paper 67-978: Milestones in Cryogenic Liquid Propellant Rocket Engines
NASA TM X-50854: The M-1 Rocket Engine Project (October 1963) (2.64 MB PDF)
NASA CR-54813: Development of Liquid Oxygen/Liquid Hydrogen Thrust Chamber for the M-1 Engine (15 May 1968)
Report # 2555-M-1-F Development of a 1,500,000-lb-thrust (nominal vacuum) Liquid Hydrogen/Liquid Oxygen Engine (30 August 1967) (27.7~ MB PDF)

Aerojet-General M-1A

Ready: 1990s?
Propellants: LOX/LH2
O/F Ratio: 5.0
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 Area Ratio (ε = Ae/At): 20

Notes: Proposed modernized sea-level startable 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)

Aerojet-General “Advanced LOX/LH2 Engine” for NOVA-class Applications

Propellants: LOX/LH2
O/F Ratio: 6.0
Thrust (sl): 24 million lbf at 383 ISP
Thrust (vac): lbf at 450 ISP
Weight: 312,000 lbs
Chamber Pressure: 2,500 PSIA
Engine Length:
63 feet
Engine Diameter:
80 feet

Notes: Advanced staged combustion engine, with a forced deflection nozzle proposed for 80 foot diameter NOVA rockets.

(Artwork of it, with specifications and citation; 717~ kb GIF)

References:
Advanced Engine Systems by R.C. Stiff, Aerojet-General Corporation (11.7 MB PDF)

Aerojet Dual Fuel Mixed-Mode High Pressure Engine for SSTO Vehicles

Dry Weight: 9,223 lbs

Mode 1 (Hydrocarbon)
Type:
Staged Combustion
Propellants
: LOX/RP-1
O/F Ratio: 2.9
Thrust (sl): 607,000 lbf at 322.9 ISP
Thrust (vac): 657,700 lbf at 349.9 ISP
T/W (sl): 65.81
T/W (vac): 71.31
Chamber Pressure: 4000 psia
Engine Length: 289 inches (with nozzle in stowed position)
Nozzle Diameter: 69.5 inches
Expansion Area Ratio (ε = Ae/At): 40

Mode 2 (Hydrolox)
Type:
Staged Combustion
Propellants
: LOX/LH2
O/F Ratio: 7.0
Thrust (vac): 515,250 lbf at 459.2 ISP
T/W (vac): 55.86
Chamber Pressure: 3000 psia
Engine Length: 346 inches (with nozzle in deployed position)
Nozzle Diameter: 154.5 inches (extension deployed)
Expansion Area Ratio (ε = Ae/At): 200

Notes: This concept uses a common thrust chamber and LOX feed system. Mode 2 utilizes a 200:1 retractable nozzle extension.

(Engine Drawing)
(Engine Cycle [Colorized] [Original])

References:
Advanced High Pressure Engine Study for Mixed-Mode Vehicle Applications (NASA CR-135141) (January 1977) (9.27~ MB PDF)

Aerojet AJ-550 Space Shuttle Main Engine (SSME)

Notes: Appears to have been Aerojet’s proposal for the SSME in the 1970s.

References:
AIAA Paper 71-660: Aerojet AJ-550 Space Shuttle Main Engine (SAE 7th Joint Propulsion Conference)

Aerojet OTV Engine – Advanced Expander Cycle Baseline

Propellants: LOX/LH2
O/F Ratio:
6.0
Thrust (vac):
10,000 lbf at 475.1 ISP
Weight: 437 lbs
T/W Ratio (vac): 22.88
Chamber Pressure: 1,300 psia
Engine Length:
60 inches (Nozzle Retracted) 120 inches (Nozzle Deployed)
Combustion Chamber Diameter:
4.18 inches
Nozzle Length: 95.6 inches
Nozzle Exit Diameter:
61.5 inches
Expansion Area Ratio (ε = Ae/At):
297 (Nozzle Retracted) / 792 (Nozzle Deployed)

Notes: These were the recommended engine specifications based upon 1980 state-of-the-art for a twin engined orbital transfer vehicle (OTV).

(Cycle Diagram)

References:
Aerojet Report 32999-F Orbit Transfer Vehicle (OTV) Engine, Phase A Study Final Report Volume II Study Reports (29 June 1979)

Aerojet AJ-60

LOX/LH2 Expander Cycle
Thrust (Vac):
60,000 lbf at 461 ISP
Chamber Pressure: 1,800 psia
Expansion Area Ratio (ε = Ae/At): 250

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Aerojet AJ-800

Notes: Derived from the NK-33, and uses dual NK-33 turbopump assemblies, one on each side of a new thrust chamber. Roughly twice as powerful as the NK-33.

References:
Booster Main Engine Selection Criteria for the Liquid Fly-Back Booster

Aerojet AJ-1200 Pressure Fed Booster Engine

Propellants: LOX/RP-1
O/F Ratio: 2.4
Thrust (sl): 1,200,000 lbf at 237.8 ISP at NPL
Thrust (vac): 1,466,000 lbf at 289.8 ISP at NPL / 1,020,000 lbf at 288.1 ISP at 70% PC
Weight: 17,484 lbs dry
T/W Ratio (sl): 68.63
T/W Ratio (vac): 83.84
Chamber Pressure: 250 psia
Expansion Area Ratio (ε = Ae/At): 5
Engine Length: 256 inches
Engine Diameter:
160 inches
Nozzle Diameter: 155 inches outside diameter

Notes: Designed for a minimum of 100 uses with a service life of 15,000 seconds. Designed to operate at two thrust levels – Normal Power Level (NPL) and a reduced thrust level of approximately 70% PC of NPL during “Max Q” transition points.

(AJ-1200 Design Features Drawing)
(AJ-1200 Engine Assembly Drawing)
(AJ-1200 Engine Envelope Drawing)
(AJ-1200 Interface Drawing)
(AJ-1200 Thrust Chamber Drawing)

References:
MSFC Report SE-019-012-2H (15 March 1972) (2.21 MB PDF)

Aerojet XLR-134 Cryogenic Upper Stage Engine

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

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 Area Ratio (ε = Ae/At): 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)
Aerojet Redmond Operations (October 2011) (23.9 MB PDF) (gives alternate specs on PDF page 52)

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 Area Ratio (ε = Ae/At):
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-11

Notes: Japanese HTV’s 1, 2 and 4 used this; and ESA’s ATV’s use this.

Aerojet R-4D-15 (HiPAT)

Propellants: NTO (N2O4) / MMH
O/F Ratio: 1
Thrust (vac): 100 lbf at 323 ISP
Weight: 12 lbs
T/W Ratio (vac): 8.33
Chamber Pressure: 137 psia
Engine Length: 28.57 Inches
Nozzle Exit Diameter: 14.25 inches
Expansion Area Ratio (ε = Ae/At):
40
Rated Restarts: 500
Maximum Single Firing Length: 500 seconds

Notes: There is a option for this engine with a 300:1 expansion ratio nozzle; which has the same thrust, an ISP of 320 seconds, weighs weighs 11.5 lbs and is 24.72” long with a nozzle exit diameter of 12.8”.

References:
Advanced Chemical Propulsion for Science Missions (NASA TM-2008-215069)
Aerojet Redmond Operations (October 2011) (23.9 MB PDF) (gives more detailed information)

Aerojet R-6D

Propellants: NTO (N2O4) / MMH
Thrust (vac): 22 Newtons (4.94 lbf) at 294 ISP
Weight:
1 pound
T/W Ratio (vac): 4.94
Chamber Pressure: 97 psia
Engine Length: 7.34 inches
Nozzle Exit Diameter: 2.17 inches
Expansion Area Ratio (ε = Ae/At):
100
Rated Restarts: 336,331
Maximum Single Firing Length: Unlimited

References:
Aerojet Redmond Operations (October 2011) (23.9 MB 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 Area Ratio (ε = Ae/At): 22 to 30 (depending on whether it’s a long, short or no scarf configuration).
Rated Restarts: 50,000
Maximum Single Firing Length: 23,000 seconds

Notes: Original Space Shuttle Program (SSP) specifications were 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)
Aerojet Redmond Operations (October 2011) (23.9 MB PDF) (Gives differing information)

Aerojet R-40B

Propellants: NTO (N2O4) / MMH
Thrust (vac): 900 lbf at 293 ISP
Weight: 15 lbs
T/W Ratio (vac): 60
Chamber Pressure: 150 psia
Expansion Area Ratio (ε = Ae/At): 60
Rated Restarts: 50,000
Maximum Single Firing Length: 23,000 seconds

Notes: Derived from the Space Shuttle’s R-40A RCS engine, but in a much more conventional configuration.

References:
Aerojet Redmond Operations (October 2011) (23.9 MB PDF)

Aerojet R-42

Propellants: NTO (N2O4) / MMH
Thrust (vac): 200 lbf at 303 ISP
Weight: 10 lbs
T/W Ratio (vac): 20
Chamber Pressure: 103 psia
Engine Length: 31 Inches
Nozzle Exit Diameter: 15.34 inches
Expansion Area Ratio (ε = Ae/At):
160
Rated Restarts: 134
Maximum Single Firing Length: 3,940 seconds

References:
Aerojet Redmond Operations (October 2011) (23.9 MB PDF)

Aerojet BPT-2000 Hall Effect Thruster

Propellants: Xenon (stored at 3,100 kg/m3)
Electrical Input: 2700 W at 400 V
Thrust (vac): 123 mN (0.027 lbf) at 1,765 ISP
Weight: 11.464 lbs
T/W Ratio (vac): 0.0023 (thruster alone, excludes power supply)
Fuel Consumption: 7.1 milligrams/second
Lifetime (Continuous): 6,000+ Hours
Rated Restarts: 6,000

References:
Aerojet Redmond Operations (October 2011) (23.9 MB PDF)

Aerojet BPT-4000 Hall Effect Thruster

Ready (Flight Certified): August 2010 (USA-214 AEHF Military Satellite)
Propellants: Xenon (stored at 3,100 kg/m3)
Electrical Input: 4500 W at 400V
Thrust (vac): 254 mN (0.057 lbf) at 2,020 ISP
Weight: 27.11 lbs
T/W Ratio (vac): 0.0021 (thruster alone, excludes power supply)
Lifetime (Continuous): 10,000+ Hours
Rated Restarts: 6,700

References:
Aerojet Redmond Operations (October 2011) (23.9 MB PDF)

Aerojet PPT / PRS-1

Maximum Input Power: 70W
Maximum Thrust:
860 μN at 650 to 1,350 ISP
Mass:
4.9 kg (of which 0.07 kg is PTFE Teflon)

Notes: This PPT relies on the Lorentz force generated by an arc passing from anode to cathode and the self-induced magnetic fields to accelerate a small quantity of chlorofluorocarbon propellant. Teflon has been used as the propellant to date.

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs
(2006) – Appendix D

Aerojet/GRC – NASA Evolutionary Xenon Thruster (NEXT)

Propellants: Xenon (stored at 3,100 kg/m3)
Electrical Input: 6900 W
Thrust (vac): 235 mN (0.052 lbf) at 4,100 ISP

Weight: 29.32 lbs
T/W Ratio (vac): 0.00177 (thruster alone, excludes power supply)
Rated Restarts: 3,650

Active Optics Area: 36 cm diameter

References:
Aerojet Redmond Operations (October 2011) (23.9 MB PDF)

Aerojet/GRC – NASA High Power Electric Propulsion (HiPEP)

Ready: 2010s
Propellants: Xenon (stored at 3,100 kg/m3)
Power Needed: 39.3 kilowatts
Thrust (vac): 0.15 lbf (670 mN) at 9,620 ISP
Fuel Consumption: 7 milligrams/second at full power.

Notes: Was intended for use on the Jupiter Icy Moons Orbiter (JIMO) probe, which was cancelled in 2005.

References:
NASA TM 2004-213194
The High Power Electric Propulsion (HiPEP) Ion Thruster (588~ kb PDF)
Aerojet Redmond Operations (October 2011) (23.9 MB PDF) (is not detailed, but mentioned in a list, showing it’s an Aerojet engine)

Ad Astra

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

Bell Aircraft Company

Fluorine Pump/Pressure-Fed Propulsion System Proposal for GE’s Apollo

Propellants: LF2/LH2
Thrust (vac): 12,083 lbf at 446.2 ISP (pump-fed mode) with an O/F of 11.92
Thrust (vac): 3,983 lbf at 445.8 ISP (pressure-fed mode) with an O/F of 10
Weight: 555~ lbs
T/W Ratio (vac): 21.77 (pump-fed mode); 7.17 (pressure-fed mode)
Chamber Pressure: 300 psia in Pump-Fed, 100 psia in Pressure-Fed
Expansion Area Ratio (ε = Ae/At): 45
Rated Restarts: 15 (pump fed)
Rated Burn Time:

Notes: Combination pump/pressure fed system. The pressure fed system is used for mid-course corrections.

(Drawing of Bell’s LF2 Engine)

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System – Final Report: Volume IV: On-Board Propulsion, Book 3 – Appendix P-B (15 May 1961) by General Electric (17.7 MB PDF containing complete book here, as it deals in depth with Bell’s LF2 Proposal)

Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System – Final Report: Volume IV: On-Board Propulsion, Book 1 – Text and Appendix P-C (15 May 1961) by General Electric (2.23 MB PDF excerpt dealing with Bell’s LF2 Proposal)

Lockheed Propulsion Company

Space Shuttle 156”, 7-Segment SRM Proposal

Initial Thrust (sl): 2,679,163 lbf at 262.6 ISP
Propellant: PBAN (LPC-580 Modified)
Chamber Pressure (Average) 631 psia
Booster Weight (Loaded): 1,384,793 lbs
Propellant Weight: 1,231,030 lb
Total Inert Weight: 150,000~ lbs
      Case Weight: 95,373 lbs in seven D6AC steel-cased segments using Abestos/NBR insulation
      Nozzle Weight: 17,004 lbs
Length (Motor): 110 feet
Length (Booster): 125 feet
Expansion Area Ratio (ε = Ae/At): 8.33

Notes: Up to 26% more SRM thrust can be added by adding up to two additional segments and enlarging the throat of the nozzle. The 8 segment motor would have 1.41 million lbs of propellant and the 9 segment motor would have 1.58 million lbs of propellant.

References:
LPC Document No. 629-6 Vol II, Book I – Study of Solid Rocket Motors for A Space Shuttle Booster (15 March 1972)

Pratt & Whitney

Pratt & Whitney RL10A-1 (LR-115)

Ready (Flight Certified): November 1961
Propellants: LOX/LH2
Thrust (Vac): 15,000 lbf at 422 ISP
Chamber Pressure: 300 psia
Expansion Area Ratio (ε = Ae/At): 40

Notes: RL10 is the P&W civilian designation, while LR-115 is the military designation. Design work began 1958 with static firing tests in 1959, preliminary flight rating tests late 1961 and first Centaur flight in November 1963.

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
History of Liquid Propellant Rocket Engines by George Paul Sutton

Pratt & Whitney RL10A-3

Ready (Flight Certified): June 1962
Propellants: LOX/LH2
Thrust (Vac): 15,000 lbf at 427 ISP
Weight: 330.6 lbs (Saturn), 317.6 lbs (Centaur)
T/W Ratio (Vac): 45.37 (Saturn), 47.22 (Centaur)
Chamber Pressure: 300 psia
Expansion Area Ratio (ε = Ae/At): 40

Notes: Used on early Centaurs and the Saturn S-IV stage.

(RL10A-3 Drawing)

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
PWA FR-914:
RL10 Engine Reliability Studies for Apollo Mission (12 March 1964)
RL10 Liquid Rocket Engine: Installation Handbook (March 1966) (38~ MB PDF)

Pratt & Whitney RL10A-3A

Notes: RL10A-3 engine with some modifications to make it more suitable for the new requirements of the Apollo Mission.

References:
PWA FR-914: RL10 Engine Reliability Studies for Apollo Mission (12 March 1964)

Pratt & Whitney RL10A-3CM-1

Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (sl):
7,000 lbf
Thrust (Vac):
15,000 lbf at 433 ISP
Chamber Pressure: 300 psia
Expansion Area Ratio (ε = Ae/At): 40

Notes: Used on Atlas-Centaur 10, the first operational Centaur launch with a true payload (Surveyor I).

References:
NASA TN D-4580 Atlas-Centaur Flight Performance For Surveyor Mission A (May 1968)
GD/C-BTD65-017 External Design Loads – Operational Centaur Vehicles (AC-6 through AC-15) (1 May 1965)

Pratt & Whitney RL10A-3S

Notes: RL10A-3 engine with modifications for interfacing with Saturn S-IV stage. Achieved an average ISP of 431.3 seconds in an actual launch, compared to the value of 425.3 that MSFC used before launch in computer simulations.

References:
Saturn I Launch Vehicle SA-8 and Launch Complex 37B Functional Systems Description Volume IX
(May 1964)
MPR-SAT-FE-64-19
Results of the Seventh Saturn I Launch Vehicle Test Flight: SA 7 (31 January 1966)

Pratt & Whitney RL10A-3-1

Ready (Flight Certified): September 1964
Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (sl): 7,000 lbf
Thrust (Vac): 15,000 lbf at 433 ISP
Weight: 306.3 lbs
T/W Ratio: 48.97
Chamber Pressure: 300 psia
Expansion Area Ratio (ε = Ae/At): 40

(RL10A-3-1 Drawing)
(RL10A-3-1 Drawing – Alternate)

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)
RL10 Liquid Rocket Engine: Installation Handbook (March 1966) (38~ MB PDF)

References:
Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967) (14.1 MB PDF)
GD/C-BTD65-017 External Design Loads – Operational Centaur Vehicles (AC-6 through AC-15) (1 May 1965)

Pratt & Whitney RL10A-3-2A

Notes: This is an RL10A-3A engine with the capability of operation in the low idle (pressurized) mode at very low thrust levels.

References:
PWA FR-914:
RL10 Engine Reliability Studies for Apollo Mission (12 March 1964)

Pratt & Whitney RL10A-3-3

Ready (Flight Certified): October 1966
Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (sl): 7,000 lbf
Thrust (Vac): 15,000 lbf at 444 ISP
Weight: 301.5 lbs
T/W Ratio: 49.75
Chamber Pressure: 400 psia
Expansion Area Ratio (ε = Ae/At): 57
Restart Capability: 3
Continuous Operating Duration: 450 seconds
Engine Total Life: 4,000 seconds

(RL10A-3-3 Drawing)
(RL10A-3-3 Drawing – Alternate)

References:
GDCA BNZ72-020 Centaur D-1A Systems Summary (September 1972) (26.9 MB PDF) (Page 3-28)
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
GD/C-BTD65-017
External Design Loads – Operational Centaur Vehicles (AC-6 through AC-15) (1 May 1965)
RL10 Liquid Rocket Engine: Installation Handbook (March 1966) (38~ MB PDF)
(RL-10A-3-3 Specifications from ‘Design Report for RL10-A-3-3’ – 600~ kb PDF Extract)

RL10A-3-7

Notes: Mentioned in the reference below.

References:
FR-6011 Design Study of RL10 Derivatives: Final Report, Volume II Engine Design Characteristics (15 December 1973) (36.2 MB PDF)

Pratt & Whitney Space Tug RL10 Derivative – Category IIA

Propellants: LOX/LH2
O/F Ratio: 6.0
Thrust (Vac): 15,000 lbf at 459.2 ISP (Full Thrust)
Thrust (Vac): 3,750 lbf at 437.5 ISP (Maneuvering Thrust)
Weight: 513 lbs
T/W Ratio: 29.23
Chamber Pressure: 400 psia (full thrust); 102 psia (Maneuvering Thrust)
Expansion Area Ratio (ε = Ae/At): 66.3 (with nozzle retracted) 262 (with nozzle extended)

Notes: Derived from RL10A-3-3 engine with the following major changes:

(RL10 Derivative Category IIA Drawings and Weight Breakdown)

References:
FR-6011 Design Study of RL10 Derivatives: Final Report, Volume II Engine Design Characteristics (15 December 1973) (36.2 MB PDF)

Pratt & Whitney RL200 (1960)

Propellants: LOX/LH2
Thrust (Vac): 200,000 lbf
Total Installation Length: 177 inches
Engine Machinery Length: 65 inches
Nozzle Length: 94 inches
Nozzle Diameter: 90 inches

Notes: This engine was Pratt & Whitney’s proposal for the 200K LH2 Second Stage Engine contract for Apollo that was eventually won by Rocketdyne’s J-2. Essentially, the RL200 was a “big” RL10; utilizing the shunt expander cycle, but modified with the addition of a separate gear driven inducer pump to reduce the launch vehicle’s job regarding propellant supply.

P&W submitted their proposal in mid-March 1960 for the RL200 to NASA, with a total development cost estimated to be $138 million. Development cost estimates for the other two competitors were $66 million for Aerojet and $44 million for Rocketdyne.

Apparently in the late 1990s, the RL200 designation was reused by P&W for a speculative engine design.

(RL200 Cycle Schematic)
(RL200 Installation Schematic)

References:
Advanced Engine Development at Pratt and Whitney: The Inside Story of Eight Special Projects, 1946-1971 by Dick Mulready

Pratt & Whitney Space Tug RL10 Derivative – Category IIB

Propellants: LOX/LH2
O/F Ratio: 6.0
Thrust (Vac): 15,000 lbf at 459.2 ISP (Full Thrust)
Thrust (Vac): 3,750 lbf at 437.5 ISP (Pumped Idle Performance)
Weight: 474 lbs
T/W Ratio: 31.64
Chamber Pressure: 400 psia (full thrust); 102 psia (Maneuvering Thrust)
Expansion Area Ratio (ε = Ae/At): 66.3 (with nozzle retracted) 262 (with nozzle extended)

Notes: Derived from RL10A-3-3 engine and similar to the Derivative IIA engine, except it doesn’t have the requirement for two-phase pumping capability at full thrust. The following major changes were made compared to the RL10A-3-3:

(RL10 Derivative Category IIB Drawings and Weight Breakdown)

References:
FR-6011 Design Study of RL10 Derivatives: Final Report, Volume II Engine Design Characteristics (15 December 1973) (36.2 MB PDF)

Pratt & Whitney Space Tug RL10 Derivative – Category IV

Propellants: LOX/LH2
O/F Ratio: 6.0
Thrust (Vac): 15,000 lbf at 470 ISP
Thrust (Vac): 3,750 lbf at 447 ISP (Maneuvering Thrust)
Weight: 424 lbs
T/W Ratio: 35.37
Chamber Pressure: 915 psia (Full thrust), 234 psia (Maneuvering Thrust)
Expansion Area Ratio (ε = Ae/At): 125 (with nozzle retracted) 401 (with nozzle extended)

Notes: Developed using 1973 Technology Base. Uses the expander cycle and many RL10 technologies, but is a “clean sheet” design intended to provide maximum Space Tug capability. Has the same operational features as the Derivative IIA engine (listed below), and is constrained by the same test facility/installed length constraints as the Derivative IIA.

(RL10 Derivative Category IV Drawings and Weight Breakdown)

References:
FR-6011 Design Study of RL10 Derivatives: Final Report, Volume II Engine Design Characteristics (15 December 1973) (36.2 MB PDF)

Pratt & Whitney Space Storable FLOX/Methane Pump Fed Engine

Propellants: Liquid Flox (82.6% Fluorine)/Liquid Methane
O/F Ratio: 5.076
Thrust (Vac): 5,000 lbf at 399.4 ISP
Mass: 98 lb
T/W Ratio (vac): 51
Chamber Pressure: 500 psia
Total Length: 46.316 inches
Nozzle Diameter: 21.95 inches
Expansion Area Ratio (ε = Ae/At): 70

(Engine Side View)

References:
NASA CR-72552 Space Storable Engine Characterization – Final Report (24 November 1969) (56.1 MB PDF)

Pratt & Whitney XLR-129-P-1 / P&W Space Shuttle Main Engine Proposal

Propellants: LOX/LH2
O/F Ratio: 6.0 (for data shown – could operate at 5.0 or 7.0 ratios)
Thrust (vac): 244,000 lbf at 450 ISP with an area ratio of 75:1
Thrust (sl):   209,000 lbf at 385.68 ISP with an area ratio of 35:1 (check-estimated data via ISP = thrust / mdot)
Weight: 3,520 lbs with 75:1 nozzle.
Chamber Pressure: 2,740 psia
Envelope Diameter: 69.25 inches
Envelope Length (nozzle retracted): 80 inches
Envelope Length (Nozzle extended): 131.7 inches
Exit Diameter: 46 inches (Primary Nozzle)
                          66 Inches (Secondary Nozzle)
Expansion Area Ratio (ε = Ae/At): 35:1 (Primary Nozzle)
                                                          75:1 (Secondary Nozzle Fully Deployed)

Notes: Originally developed as a 250K thrust engine for the USAF ISINGLASS rocketplane, but later uprated to 350K for use as P&W’s entry into the SSME competition. Apparently, the original design concept was to design an engine family that could be tailored to any desired aerospace vehicle application via mix and matching interchangeable one position or two position nozzle extensions to a standard turbopump/combustion chamber.

First known engine using extensible nozzles, which had been invented in 1948 by a scientist at the United Aircraft Research Laboratory in 1948. The first hot fire of an extensible nozzle was the ground tests of the XLR-129 in early 1967.

Data shown above is for the USAF XLR129-P-1. The LR129-P-1 flight engine would have increased chamber pressure to 3000 psi, resulting in various performance increases (see LR129-P-1 Performance Extract below).

Drawings, Data Extracts:
XLR129-P-1 Cutaway (GIF)
XLR129-P-1 Cutaway, Alternate (GIF)
XLR129-P-1 Side Drawing (GIF)
XLR129-P-1 Nozzle Shape (GIF)
XLR129-P-1 Nozzle Details (GIF)
XLR129-P-1 Performance Extract (280~ kb PDF)
   LR129-P-1 Performance Extract (140~ kb GIF)

References:s
AIAA Paper 71-658: Pratt and Whitney Aircraft’s Space Shuttle Main Engine (SAE 7th Joint Propulsion Conference) ← FIND for details.
Advanced Engine Development at Pratt and Whitney: The Inside Story of Eight Special Projects, 1946-1971 by Dick Mulready
Air Force Reusable Rocket Engine Program: XLR129-P-1 Engine Performance (April 1969) (2.3 MB PDF)
Air Force Reusable Rocket Engine Program: XLR129-P-1 Demonstrator Engine Design, Volume I (April 1970) (4~ MB PDF)
Air Force Reusable Rocket Engine Program: XLR129-P-1 Demonstrator Engine Design, Volume III (April 1970) (12.3~ MB PDF)
Air Force Reusable Rocket Engine Program: XLR129-P-1 Final Report, Volume I (January 1971) (18.5 MB PDF)
Air Force Reusable Rocket Engine Program: XLR129-P-1 Final Report, Volume II (January 1971) (24.5 MB PDF)
Air Force Reusable Rocket Engine Program: XLR129-P-1 Final Report, Volume III (January 1971) (23.3 MB PDF)
History of Liquid Propellant Rocket Engines by George Paul Sutton

Pratt & Whitney RL100 Advanced Expander Cycle Engine (AEE) Family

Propellants: LOX/LH2
O/F Ratio:
6.0
Thrust (vac):
15,000 lbf at 482 ISP (mainstage), 1,500 lbf at 455 ISP (pumped idle)
Weight: 427 lbs
T/W Ratio (vac): 35.12
Chamber Pressure: 1,500 psia (mainstage), 154 psia (pumped idle)
Engine Length:
60 inches (stowed) 120 inches (deployed)
Nozzle Diameter:
36.72 inches (stowed) 63.65 inches (deployed)
Expansion Area Ratio (ε = Ae/At):
210 (Nozzle Stowed) 640 (Nozzle Deployed)

Notes: Developed using 1980 Technology Base as a “clean sheet” engine for a man-rated Orbital Transfer Vehicle. A later study looked at a downrated version with the following specifications:

Thrust (vac): 7,500 lbf at 474 ISP
Weight: 300 lbs
T/W Ratio (vac): 25
Chamber Pressure: 1,200 psia
Expansion Area Ratio (ε = Ae/At):
600 (Nozzle Deployed)

(AEE Engine Drawing)

References:
Orbit Transfer Vehicle (OTV) Advanced Expander Cycle Engine Point Design Summary (15 March 1981)

Orbital Transfer Vehicle: Concept Definition and System Analysis Study NAS8-36108 First Quarterly Review Briefing (30 Oct 1984) (476~ kb PDF excerpt)

Pratt & Whitney RL10A-3-3A

Ready (Flight Certified): November 1981
Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (Vac): 16,500 lbf at 444.4 ISP
Chamber Pressure: 475 psia
Expansion Area Ratio (ε = Ae/At): 61

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
The Centaur Upper Stage Vehicle
by Thomas J. Rudman and Kurt L. Austad (848 kb PDF)

Pratt & Whitney RL10A-3-3B

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 ESCORT Bi-Modal Nuclear Thermal Rocket

Propellants: LH2
Thrust (Vac): 15,000 lbf at 900 ISP
Power Generation Capability: 10-20 kWe

Notes: Grew out of the XNR2000 NTR family.

References:
AIAA 2004-3863: TRITON: A TRImodal capable, Thrust Optimized, Nuclear Propulsion and Power System for Advanced Space Missions (249~ kb PDF)

Pratt & Whitney TRITON Bi-Modal Nuclear Thermal Rocket

Propellants: LH2
Thrust (Vac): 15,000 lbf at 900+ ISP
Power Generation Capability: 25 kWe

Notes: Grew out of the ESCORT Bi-Modal NTR concept. Designed to use LANTR (LOX-Augmented Nuclear Thermal Rocket) to achieve 40 to 50 klbf of thrust without having to scale up reactor size which increases costs.

References:
AIAA 2004-3863: TRITON: A TRImodal capable, Thrust Optimized, Nuclear Propulsion and Power System for Advanced Space Missions (249~ kb PDF)

Pratt & Whitney RL10A-4

Ready (Flight Certified): December 1990
Propellants: LOX/LH2
O/F Ratio: 5.5
Thrust (Vac): 22,300 lbf at 451 ISP
Weight: 370 lbs
T/W Ratio (vac): 60.27
Chamber Pressure: 578 psia
Expansion Area Ratio (ε = Ae/At): 84

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
RL10A-4 Product Brochure (1.07 MB PDF)

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 Area Ratio (ε = Ae/At): 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
O/F Ratio: 5.5
Thrust (Vac): 22,300 lbf at 451 ISP
Weight: 370 lbs
T/W Ratio (vac): 60.27
Chamber Pressure: 610 psia
Length: 91.5 inches
Expansion Area Ratio (ε = Ae/At):
84

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
RL10A-4-1/-A-4-2 Brochure (4.39 MB PDF)

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 inches (Stowed), 163.5 inches (Deployed)
Nozzle Diameter: 84.5 inches
Expansion Area Ratio (ε = Ae/At): 285
Rated Starts: 15
Rated Lifetime: 3,500 seconds

(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)

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
RL10B-2 Product Brochure (362~ kb PDF)
Program Status of the Pratt & Whitney RL60 Engine (448~ kb PDF)

Pratt & Whitney RL20P-3

Propellants: LOX/LH2 staged combustion
Thrust (sl): 250,000 lbf
Thrust (vac): 265,000 lbf
Chamber Pressure: 3000 psia
Exit Diameter: 36 inches
Expansion Area Ratio (ε = Ae/At): 20.5

Description: Designed around 1963.

There’s an interesting story about this engine demonstrating the “old boy” network back then in US aerospace. The model of the engine was shipped to Los Angeles for the United Aircraft Corporation Exposition 1963; but it was dropped at the airport and broken into several pieces with less than a week to go to showtime.

It happened to be that the airport in question was right next to the Saberliner division of NAA, and Saberliners were powered by P&W engines. Thus, somebody at Pratt knew somebody at North American; and the broken model of the RL20P-3 was taken into NAA’s model shop and restored in about three days. The design of the RL20P-3 impressed Rocketdyne enough that one of the chief designers was then offered a job at Rocketdyne on the spot during the UAC Expo ‘63.

P&W/UAC Brochure Description circa 1963:

Over the past four years, Pratt & Whitney Aircraft has conducted a comprehensive study to define the propulsion requirements for the next generation of launch vehicles. As a part of these studies, which have covered a wide range of sizes and missions, the engineering mockup shown here was constructed to assist in component arrangement and integration, and shows what a 250,000-pound thrust space transport engine might look like.

To provide an economic space transportation system, it appears that the next generation of launch vehicles will be developed as reusable systems. Engines for these vehicles will be significantly different from the throw-away types in current use, and must possess many of the characteristics of the engines in use in today's jet aircraft. Stability over a wide range of operating conditions, variable thrust to permit vehicle control and ground checkout, time between overhauls measured in hours, and operational dependability must be coupled with high performance to meet the propulsion needs of a space transport system.

The engine design represented by the mockup, which has been designated RL20P-3, employs oxygen/hydrogen propellants. It produces 250,000 pounds thrust at sea level, has a nozzle expansion ratio of 20.5, and an exit diameter of three feet. It operates at a main chamber pressure of 3000 psia at full thrust; thrust is variable from 10% to 100%. By using traditional transport engine design practice, the initial time between overhauls (T.B.O.) is expected to exceed 10 hours. This T.B.O. will allow 75 to 100 flights before engine removal for overhaul.

Around 1965-66, the RL20 design was modified to include a two position nozzle, as shown in one of the photographs linked below.

(RL20P-3 Mockup, Circa 1963 – 600~ K PNG)
(J-2 and RL20P-3 compared, circa 1963– 300~ K PNG)
(RL20 Mockup with Two Position Nozzle – 270~ K PNG)

Saturn Improvement Studies Description:
“A new approach to the design of a high pressure bell engine was also demonstrated in full scale model form by Pratt & Whitney.

Somewhat less exotic than Rocketdyne's approach, the P&W entry, nevertheless, exhibits some impressive innovations. Designated RL-20, it is a restartable LOX/Hydrogen engine similar in size and general appearance to the J-2.

Aside from the high pressure combustion chamber, its main feature is a two position nozzle skirt extension.

With the skirt in a retracted position, the engine takes up less axial length in the stacked vehicle. The engine can be operated in this configuration using the regeneratively cooled stub nozzle to provide expansion for low altitude operation. At high altitude or after staging, the nozzle skirt would be extended (in 15 seconds) to increase its expansion ratio.

Bleed hydrogen is used to cool the extended skirt and is expanded to local ambient pressure in miniature nozzles at the large nozzle's outlet plane. Nearly as much thrust results from this effect as would be obtained in burning the hydrogen. Pump turbines are mounted between a preburner and the film-cooled (hydrogen) main chamber where final O2 is added. The preburner, main combustion chamber arrangement provides a progressive combustion effect which may reduce the possibility of combustion instability.”

References:
Saturn Improvement Studies – A Summary – Case 330 (28 October 1966)
Advanced Engine Development at Pratt and Whitney: The Inside Story of Eight Special Projects, 1946-1971 by Dick Mulready

Pratt & Whitney RL60

Propellants: LOX/LH2
O/F Ratio: 5.0 to 6.0
Thrust (Vac): 60,000 lbf at 465 ISP
Weight: 1,100 lbs
T/W Ratio (vac): 54.54
Chamber Pressure: 1,200 psia
Engine Length: 7.3 feet
Engine Diameter: 7.5 feet
Rated Starts:
45
Rated Lifetime: 4,050 seconds

Notes: Evolved outgrowth of the earlier failed Commercial Cryogenic Advanced Upper Stage Engine (cCAUSE), RL50, and SPW2000 programs.

(RL60 Render)

References:
Program Status of the Pratt & Whitney RL60 Engine (448~ kb PDF)

Pratt & Whitney Advanced Expander Test Bed (AETB)

Propellants: LOX/LH2 Split Expander Cycle
O/F Ratio: 6 (can operate on mixtures between 5.0 and 7.0)
Thrust (Vac): 20,000 lbf @ 480 ISP
Thrust (sl): 15,059 lbf @ 361.42 ISP
Weight: 2,200 lbs (est)
T/W Ratio (vac): 9.09
T/W Ratio (sl):
6.845
Chamber Pressure: 1,200 PSIA
Expansion Area Ratio (ε = Ae/At): 7.5 (Conical nozzle, instead of bell)
Rated Starts:
100
Rated Lifetime: 2 Hours (5 Hours Desirable)

Notes: Preliminary design began on 27 April 1990 and was completed January 1991. The PDR was held 29-31 January 1991 at NASA Lewis. The engine was designed around a 1,500 PSIA, 25K vacuum thrust point to add in component design margin and flexibility for operating in off-design conditions.

The goals of the AETB program were to:

(AETB Drawing)

References:
Advanced Expander Test Bed Program: Preliminary Design Review Report, May 1991 (9.66~ MB PDF)

Pratt & Whitney / Aerojet COBRA
(Co-Optimized Booster for Reusable Applications)

Propellants: LOX/LH2
O/F Ratio: 6
Thrust (sl):
492,590 lbf at 373.3 ISP
Thrust (vac):
600,000 lbf at 454.7 ISP
Chamber Pressure: 3,000 psia
Engine Length: 180 inches
T/W (Vac): 75
Engine Mass: 8,000 lbs

Notes: The goal of COBRA was to produce a rocket engine prototype that would be simple to operate, provide high reliability and a long life, and reduce cost per launch by virtue of being reusable. Staged combustion with a 100 mission lifetime and a maintenance checkup every 50 missions.

(COBRA Render)
(COBRA Engine Cycle)

References:
FS-2002-09-141-MSFC
Main Engine Candidates for a Second Generation Reusable Launch Vehicle (Sep. 2002) (535~ kb PDF)
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Pratt & Whitney AR1000

Notes: Developed using the Russian RD-180 as a point of departure.

(AR-1000 Side CAD Render)
(AR-1000 Perspective CAD Render)

References:
Development of Reusable Engines by Frederick (Rick) Bachtel (862~ kb PDF)

Reaction Motors

RMI 6000C-4 “Black Betsy” (Bell X-1)

Propellants: LOX / 75% Alcohol, 25% Water
Thrust (sl):
6000 lbf at 209 ISP
Chamber Pressure: 220 psia
Rated Lifetime:
seconds

Notes: Gained its nickname because the engine was painted black to prevent rust. Had four individual thrust chambers, each contributing 1,500 lbf to engine totals; and thrust was controlled by turning on/off individual thrust chambers. Turbopump was powered by HTP.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

RMI XLR-99 (North American X-15)

Propellants: LOX / Ammonia
O/F Ratio: 1.25
Thrust (18,700 ft):
50,000 lbf at 230 ISP
Thrust (Vac): 50,870 lbf
Chamber Pressure: 600 psia
Dry Mass: 910 lbs
Overall Length: 82 inches
Overall Diameter: 39.3 inches
Expansion Area Ratio (ε = Ae/At): 9.8 (design altitude of 18,500 ft)

Notes: Was throttleable down to 30% via lowering chamber pressure to 225 psia. Had significant and severe problems during development to the point where NASA ordered a competing engine. Never reached design goal of 240~ ISP.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

RMI 20K Engine (Viking Rocket)

Propellants: LOX/Alcohol
Thrust: 20,000 lbf

Notes: First liquid rocket engine to have gimballed thrust chamber.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

RMI 400 lb / 220 lb Engine (Lark SAM)

Propellants: Nitric Acid / Aniline
Thrust (Boost Chamber): 400 lbf
Thrust (Sustainer Chamber): 220 lbf

Notes: Boost chamber turned off when the missile reached Mach 0.85, while the sustainer operated continually. Despite developing the engine, Rocket Motors International lost out on the product improvement and follow on production work to Rocketdyne.

(Lark Engine)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

RMI 5K Engine (Sparrow III AAM)

Propellants: RFNA / 20% NTO / Hydrazine/ammonium thiocyanite additive
Thrust: 5,000 lbf for 2.5 seconds

Notes: Alternative propulsion system contracted out by Raytheon to provide a smokeless propulsion system capable of being cold soaked to -65F and still fire. Was to be interchangeable with existing solid propellant motor for Sparrow III. Cancelled 1960 after 400 engines were produced after solid propellants capable of being cold-soaked to -65F were developed.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

RMI LR58 (Bullpup A Missile)

Propellants: RFNA / 50.5% Diethylnetramine, 40.5% UDMH, 9% acetonitrile
Thrust: 12,000 lbf for 1.9 seconds
Diameter: 12.1 inches
Length: 40.5 inches
Weight (loaded): 203 lb
Weight (Dry): 92 lb
Total Impulse of Unit: 22,800 lb-s

Notes: Consisted of a unitary fuel tank that was also part of the engine. Minimum storage life of five years. Reliability estimated at 0.9972

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

RMI LR62 (Bullpup B Missile)

Propellants: RFNA / 50.5% Diethylnetramine, 40.5% UDMH, 9% acetonitrile
Thrust: 30,000 lbf for 2.3 seconds
Diameter: 17.3 inches
Length: 61.2 inches
Weight (loaded): 563 lb
Weight (Dry): 205 lb
Total Impulse of Unit: 69,000 lb-s

Notes: Consisted of a unitary fuel tank that was also part of the engine. Minimum storage life of five years. Reliability estimated at 0.9972

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

RMI 1K Engine (Corvus ASM)

Propellants: Nitric Acid / Mixed Amine
Thrust: 1,030 lbf for 177 seconds

Notes: Work started 1957, first flight 1959. Cancelled by USN.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

General Electric

GE 13.5K Engine (Hermes A-1)

Propellant: LOX/ 75% Alcohol, Pressure Fed
Thrust: 13,500 lbf

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 21K Engine (Hermes A-3)

Propellant: LOX/ 75% Alcohol, Turbopump Fed
O/F Ratio: 1.12
Thrust (sl): 21,000 lbf at 200 ISP

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE X-400 Engine

Propellant: LOX/ RP-1
Thrust: 27,000 lbf

Notes: Designed as a follow on to the Hermes A-3 21K engine.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 80K Engine (Hermes C-1)

Thrust: 80,000 lbf

Notes: Designed and undergoing component testing when cancelled so that GE could concentrate on the 27.8K Vanguard engine.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 27.8K Engine (Vanguard Booster)

Propellants: LOX/Kerosene
Thrust:
27,000 lbf at 254 ISP

Notes: First US engine that used turbine exhaust for roll control.

(GE Vanguard Booster Engine)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 405H Vega Upper Stage Engine

Propellants: LOX/RP-1
Thrust (vac):
33,840 lbf at 300 ISP
Expansion Area Ratio (ε = Ae/At): 25

Notes: Designed for Atlas/Vega interim upper stage which was to be used before Atlas/Centaur became operational. First flight was scheduled for 1960, but the Atlas/Vega SLV was cancelled in favor of Atlas/Agena SLV.

(GE Vega Upper Stage Engine)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 412A Engine (North American X-15)

Thrust: 52,000 lbf

Notes: Ordered as backup engine for X-15 after problems with XLR-99 engine arose during development. Consisted of two thrust chambers of 26,000 lbf each. Effectively a modified version of the Vanguard engine. Cancelled.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 16K Plug Nozzle

Thrust: 16,000 lbf

Notes: Early test demonstrator for aerospike concept. Built and fired.

(Photo of 16K Spike)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 50K Plug Nozzle

Thrust: 50,000 lbf

Notes: Next-stage of demonstration for the aerospike concept. Built and fired.

(Photo of 50K Spike firing)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

GE 1000K Plug Nozzle

Thrust: 1,000,000 lbf

Notes: GE built sections of this nozzle design and tested the sections before management made the decision to terminate General Electric’s liquid propellant rocket division in mid-1966.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Curtiss-Wright

Curtiss-Wright XLR-25 (X-2)

Propellants: LOX / 75% Alcohol
Thrust:
15,000 lbf

Notes: Could be throttled as low as 2,500 lbf, because it had two thrust chambers. One ran at 10,000 lbf and the other at 5,000 lbf. Chamber pressures could be reduced in one or both chambers to support throttling. Designed with some input from Robert Goddard.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne

Navaho Phase I

Notes: Two German-built V-2 engines were reconditioned for this phase.

Rocketdyne XLR-41-NA-1 (Navaho Phase II)

Notes: Simplified V-2 engine made from US-built parts and built to US pipe size and fitting standards. Was water-flow tested, and as it was ready to be hot fired; the program was cancelled in favor of hot testing the early US-designed thrust chambers that would eventually lead to the Redstone NAA-75-110.

(Photo of XLR-41-NA-1 next to V-2 engine – 1.6 MB PNG)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne XLR-43-NA-1 (Navaho Phase III)

Propellants: LOX/Alcohol-Water
Thrust (sl):
75,000 lbf
Chamber Pressure: 300 psia
Rated Lifetime:
60 seconds

Notes: This engine formed the basis for the Redstone’s NA 75-110 A-1 to 7 engine.

(Photograph of XLR-43-NA-1)

References:
Historic American Engineering Record AL-129-A: Marshall Space Flight Center, Test Stand (1.04 MB PDF)
Propulsion: The Key to Space Travel by P.D. Castenholz and H.K. Griggs
AMC 23M: History of the Redstone Missile System by John W. Bullard

Rocketdyne XLR-43-NA-3 (Navaho Phase IV)

Thrust (sl): 120,000 lbf at 230 ISP
Dry Weight: 1,230 lb
T/W (sl): 97.5

Notes: First US engine to use “spaghetti” (tubular wall) configuration for the thrust chamber, which reduced weight by 50%.

Reference:
Encyclopedia of 20th Century Technology, Volume 2
Modern Engineering for Design of Liquid Propellant Rocket Engines by Harry Arbit

Rocketdyne XLR-71-NA-1 (G-26 Navaho)

Propellants: LOX / 92.5% Alcohol, 7.5% Water, Gas Generator
O/F Ratio: 1.375
Thrust (sl):
240,000 lbf @ 229 ISP
Thrust (vac): 278,000 lbf @ 265 ISP
Chamber Pressure: 438 psia
Dry Weight:
2,501 lbs
T/W Ratio (sl): 95.96
T/W Ratio (vac): 111.15
Engine Length: 117 inches
Engine Diameter: 77 inches
Expansion Area Ratio (ε = Ae/At): 4.6
Rated Lifetime: 65~ seconds

Notes: “Doubled up” version of XLR-43-NA-3 with two thrust chambers.

(XLR-71-NA-1 on Test Stand – 1.37 MB PNG)

References:
Rocketdyne: Powering Humans into Space by Robert S Kraemer
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne XLR-83-NA-1 (G-38 Navaho)

Thrust (sl): 405,000 lbf @ 245 ISP
Propellants
: LOX / JP-5
Rated Lifetime: 93 to 100 seconds

Notes: “Tripled up” version of XLR-43-NA-3 with three thrust chambers and modified with better cooling to allow the burning of hydrocarbon fuel.

(XLR-83-NA-1 Perspective Drawing)

References:
SM-64A Navaho SAC – 14 September 1955.
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne NAA 75-110 ‘A-Series’

Propellants: LOX / 75% Alcohol, 25% Water, Gas Generator
O/F Ratio: 1.354
Thrust (sl):
78,000 lbf @ 218 ISP
Thrust (vac): 89,000 lbf @ 249 ISP
Chamber Pressure: 318 psia
Dry Weight:
1,478 lbs
T/W Ratio (sl): 52.77
T/W Ratio (vac): 60.21
Engine Length: 131 inches
Engine Diameter: 68 inches
Engine Throat Diameter: 15.5 inches
Expansion Area Ratio (ε = Ae/At): 3.61
Rated Lifetime: 117 seconds

Variants:

A-1

Two built and flown in Redstones RS-1 and RS-2.

A-2

Three built, flown in Redstones RS-3 to RS-5.

A-3

Five built, flown in Redstones RS-8 to RS-12.

A-4

Flown on Redstone RS-18, the first Chrysler-built missile on 14 March 1956. Was briefly tested with Hydyne fuel, which was composed of 60% UDMH and 40% diethylenetriamine (DETA).

A-5

Never Flown

A-6

Major Production Variant #1. Flown from 2 October 1957.

A-7

Major Production Variant #2. Flown from 24 June 1958.

Notes: Used on the Redstone missile. Based on Rocketdyne’s earlier Navaho engine; the XLR-43-NA-1. The designation referred to the manufacturer – North American Aviation, the thrust – 75,000 lbf, and the burn time – 110 seconds. Seven different types of engines in the family were developed – NAA 75-110-A-1 through A-7 for testing and evaluation of different components. All seven engines were interchangeable with each other, with only minor tubing modifications needed to mate the engine to the Redstone. Data shown above is for A-6/A-7 versions.

(Photograph of NAA 75-100 – Unknown Model Number)
(Drawing of NAA 75-110 – Unknown Model Number)

References:
Rocketdyne: Powering Humans into Space by Robert S Kraemer
Historic American Engineering Record AL-129-A: Marshall Space Flight Center, Test Stand (1.04 MB PDF)
AMC 23M: History of the Redstone Missile System by John W. Bullard
http://www.enginehistory.org/Museums/USSRC/USSRC_Redstone.shtml
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne LR43-NA-3 (MA-1 Atlas A/B Booster Engine)

Thrust (sl): 150,000 lbf at 245 ISP

Notes: Atlas Propulsion System, R&D Version – All three engines (2 boosters, 1 sustainer) were fairly independent of each other in this incarnation.

References:
The Development of Propulsion Technology for U.S. Space Launch Vehicles, 1926-1991 by J.D. Hunley
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne LR43-NA-5 (MA-1 Atlas B Sustainer Engine)

Thrust (sl): 54,000 lbf at 210 ISP

Notes: Atlas Propulsion System, R&D Version – All three engines (2 boosters, 1 sustainer) were fairly independent of each other in this incarnation.

References:
The Development of Propulsion Technology for U.S. Space Launch Vehicles, 1926-1991 by J.D. Hunley
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne LR89-NA-3 (MA-2 Atlas Booster Engine)

Thrust (sl): 154,500 lbf

Notes: Atlas Propulsion System, Production Version – Simplified version of MA-1 in which the booster engine turbopumps were mounted together so they could be powered by a single gas generator instead of two. Preferred for space launch vehicle variants of Atlas.

References:
The Development of Propulsion Technology for U.S. Space Launch Vehicles, 1926-1991 by J.D. Hunley
Characteristics Summary, SM-65D Atlas, February 1962
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne LR105-NA-3 (MA-2 Atlas Sustainer Engine)

Thrust (sl): 57,000 lbf

Notes: Atlas Propulsion System, Production Version – Simplified version of MA-1 in which the booster engine turbopumps were mounted together so they could be powered by a single gas generator instead of two. Preferred for space launch vehicle variants of Atlas.

References:
The Development of Propulsion Technology for U.S. Space Launch Vehicles, 1926-1991 by J.D. Hunley
Characteristics Summary, SM-65D Atlas, February 1962
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne LR89-NA-5 (MA-3 Atlas E/F Booster Engine)

Thrust (sl): 165,000 lbf at 250 ISP
Thrust (vac): Unknown at 290 ISP

Notes: Atlas Propulsion System, Production Version II – Developed at USAF request; put the turbopump and gas generator back on each booster so that each booster engine was identical and interchangeable, simplifying logistics.

References:
The Development of Propulsion Technology for U.S. Space Launch Vehicles, 1926-1991 by J.D. Hunley
Standard Missile Characteristics, SM-65F Atlas, January 1962
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne LR105-NA-5 (MA-3 Atlas E/F Sustainer Engine)

Thrust (sl): 57,000 lbf at 215 ISP
Thrust (vac): Unknown at 310 ISP

Notes: Atlas Propulsion System, Production Version II – Developed at USAF request; put the turbopump and gas generator back on each booster so that each booster engine was identical and interchangeable, simplifying logistics.

References:
The Development of Propulsion Technology for U.S. Space Launch Vehicles, 1926-1991 by J.D. Hunley
Standard Missile Characteristics, SM-65F Atlas, January 1962
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne E-1 / MA-4 (Titan Booster Engine)

Notes: Proposed single chamber E-1 for Titan. Never fully developed and never found any application.

References:
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne LR89-NA-7 (MA-5 Atlas Booster Engine)

Propellants: LOX/RP-1
O/F Ratio: 2.28
Thrust (sl): 165,000 lbf at 254 ISP
Thrust (vac): 187,900 lbf
Chamber Pressure: 578 psia
Expansion Area Ratio (ε = Ae/At): 8

Notes: Upgraded version of MA-2 propulsion system for Atlas Space Launch Vehicles.

References:
NASA SP-8120 Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)
GD/C-BTD65-017 External Design Loads – Operational Centaur Vehicles (AC-6 through AC-15) (1 May 1965)
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

Rocketdyne LR105-NA-7 (MA-5 Atlas Sustainer Engine)

Propellants: LOX/RP-1
Thrust (sl): 57,000 lbf
Thrust (vac): 81,900 lbf

Notes: Upgraded version of MA-2 propulsion system for Atlas Space Launch Vehicles.

References:
GD/C-BTD65-017 External Design Loads – Operational Centaur Vehicles (AC-6 through AC-15) (1 May 1965)
Rocketdyne: Powering Humans into Space by Robert S. Kraemer (Pages 110 to 111)

RS-27 (MA-5A Atlas II Booster Engine)

Thrust (sl): 414,000 lbf

Notes: Used on Atlas II. Replaced the RS-56A Booster Engines with RS-27. Vernier engines were deleted, their functions assumed by hydrazine thrusters on the Atlas II interstage.

References:
NASA Historical Data Book: Volume VII

RS-56SA (MA-5A Atlas II Sustainer Engine)

Thrust: 60,500 lbf

Notes: Used on Atlas II.

References:
NASA Historical Data Book: Volume VII

Thor MB-1

Thrust: 135,000 lbf

Notes: Used for initial flight testing. Had a small onboard propellant tank for initial starting, along with pyrotechnic ignitors in the thrust chamber and for the gas generator startup.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Thor MB-2?

Notes: Used hypergolic ignitor in thrust chamber and two pyrotechnic ignitors to start the gas generator. Was licensed to Rolls-Royce in 1955 back when it still had a conical exhaust nozzle and formed the basis for the Rolls Royce BLUE STREAK engine.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne LR79-NA-x (Thor MB-3 Block I)

Thrust (sl): 152,000 lbf

Notes: Developed from the Atlas MA-3 booster engine via deleting a thrust chamber, essentially.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton
To Reach the High Frontier: A History of U.S. Launch Vehicles

Rocketdyne LR79-NA-11 (Thor MB-3 Block II)

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 Area Ratio (ε = Ae/At): 8
Diameter: 49 inches

Notes: First used on Delta A in 1962.

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)
Rockets by Joseph A. Angelo
To Reach the High Frontier: A History of U.S. Launch Vehicles

Rocketdyne LR79-NA-x (Thor MB-3 Block III)

Thrust (sl): 172,000 lbf

Notes: First used on Delta E in 1965. Improved reliability and slightly improved thrust.

Reference:
To Reach the High Frontier: A History of U.S. Launch Vehicles

Rocketdyne AR-1

90% Hydrogen Peroxide and JP-4 or JP-5

Notes: Pressure fed. Tested on FJ-4F Fury. Not a success due to tank mass from a pressure fed system.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne AR-2

90% Hydrogen Peroxide and JP-4 or JP-5

Notes:

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne AR2-1

90% Hydrogen Peroxide and JP-4 or JP-5

Notes:

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne AR2-2

90% Hydrogen Peroxide and JP-4 or JP-5

Notes:

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne AR2-3

90% Hydrogen Peroxide and JP-4 or JP-5
O/F ratio: 6.5
Thrust (100,000 ft): 6,600 lbf at 246 ISP
Chamber Pressure: 560 psia
Expansion Ratio: 12
Dry Weight: 225 lbs

Notes: Pump-fed, allowing significantly lighter oxidizer tank with fuel coming directly from aircraft tanks.

FJ-4F
Used on 103 flights with a total operating time of 3.5 hours of rocket operation. Allowed FJ-4F to reach Mach 1.31 in level flight.

F-86:
Used on 31 flights to Mach 1.22 in horizontal flight.

NF-104A
Used for 302 flights for a total operating time of 8.6 hours.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne NOMAD

Propellants: Liquid Fluorine (LF2) / N2H4 (Hydrazine)
O/F Ratio: 13
Thrust (vac): 12,000 lbf at 368 ISP (Apollo Proposal)
                      12,000 lbf at 357 ISP (Upper Stage Proposal)
Diameter: 6 feet? (for stage or engine? No idea)
Chamber Pressure: 150 psia
Expansion Area Ratio (ε = Ae/At): 20

Notes: Development work began in 1958 as part of a complete integrated upper stage design effort by Rocketdyne. Was briefly investigated by General Electric for their Apollo proposal after Rocketdyne made an unsolicted proposal.

(Photo of NOMAD Stage)

References:
Project Apollo: A Feasibility Study of an Advanced Manned Spacecraft and System – Final Report: Volume IV: On-Board Propulsion, Book 1 – Text and Appendix P-C (15 May 1961) by General Electric (1.94 MB PDF Excerpt)
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne RS-2101A Engine (Modified with 60:1 Nozzle Extension)

Propellants: NTO (N2O4) / MMH (Monomethylhydrazine)
O/F Ratio: 1.55
Thrust (vac): 300~ lbf at 291.1 ISP
Dry Weight: 18 lbs (+/- 0.3 lbs)
T/W Ratio (vac): 16.66
Chamber Pressure: 114.7 psia
Expansion Area Ratio (ε = Ae/At): 60
Diameter: inches

Notes: During a typical mission the engine will be subjected to about 25 starts for a total firing time of about 3500 seconds. There will be one long duration firing of approximately 3000 seconds. Duration of intermediate firings will be from 0.4 to 100 seconds. Between firings the chamber is expected to cool to an equilibrium temperature below 200 F. Thus the objective will be to design an engine which will survive one (1) mission duty cycle (25 starts with one 3000 second firing) within specified performance limits: and three (3) mission duty cycles without a catastrophic failure

(RS-2101 / RS-2101A Comparison Picture)
(RS-2101A Drawing)

References:
Modified RS2101 Rocket Engine Study Program: Final Report (7 December 1971) on NTRS

Rocketdyne S-3D

Propellants: LOX / RP-1 Kerosene Gas Generator
O/F Ratio: 2.4
Thrust (sl):
150,000 lbf @ 248 ISP
Thrust (vac): 174,000 lbf @ 288 ISP
Chamber Pressure: 527 psia
Dry Weight:
2,008 lbs
T/W Ratio (sl): 74.7
T/W Ratio (vac): 86.65
Engine Length: 142 inches
Engine Diameter: 67 inches
Expansion Area Ratio (ε = Ae/At): 8
Rated Lifetime: 180 seconds

References:
Rocketdyne: Powering Humans into Space by Robert S Kraemer

Rocketdyne S-3E

Propellants: LOX/RP-1 Kerosene

Notes: Used on Thor/Delta.

Rocketdyne S-3F

Propellants: LOX/RP-1 Kerosene

Rocketdyne E-1

Propellants: LOX/RP-1 Kerosene
Thrust (sl):
400,000 lbf

Notes: Proposed engine for Saturn. Cancelled in favor of H-1.

Rocketdyne E-1 (Uprated Dynasoar RBS Version)

Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.25
Thrust (sl): 468,000 lbf at 256 ISP
Thrust (vac): 524,000~ lbf at 290~ ISP (inexact)
Dry Weight: 4,000~ lbs
T/W Ratio (sl): 117
T/W Ratio (vac): 131
Chamber Pressure: 790 psia
Expansion Area Ratio (ε = Ae/At): 12
Length: 156 inches
Diameter:
71 inches

Notes: Proposed uprated version of E-1 suggested for a fly-back reusable booster by Boeing for its Dyna-Soar Weapons System. Downrated from 500 klbf baseline for manned flight safety, as the booster would have been piloted.

References:
MD 59-44: Operational Dyna Soar Recoverable Booster Study – Selected Booster (1 March 1959) (980~ kb PDF Excerpt)

Rocketdyne H-1 (Engines H-1001 through H-1009)

Propellants: LOX/RP-1 Kerosene
Thrust (sl):
165,000 lbf

Notes: These engines were non-flight capable engines and used for ground testing with Jupiter hardware accessories installed for quick delivery and testing.

References:
(H-1 Data Sheet by Rocketdyne – 500~ kb PDF)
Saturn H-1 Engine Design Features and Proposed Changes – 21 September 1959 – Via heroicrelics.com (LINK)
Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967) (14.1 MB PDF)

Rocketdyne H-1 (SA-1 through SA-10)

Propellants: LOX/RP-1 Kerosene
Thrust (sl): 165,000 lbf at 252.7 ISP (SA-1 through SA-4)
Thrust (sl):
188,000 lbf at 257 ISP (SA-5 through SA-10)

Notes: The first engines were downrated to 165,000 lbf, even though they had been designed and built for a thrust level of 188,000 lbf to improve reliability. Later Saturn I launches used the full thrust of this engine.

(H-1 Drawing)

References:
(H-1 Data Sheet by Rocketdyne – 500~ kb PDF)
Saturn H-1 Engine Design Features and Proposed Changes – 21 September 1959 – Via heroicrelics.com (LINK)
Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967) (14.1 MB PDF)
The Development of Propulsion Technology for U.S. Space Launch Vehicles, 1926-1991 by J.D. Hunley
MPR-SAT-WF-61-8 Saturn SA-1 Flight Evaluation (14 December 1961)
MPR-SAT-FE-64-19 Results of the Seventh Saturn I Launch Vehicle Test Flight: SA 7 (31 January 1966)

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

Development Began: September 1958
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 Area Ratio (ε = Ae/At): 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)
AIAA Paper 68-569 Development of LOX/RP-1 Engines for Saturn/Apollo Launch Vehicles
Launch Vehicle Engines Project Development Plan
MA001-A50-2H (1 January 1967) (14.1 MB PDF)
(H-1 Data Sheet by Rocketdyne – 500~ kb PDF)

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
Thrust (vac): 228,800 lbf at 301 ISP
Dry Weight (Inboard/Outboard Engines): 2,100 lbs
T/W Ratio (sl): 97.61
T/W Ratio (vac): 108.95
Length: 8.8 Feet (105.6”)
Diameter:
4.9 Feet (58.8”)
Expansion Area Ratio (ε = Ae/At): 8

Notes: The H-1C was the fixed inboard engine variant, while the H-1D was the outboard gimbaling engine variant. They were identical with the exception of the exhaust system (H-1C has a partial aspirator, the H-1D has full aspirators) and vehicle attach-hardware. After the end of the Apollo program, several dozen H-1s in unfired condition were transferred to the Delta EELV program, where they became the RS-27.

(H-1 Side Drawing)

References:
Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967) (14.1 MB PDF)
(H-1 Specifications)
NASA SP-8120 Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)
“Extended Long Tank Delta” by Ed Kyle spacelaunchreport.com (LINK)

Rocketdyne RS-27

Thrust (sl): 207,000 lbf
Expansion Area Ratio (ε = Ae/At): 8

Notes: Rebadged H-1 engines from the Apollo program that were refurbished and slightly tuned up as they were no longer flying on man-rated vehicles. 83 were flown on the Delta 2000/3000/5000 series rockets.

References:
“Extended Long Tank Delta” by Ed Kyle spacelaunchreport.com (LINK)
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne RS-27A

Propellants: LOX/RP-1 Kerosene
O/F Ratio: 2.245
Thrust (sl): 200,000 lbf at 255 ISP
Thrust (vac): 237,000 lbf at 302 ISP
Dry Weight: 2,528 lbs
T/W Ratio (sl): 79.11
T/W Ratio (vac): 93.75
Chamber Pressure: 700 psia
Length: 149 inches
Diameter:
67 inches
Expansion Area Ratio (ε = Ae/At): 12

Notes: Further development of the H-1/RS-27 and simplified for cheaper production. Has two vernier engines which produce 1,012 lbf of thrust each and provide roll control throughout vehicle flight.

(Rocketdyne RS-27A Brochure – 500~ kb PDF)

References:
History of Liquid Propellant Rocket Engines
by George Paul Sutton

Rocketdyne J-2 (SA-201 through SA-203)

Propellants: LOX/LH2
O/F Ratio: 5
Thrust (vac): 200,000 lbf at 418 ISP
Weight: 3,480 lbs
T/W Ratio (vac): 57.47
Expansion Area Ratio (ε = Ae/At): 27.5

References:
(J-2 Specifications)

Rocketdyne J-2 (SA-204 through SA-207 and SA-501 through SA-503)

Propellants: LOX/LH2
O/F Ratio: 5.5
Thrust (vac): 225,000 lbf at 419 ISP
Weight: 3,480 lbs
T/W Ratio (vac): 64.65
Expansion Area Ratio (ε = Ae/At): 27.5

References:
(J-2 Specifications)

Rocketdyne J-2 (SA-208 to final Saturn IB, SA-504 to final Saturn V)

Development Begun: September 1960
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 Area Ratio (ε = Ae/At): 27.5

(J-2 Side Views)
(J-2 Cleaned up Side View)

References:
(J-2 Specifications)
Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967) (14.1 MB PDF)
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
NASA SP-8120
Liquid Rocket Engine Nozzles (July 1979) (140~ kb PDF excerpt)
(J-2 Fact Sheet from Saturn V News Reference)

Rocketdyne J-2X (First Program)

Notes: The goal of the first J-2X program, which was operational from 1964 to 1968, were to simplify the engine, the stage equipment associated with pre-launch operations, reduce engine ground support equipment. In addition, the engine was to have higher thrust and ISP to enable greater operational flexibility. All these goals were to be achieved without significant modifications to the S-II or S-IVB stages.

To achieve these goals, the J-2X program investigated the following areas of improvement over the J-2:

There were three extensible nozzle concepts evaluated within the ground rules of adding no more than 750 pounds to engine weight, being able to fit into the S-IC/S-II or S-II/S-IVB interstages, and having a 2 second (or less) actuation time. They were the:

Due to the premature cancellation of the J-2X program, an immediate downselect of nozzle extension concepts was forced.

Because the Aerobell concept was similar in concept to the XLR-129 nozzle extension program already under research by the USAF, it was eliminated.

The telescoping extension concept had already been modeled in full scale on a nonoperational J-2 engine to demonstrate the actuation of the spring driven scissor arms. Due to the complexity of the multiple section telescoping assembly, this concept was also eliminated, leaving only the Airmat extension for further study.

The Airmat extension would have been attached to the existing engine nozzle at the 27.5:1 expansion ratio point, and upon inflation, would have extended it to an expansion ratio of 48:1. Payload gains for the S-IVB stage were estimated to be 3,200 pounds from this effort.

Two full scale J-2 Airmat nozzle extensions were fabricated; one to 48:1, the other to a more conservative 41.3:1 extension. It was planned to fire a J-2 with first the 41.3:1 unit then the 48:1 unit at Arnold Engineering Development Center, but J-2 engine firings at that facility were prematurely canceled before the nozzle extensions could be fabricated. It was decided to use gaseous nitrogen on a non operational J-2 to test the nozzle extension process.

Ultimately, the lessons learned from the experimental J-2X were applied to the production-capable J-2S design.

(J-2X Nozzle Extension Concepts Studied)
(J-2X Airmat Nozzle Extension Drawing)

Reference:
AIAA Paper 67-978: Milestones in Cryogenic Liquid Propellant Rocket Engines
NASA TM-X-64690 Chemical Propulsion Research at MSFC: Research Achievements Review Vol IV No 6 (August 1972) (18.2~ MB PDF Excerpt)

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: 1,245 PSIA
Expansion Area Ratio (ε = Ae/At):
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 topping (tap-off) cycle to eliminate the gas generator. The starting bottle system was eliminated, being replaced with a solid propellant turbine starter (SPTS) system with multiple sealed SPTS units allowing multi-restart capability in space.

Throttling capability was added as an option for applications other than the Saturn Program. Throttling was achieved via simply closing down the hot gas valve, which in turn reduced the available power to the turbines. This scheme was verified during hot-fire testing down to 6:1 throttling levels.

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 (21,400> seconds of mainstage duration and 6,600> seconds of low-thrust idle mode operation). Engine lifetime was demonstrated to 12,000 seconds and 30 starts between overhaul.

In 1972 an engine qualification program consisting of a 36 month test program using one engine and a spare engine was proposed. It was rejected and the J-2S program terminated before it could be flight certified.

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.

Mid 1990s Restart Efforts:

It was concluded that the J-2S could be produced entirely from existing drawings with the exception of replacing obsolete, outdated, and unavailable electronics. All of the materials used then were still available, as were their manufacturing processes. However, 67 changes were suggested for producibility/reliability enhancements for the restart.

Studies also showed that the turbopumps could be uprated to high enough chamber pressures to deliver 320 klbf of vacuum thrust merely by increasing the thickness of structural housings.

If go-ahead was given in FY94, engine testing would begin in Q4 FY96 and certification would be complete by mid-FY98. Non-recurring costs for the restart would have been $245 million in FY92 dollars. Average engine costs were undetermined, as it depended on production rate. (see Costing image below).

Resources:
(J-2S Side Drawing)
(J-2S Sheet from Rocketdyne Circa 1993 for SEI Restart Studies – 579 kb GIF)
(J-2S Restart Costing and Schedule – 175 kb GIF)
(J-2 vs J-2S, McDonnell Phase B Shuttle Study)
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 3: Alternate Propulsion Subsystem Concepts (NAS8-39210) DCN 1-1-PP-02147
J-2S Restart Study: Task Final Report DR-4 (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)

AIAA Paper 67-978: Milestones in Cryogenic Liquid Propellant Rocket Engines

Rocketdyne F-1 (SA-501 through SA-503)

Propellants: LOX/RP-1 Kerosene Gas Generator
Thrust (sl): 1,500,000 lbf at 260 ISP
Weight: 18,416 lbs
T/W Ratio (sl): 81.45
Engine Length: 220 inches
Nozzle Diameter: 148 inches
Expansion Area Ratio (ε = Ae/At): 16

Notes: Designed for a mission duration of 150 seconds.

(NASA F-1 Engine Specification Photo)

References:
Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967) (14.1 MB PDF)

Rocketdyne F-1 (SA-504 to SA-515)

Development Begun: January 1959
Ready (Flight Certified):
9 May 1967
Ready (Man Rated): April 1968 (Apollo 6)
Propellants: LOX/RP-1 Kerosene Gas Generator
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 Area Ratio (ε = Ae/At): 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.)

(F-1 Top and Side Drawing)
(NASA F-1 Engine Specification Photo)

References:
Remembering the Giants: Apollo Rocket Propulsion Development (NASA SP-2009-4545)
AIAA Paper 68-569
Development of LOX/RP-1 Engines for Saturn/Apollo Launch Vehicles
(F-1/F-1A Comparison Sheet from Rocketdyne Circa 1993 for SEI Restart Studies – 473 kb GIF)
(F-1 Fact Sheet from Saturn V News Reference)

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 Area Ratio (ε = Ae/At):
16

Notes: The F-1A program began on 1 April 1968 and resulted in an engine with the following simplifications over the production “Qualification II” F-1:

Areas of Comparison

Qualification II Engine

F-1A Engine

Stage Connections and Engine Interconnections

39

21

Servicing Connections

37

24

Acceptance Instrumentation

129

8

Flight Instrumentation

34

6

Propulsion Redlines

16

6

Additional Ground Test Restrictions

7

3

Monitored Countdown Events (Launch Interlocks)

15

2

Throttling: The F-1A would have been capable of being calibrated to produce between 1.35 and 1.8 mlbf of thrust via adjustment of various components of the engine. Throttling ability would have been up to 300,000 lbf below the calibration rating except as limited by the minimum rating.

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.

(F-1A Simplified Drawing)

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)
R-8102 F-1A Task Assignment Program Final Report (15 January 1970)
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)

Pratt & Whitney Rocketdyne / Dynetics F-1B

Propellants: LOX/RP-1 Kerosene
Thrust (sl): 1,800,000 lbf
Expansion Area Ratio (ε = Ae/At): 12

Notes: Simplified, modernized F-1A proposed for NASA’s Space Launch System (SLS). Can throttle down to 1.3 mlbf without problems, and uses new processes where possible to simplify engine production.

(F-1[B] Diagram from Aviation Week)

Rocketdyne X-1

Notes: Developed from 1957-1959 to investigate product improvements for Rocketdyne engines. Amongst the innovations tested on this engine were:

In addition to these enhancements, the X-1 was used to investigate off-design operating conditions, as it was operated at 198K, despite being designed for 165K.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne X-8

LOX-LH2 Gas Generator Cycle.
Thrust: 90,000 lbf at 408 ISP

Notes: Developed 1961. Tested before the J-2 design was finalized.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne 1961 50K E-D Engine

NTO/Aerozine-50
Thrust: 50,000 lbf
Chamber Pressure: 300 psia

Notes: Ran for 300 seconds with an uncooled nozzle.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne 1961 10K E-D Engine

NTO/Aerozine-50
Thrust: 10,000 lbf
Chamber Pressure: 225 psia
Expansion Ratio: 300

Notes: Had a cooled nozzle to allow prolonged operation. Tested at Arnold ETC in Tennessee at pressure ratios between 400 and 10,000. Data correlated well with predictions.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne 9.9K Hydrolox E-D Engine

LOX/LH2
Thrust (Vac): 9,900 lbf

Notes: Tested at Arnold ETC from sea level to 100,000 feet.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

NASA/USAF Advanced Development Program (ADP) Aerospike Engine
a/k/a
Advanced Cryogenic Rocket Engine Program, Aerospike
a/k/a
Advanced Engine Aerospike
a/k/a
250K Hydrolox Aerospike by Rocketdyne

Propellants: LOX/LH2
O/F Ratio: 6.0
Thrust (vac): 250,000 lbf at 450.1 ISP (throttleable down to 50,000 lbf at 443.8 ISP)
Thrust (sl): 206,000 lbf at 371 ISP (throttleable down to 32,400 lbf at 289 ISP)
Weight: 3,950 lbs
T/W Ratio (vac): 63.29
T/W Ratio (sl): 52.15
Chamber Pressure: 1,500 psia (323 psia at 20% thrust level)
Length: 56.55 inches
Diameter: 100 inches
Expansion Area Ratio (ε = Ae/At):
74.1

Notes: Program began on 1 March 1966 to develop a preliminary design for a LH2/LOX engine with the following specifications:

The above specifications listed are for the Demonstrator Module (DM). The Flight Module (FM) would have been the same 100” diameter, but length would have reduced to 48.3 inches, along with weight dropping to 2,939 lbs, and the ε increasing to 75.8; making vacuum ISP increase slightly to 452 seconds.

The demonstrator module was fired 34 times at chamber pressures from 200 to 1,055 PSIA; while components were test fired at chamber pressures from 340 to 2,045 PSIA.

(ADP Assembly Components)
(ADP Preliminary Characteristics)
(ADP Characteristics)
(ADP Cycle)
(ADP Top/Side Drawing)
(ADP Top/Side Drawing 2)
(ADP 250K Tube Wall Test and Firing at Nevada Field Laboratory)
(AEE 250K Test Firing)
(AEE 250K ISP Graph)

References:
AFRPL-TR-67-280 Final Report, Advanced Cryogenic Rocket Engine Program, Aerospike Nozzle Concept, Volume I (January 1966) (17.1~ MB PDF)
NASA TM-X-64690 Chemical Propulsion Research at MSFC: Research Achievements Review Vol IV No 6 (August 1972) (18.2~ MB PDF Excerpt)

Rocketdyne Split Combustor Linear Aerospike Engine


Concept 1
(Hybrid Engine)

Concept 2
(All-Hydrolox)

Propellants

LOX/RP-1 (Outer)
LOX/LH2 (Inner)

LOX/LH2 (Outer)
LOX/LH2 (Inner)

Type of Cycle

Gas Generator (Outer)
Staged Combustion (Inner)

Staged Combustion (Outer)
Staged Combustion (Inner)

O/F Ratio

2.8 (Outer)
7.0 (Inner)

7.0 (Outer)
7.0 (Inner)

Mode 1 Thrust (sl)

4,000,000 lbf at 321.3 ISP

4,000,000 lbf at 389.5 ISP

Mode 1 Thrust (vac)

4,500,000 lbf at 367.2 ISP

4,500,000 lbf at 438.7 ISP

Mode 2 Thrust (vac)

1,650,000 lbf at 455.4 ISP

1,650,000 lbf at 455.4 ISP

Chamber Pressure

2,000 psia (Outer)
2,500 psia (Inner)

2,500 psia (Outer)
2,500 psia (Inner)

Expansion Area Ratio

40:1 (Mode 1)
114:1 (Mode 2)

40:1 (Mode 1)
114:1 (Mode 2)

Engine Dry Mass

44,250 lbs

50,380 lbs

T/W Ratio (sl) (Mode 1)

90.39

79.39

T/W Ratio (vac) (Mode 2)

37.28

32.75

Engine Envelope

346 inches wide
172 inches high
84.5 inches long

346 inches wide
175 inches high
84.5 inches long

Notes: Each engine is made up of four modules, each module’s dimensions are:

Width: ½ that of total engine width
Height: ½ that of total engine height
Length: 84.5 inches

Mode 1 Operation: Both Combustors Fire
Mode 2 Operation: Inner Combustor only

(Engine Altitude vs ISP Tables for Mode 1 Operation)
(SCLA Concept 1 Engine Cycle)
(SCLA Concept 1 Engine Drawing)
(SCLA Concept 2 Engine Drawing)

References:
NASA CR-135231 Final Report: Linear Aerospike Engine Study (November 1977) (10.9 MB PDF)

USAF Advanced Maneuvering Propulsion System (AMPS)

Main Engine (Toroidal Aerospike)
     Propellants: LF2/LH2
     O/F Ratio: 12
     
Thrust (vac): 30,000 lbf at 460 ISP
     
Chamber Pressure: 650 psia
     Expansion Area Ratio (ε = Ae/At):
60

Secondary Engine (Bell)
     Propellants: LF2/LH2
     O/F Ratio: 12
     
Thrust (vac): 3,300 lbf at 457.5 ISP
     
Chamber Pressure: 750 psia
     Expansion Area Ratio (ε = Ae/At):
60

Notes: Designed for a high energy upper-stage that would replace the Transstage on the Titan IIID. More specifically, the AMPS would be:

A high performance LF2/LH2 maneuvering propulsion system has been defined, which provides an in orbit delta V capability in excess of 21,000 ft/sec with a 2000 pound payload. The system, which is designed for launch into a 100 nautical mile polar orbit with the Titan IIID, weighs 18,000 pounds fully loaded, has instant response and multi-start capability, while providing a nominal space residence duration of 14 days. Longer orbital stay times (up to 180 days) can be accommodated with some design modification without significantly degrading performance.”

(AMPS Engine Diagram and Specifications)
(AMPS Upper-Stage Specifications)
(AMPS Upper Stage Diagram)

References:
AFRPL-TR-68-20 Advanced Maneuvering Propulsion Technology Program (Second Quarterly Report) (May 1968)

Rocketdyne Advanced Modular Propulsion System (AMPS)

Thrust: 20,000 lbf expander cycle, annular aerospike

Notes: Developed and tested in the early 1970s.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne RS-23 Shuttle Orbit Maneuver Engine (OME)

Propellants: NTO / Aerozine-50
Thrust (vac): 6000 lbf at 313 ISP

Notes: Rocketdyne's proposal for STS orbiter OME. Program started 1972 mostly on company money. Lost out to Aerojet’s proposal.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne RS-30 Advanced Space Engine (ASE)

Propellants: LOX/LH2
O/F Ratio: 6.0
Thrust (vac): 20,000 lbf at 473.43 ISP (Mainstage Mode), 1,332 lbf at 442.5 ISP (Pumped Idle Mode)
Weight: 337 lbm
T/W Ratio (vac): 59.34
Chamber Pressure: 2,233 psia (Mainstage Mode); 176 psia (Pumped Idle Mode)
Engine Length: 92.248 inches
Engine Nozzle Diameter: 48.484 inches
Expansion Area Ratio (ε = Ae/At): 400

Notes: Resulted from a 10-month study to select an optimum configuration for a 20 klbf Hydrolox Topping-cycle engine. Used as a test bed to estimate scaled down features/components of SSME and to also demonstrate the feasibility of a high pressure hydrolox upper stage engine.

(Advanced Space Engine Drawing)

References:
NASA CR-121236 Advanced Space Engine Preliminary Design by A.T. Zachary (October 1973) (13.6 MB PDF)
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne RS-25 Space Shuttle Main Engine (SSME)

Ready (Man Rated): April 1981 (at 100% RPL)
                                            1983 (at 104% RPL – 109% FPL available for in flight emergencies)
                                            1998 (at 104.5% RPL for ISS construction)
                                            2001 (at 109% FPL certified for regular duty)

Cycle: Staged Combustion
Propellants: LOX/LH2
O/F Ratio: 6.0

Thrust (sl): 373,500 lbf at 362 ISP (100% RPL)
                    390,000 lbf at 364.8 ISP (104% RPL)
                    418,000 lbf at (109% EPL/FPL)

Thrust (vac): 468,400 lbf at 454 ISP (100% RPL)
                      488,800 lbf at 452.9 ISP (104% RPL)
                      512,300 lbf at 452.3 ISP (109% EPL/FPL)

Weight: 6,970 lbs (RS-25 Phase I SSME) ~Approximate, extracted from graph~
              7,110 lbs (RS-25A Phase II SSME)
~Approximate, extracted from graph~
              7,460
lbs (RS-25B Block I SSME) ~Approximate, extracted from graph~
              7,620 lbs (RS-25C Block IIA SSME) ~Approximate, extracted from graph~
              7,748 lbs
(RS-25D Block II SSME)
              8,148 lbs (RS-25    Block III SSME) ~Approximate, based on projected weight growth of 400 lbs over Block II~

Chamber Pressure: 3,319 PSIA at 109% RPL for Block I SSME? (Uncertain)
                                 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 Area Ratio (ε = Ae/At): 77.5 (or sometimes given as 69)

Notes: Rocketdyne began the SSME contract in August 1972, with full scale engine ignition testing occurring in June 1975. It took until March 1977 to get a SSME up to 100% of full power for 60 seconds of continuous operation. The first cluster test of three engines occurred in April 1978.

SSME specifications originally called for each engine to be able to survive no less than six excursions into 109% EPL over a 100 mission lifespan.

Following a careful study of fatigue following a STS redefinition, it was found that if engine life was limited to 55 missions, then there would be no limit to the number of excursions an engine could make into the 109% power regime; and thus Emergency Power Level was renamed to Full Power Level.

Even while STS-1 was in orbit, work began on certifying the SSME to 109% FPL; but after a few years, certification work was halted after many test engine failures on the stand. 109% power was still on the table as an option for an emergency abort scenario, but it was not certified for regular flight use.

As part of their proposal for the 1995 EELV competition, Boeing took Engine #2107, dropped it into the Mississippi river, dried it out, inspected it and then hot fired it to prove the viability of a recoverable engine module utilizing SSMEs (or SSME technology variant engines) for their EELV proposal.

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. Please note that the model designations, such as -25A, -25B, -25C were backfitted much later in the program.

RS-25 Phase I First Manned Orbital Flight (FMOF) SSME: Used on STS-1 through STS-5. Certified to 100% RPL (470,000 lbf vac). Many components of the engine had limited operational life and had to be replaced well before the 55-flight requirement.

RS-25 Phase I Full Power Level (FPL) SSME: First Flight STS-6 on 4 April 1983. Last flight STS-51L on 28 January 1986. 147 design changes were required to enable FPL operation. 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. Estimated to have a 1/404 catastrophic failure risk in a three engine cluster.

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. Estimated to have a 1/608 catastrophic failure risk in a three engine cluster.

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. The last Block IIA flight was STS-109 in March 2002. Estimated to have a 1/999 catastrophic failure risk in a three engine cluster.

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. Estimated to have a 1/1283 catastrophic failure risk in a three engine cluster.

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.

Block III SSME: Proposed around 2000 with a goal of first flight by 2005. The original proposals had it being capable of 120% RPL to improve abort scenarios. This was later changed to improve safety at nominal power levels without cost risks associated with changes to HP pumps and powerhead via reducing stresses on the turbopumps. Estimated to have a 1/2333 to 1/2960 catastrophic failure risk in a three engine cluster. Block III would have used most of the Block II components and included:

Specific impulse would have dropped by 0.5 to 1.5 seconds, and engine weight would have increased by about 400~ pounds per engine.

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 (6.52 MB PDF) 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 (610~ kb Word Document)
Space Shuttle Main Engine – Thirty Years Of Innovation by Fred H. Jue (36.5~ MB Powerpoint)
Atlantis STS-104 SSME Flight Readiness Review, 28 June 2001 (2.1 MB PDF)
SSME: Space Shuttle Main Engine Brochure (506~ kb PDF)
Space Shuttle Main Engine Orientation (June 1998) (9.19~ MB PDF)
The Space Shuttle Main Engine and the Solid Rocket Booster (14 October 1980)
Block III SSME Upgrade Project Overview, July 2000 (700~ kb PDF)
Space Shuttle Main Engine: The First Ten Years by Robert E. Biggs (Development of SSME to STS-1)
Space Shuttle Main Engine: The Second Decade by Robert E. Biggs
Space Shuttle Main Engine: The First Twenty Years and Beyond by Robert E. Biggs (collects two above papers)
Lessons in Systems Engineering – The SSME Weight Growth History (2.65 MB PDF)

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 (or 165 psia)
Length: 47 inches
Nozzle Exit Diameter: 34 inches
Expansion Area Ratio (ε = Ae/At):
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)

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

Ready (Man Rated): July 1966 (AS-203)
Propellants:
NTO (N2O4) / MMH
O/F Ratio: 1.27
Thrust (vac):
72 lbf at 274 ISP
Chamber Pressure: 101 psia
Expansion Ratio (Ac/At): 40

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 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 X-8

LOX-LH2 Gas Generator Cycle.
90,000 lbf at 408 ISP.

Notes: Developed 1961. Tested before the J-2 design was finalized.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne XLR-132

Propellants: NTO (N2O4) / MMH
Type: Gas Generator
O/F Ratio: 2.0
Thrust (vac):
3,750 lbf at 342.4 ISP
Engine Weight: 114 pounds
T/W Ratio (vac): 32.89
Chamber Pressure: 1500
Engine Length: 52 inches
Nozzle Diameter: 26 inches (exit)
Expansion Ratio (Ac/At):
400

References:
Rocketdyne/Westinghouse Nuclear Thermal Rocket Engine Modeling by Jim Glass (22 October 1992)
Orbital Transfer Vehicle: Concept Definition and System Analysis Study NAS8-36108 First Quarterly Review Briefing (30 Oct 1984) (476~ kb PDF excerpt)

Rocketdyne RS-2 (Mariner '71 PBCS)

Notes: Uses a machined beryllium thrust chamber.

References:
Report of the Ad Hoc Committee on Beryllium, National Materials Advisory Board, Division of Engineering, National Research Council

Rocketdyne RS-14 (Axial Engine, MM3 PBCS)

Notes: Axial thruster of the Minuteman III Post Boost Control System (PBCS). Uses a machined beryllium thrust chamber.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton
Report of the Ad Hoc Committee on Beryllium, National Materials Advisory Board, Division of Engineering, National Research Council

Rocketdyne Advanced Maneuvering Propulsion Technology Program (AMPTP) Aerospike

Liquid Hydrogen/Liquid Oxygen
O/F: 5.5
Thrust (vac): 25,000 lbf at 470.4 ISP
Dry Mass: 398 lbs
T/W Ratio (vac): 62.81
Chamber Pressure: 1,000 psia
Expansion Ratio: 200
Engine Diameter: 65.25 inches
Engine Length: 23 inches

Notes: This appears to be a further outgrowth of an earlier Liquid Fluorine/Liquid Hydrogen aerospike technology program conducted under the Advanced Maneuvering Propulsion Technology Program's aegis.

Work started in 1970 under USAF contract F04611-67-C-0016 to design and develop a 25,000 lbf hydrolox aerospike flightweight thrust chamber. Contract was terminated on 31 December 1975 after the thrust chamber was extensively damaged in a test stand fire. The thrust chamber was then subsequently repaired by NASA under contract NAS3-20076.

The chamber consisted of 24 regeneratively cooled combustor elements around the periphery of a regeneratively cooled nozzle.

(Drawing of AMPTP Aerospike)

Reference:
NASA CR-135169 Aerospike Thrust Chamber Program, Final Report (Dec 1976) (7.2 MB PDF)

Rocketdyne L-1 Linear Test Bed (LTB) Engine

LOX/LH2
O/F Ratio: 5.5
Chamber Pressure: 1,200 PSI
Thrust (sl): 200,000 lbf at 345.5 ISP
Thrust (vac): 263,791 lbf at 455.7 ISP
Expansion Ratio: 115.82 (estimated)
Dimensions: 120" long by 120" wide by 96" high

Notes: Work began April 1970 and concluded June 1972 with 44 successful tests over 3,113 seconds of hot fire. Test 624-028 ran for 529 seconds mainstage duration. Power package consisted of spares from the J-2/J-2S program and required at worst minor modifications for use on the linear test bed program. The thrust chamber(s) were two sets of 10 combustors for twenty in total. Overall, the L-1 LTB was a ‘battleship’ engine which used a heavy duty thrust frame and other non-flightweight components to validate the concept of a linear aerospike.

(Chart which references it – 1.72 MB PNG)
(Drawing of L-1 LTB? – I have no idea if this is actually it, but it looks close)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton
Reusable Hydrocarbon Rocket Engine Maturity for USAF RBS (February 2012) (4.9 MB PDF)
Linear Test Bed Final Report, Volume I: Test Bed No. 1 (30 Aug 1972) (12.2~ MB PDF)
NASA TM-X-64690
Chemical Propulsion Research at MSFC: Research Achievements Review Vol IV No 6 (August 1972) (18.2~ MB PDF Excerpt)

Rocketdyne L-2 Linear Test Bed (LTB) Engine

LOX/LH2
O/F Ratio: 5.5
Chamber Pressure: 1,200 PSI
Thrust (sl): 100,000 lbf at 345.5 ISP
Thrust (vac): 131,895 lbf at 455.7 ISP
Expansion Ratio: 115.82 (estimated)

Notes: Work began April 1970 and concluded October 1972 with 29 tests over 1,199.5 seconds of hot fire including a 300 second mainstage run. The thrust chamber(s) were two sets of 5 combustors for ten in total. Overall, the L-2 LTB was a ‘lightweight’ engine which used a lighter weight thrust frame and other components to develop the technology further.

The thrust levels given above are based on a "chopped" down L-1 with half the combustion chambers. In actual tests, maximum thrust was actually 95K instead of 100K, due to to Rocketdyne using the L-2 LTB to explore the various regimes of linear aerospike capabilities, including thrust vectoring via turning on/off thrust cells as necessary instead of going for raw power.

(Chart which references it – 1.72 MB PNG)

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton
Reusable Hydrocarbon Rocket Engine Maturity for USAF RBS (February 2012) (4.9 MB PDF)
Linear Test Bed Final Report, Volume II: Test Bed No. 2 (27 March 1974) (6.6~ MB PDF)
Reusable Hydrocarbon Rocket Engine Maturity for USAF RBS (February 2012) (4.9 MB PDF)

RSX

Notes: 237klbf Kerolox space engine.

References:
Rocketdyne/Westinghouse Nuclear Thermal Rocket Engine Modeling by Jim Glass (22 October 1992)

Integrated Modular Engine (IME)

Notes: 30klbf Hydrolox space engine.

References:
Rocketdyne/Westinghouse Nuclear Thermal Rocket Engine Modeling by Jim Glass (22 October 1992)
Annular Nozzle Engine Technology by Al Martinez

Rocketdyne 1K LH2 Expander – Low Thrust Chemical Rocket Engine Study

Propellants: LOX/LH2
Thrust (vac): 1,000 lbf
Chamber Pressure: 500 psia
Engine Length: 45.25 inches
Nozzle Diameter: 22.56 inches
Expansion Area Ratio (ε = Ae/At): 500

(Engine Diagram)

References:
Low Thrust Chemical Propulsion by James M. Shoji

Rocketdyne 3K LH2 Expander – Low Thrust Chemical Rocket Engine Study

Propellants: LOX/LH2
Thrust (vac): 3,000 lbf
Chamber Pressure: 660 psia

(Engine Diagram)

References:
Low Thrust Chemical Propulsion by James M. Shoji

Rocketdyne OTV Engine – Advanced Expander Cycle Performance Baseline

Propellants: LOX/LH2
O/F Ratio:
6.0
Thrust (vac):
15,000 lbf at 482.5 ISP
Weight: 415 lb
T/W Ratio (vac): 36.14
Chamber Pressure: 1,610 psia
Engine Length:
60 inches (Nozzle Retracted) 117 inches (Nozzle Deployed)
Expansion Area Ratio (ε = Ae/At):
210 (Nozzle Retracted) / 625 (Nozzle Deployed)

Notes: These were the recommended engine specifications based upon 1980 state-of-the-art.

(Engine Drawing)

References:
Orbital Transfer Vehicle (OTV) Engine Study, Phase A – Extension, Bi Monthly Progress Report No 3 (11 Jan 1980)

Rocketdyne XRS-2200 Linear Aerospike Engine

Propellants: LOX/LH2 Gas Generator
O/F Ratio:
6.0
Thrust (sl):
204,420 lbf at 339 ISP
Thrust (vac): 266,230 lbf at 436.5 ISP
Chamber Pressure: 857 PSIA
Expansion Area Ratio (ε = Ae/At): 57.7
Throttling: 57-100%

Notes: The philosophy of the X-33 program was to accept increased risk in order to achieve lower costs and quicker schedule. To do so, the XRS-2200 program relied heavily on the experience gained from Rocketdyne’s testing of a linear test bed engine from 1970 to 1972. Where possible, the X-33 program used existing hardware and/or designs. The turbopumps and the gas generator were based on J-2 and J-2S engines. Component testing was used for design development, proving margins, and qualifications. Software was tested with hardware-in-the-loop. Single engine testing on the Stennis Space Center’s A-1 test stand was used to verify the design. The two flight engines, in their dual-engine configuration, had had a short ignition test and were about to start acceptance testing when NASA decided not to renew its involvement in the cooperative agreement. It was composed of 20 thrust cells.

References:
X-33 XRS-2200 Linear Aerospike Engine Sea Level Plume Radiation
History of Liquid Propellant Rocket Engines
by George Paul Sutton
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Rocketdyne RS-2200 Linear Aerospike Engine

Propellants: LOX/LH2 Gas Generator
Thrust (sl):
431,000 lbf

Notes: Larger, productionized version of the XRS-2200 intended for the VentureStar. Seven of these engines would have been used to lift the 2.2 million lb (GLOW) vehicle.

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Rocketdyne RS-44 Advanced Expander Cycle Engine

Propellants: LOX/LH2
Type: Single Expander
O/F Ratio: 6.0
Thrust (vac):
15,000 lbf at 463 ISP
Engine Weight: 342 pounds
T/W Ratio (vac): 43.85
Chamber Pressure: 1,540 psia
Expansion Ratio (Ac/At):
225

Notes: Rocketdyne proposed the following growth options for the RS-44:

Incremental Capacity Upgrade
Thrust (vac):
15,000 lbf at 481 ISP
Engine Weight: 461 pounds
T/W Ratio (vac): 32.53
Chamber Pressure: 1,540 psia
Expansion Ratio (Ac/At):
225 (Nozzle stowed), 625 (Nozzle extended)

Full Capacity Upgrade
Thrust (vac):
15,000 lbf at 492 ISP
Engine Weight: 407 pounds
T/W Ratio (vac): 36.85
Chamber Pressure: 2,052 psia
Expansion Ratio (Ac/At):
1,175 (Nozzle extended)

(RS-44 Engine Cycle)
(RS-44 Testbed Hardware)
(RS-44 Mockup Model)

Notes: Designed 1981 with a two piece extensible nozzle

Notes II: The final candidate in the existing concept category was the RS-44. This was an advanced expander cycle conventional bell nozzle engine experimentally tested by Rocketdyne in the mid-1980s. It may have had a 320:1 nozzle at one point during the development cycle.

References:
Rocketdyne/Westinghouse Nuclear Thermal Rocket Engine Modeling by Jim Glass (22 October 1992)
Advanced OTV Engine Concepts by A.T. Zachary
Orbital Transfer Vehicle: Concept Definition and System Analysis Study NAS8-36108 First Quarterly Review Briefing (30 Oct 1984) (476~ kb PDF excerpt)
History of Liquid Propellant Rocket Engines by George Paul Sutton

Rocketdyne/Westinghouse NERVA-Derived Engine Family

Thrust Level (klbf):

25

50

75

Weight (lbm)




T/W Ratio (w/Shield):

3.6

5.3

6

ISP (Sec)


870


Reactor Exit Temperature


2,550K


Chamber Pressure (psia):

748

784


Expansion Ratio

200

200

200

Nozzle Exit Diameter (m):

1.73

2.44

2.99

Overall Engine Length (m):

6

7.66

8.74

Notes: This family was designed using the XE-Prime design as a starting baseline. XE-Prime had an ISP of 710 seconds via bleed cycle, with 2270K reactor exit temperature and an expansion ratio of 10. ISP was gained through the following changes:

Baseline engine operating life was to be 1.5 hours of operational life at 2550K; or 4.5 hours of operational life if the chamber temperature was lowered to 2450K, which in turn reduced ISP to 850.

Future developmental possibilities included changing the fuel element to a composite type based on Nuclear Furnace data; which would have changed lifetime/ISP to 1.5 hours of operatinal life at 2700K and an ISP of 900; or 4.5 hours of operational life at 2550K and an ISP of 870.

To show how much things had come since 1972 NTR testing; Phoebus 1B operated at a chamber pressure of 750 psia, and a throat heat flux of 30~ BTU/in2 sec; while SSME operated at a chamber pressure of 3000 psia and a throat heat flux of 100~ BTU/in2 sec.

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

Rocketdyne RS-60

LOX/LH2 Closed Expander
Thrust (Vac): 60,000 lbf at 460 ISP
O/F Ratio: 5.8
Chamber Pressure: 1250 psia
Weight: 1,200 lbs
Area Ratio: 285

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D (they swapped it around)

Rocketdyne RS-68

Ready (Flight Certified): November 2002 (Unmanned Delta IV Launch)
Propellants: LOX/LH2
O/F Ratio: 6
Weight: 14,876 lbs (6,748 kg)

Full Power Mode
     Thrust (sl): 656,000 lbf at 357 ISP
     Thrust (vac): 751,000 lbf at 409 ISP
     Chamber Pressure: 1,420 psia

Minimum Power Mode
     Thrust (sl): 337,000 lbf at 357 ISP
     Thrust (vac): 432,000 lbf at 409 ISP
     Chamber Pressure: 815 psia

Expansion Area Ratio (ε = Ae/At): 21.5

Notes: Can operate at/successfully transition between full and minimum power levels on command from the launch vehicle. Developed as a cheap expendable engine building off the failed STME program. Took 4.7 years and $500 million to develop, compared to 7 years and $750 million for the Gillette Mach 3 disposable razor. Has 80% fewer parts and requires 92% less labor than than the RS-25 SSME during manufacture, resulting in a recurring cost only 7% that of the SSME.

One area this is evident in is in the turbo-machinery – the Block II SSME requires about 170~ individual parts for it’s LOX turbopump, and about 200~ for it’s fuel turbopump. RS-68 requires about 40~ parts for the fuel turbopump and 25~ for the LOX turbopump. This translates into significantly less drawings that have to be created, kept track of, and checked for dimensions with the finished products.

(RS-68 Side Diagram; Labelled)
(RS-68 Side Diagram; Unlabelled)
(J-2X and RS-68 to scale – 1.58 MB Transparent PNG)
(RS-68 Operating Schematic)

References:
RS-68 Propulsion System (623~ kb PDF)
AIAA 2002-4324
Propulsion for the 21st Century – RS-68 (7.8 MB DOC)
Propulsion for the 21
st Century – Cost Driven Design/Development/Delivery (Slides for AIAA 2002-4324) (12.6 MB Powerpoint)

RS-68A

“The RS-68A program consists of two major design changes. Engine thrust is increased 39,000 pounds force by modifying the turbine nozzles from axis-symmetric to 3 dimensional to reduce turbine blade loading and to expand the operational range of both the fuel and oxidizer turbopumps. Specific impulse is improved by increasing the number of the main injector combustion elements which improves mixing and combustion efficiency. Other AATS funded improvements to the RS-68A engine include a new bearing material that is more resistant to stress corrosion cracking, improved processing of the 2nd stage fuel turbopump blisk to reduce cracking potential, improved oxidizer turbopump chill sensor and an improved hot gas temperature sensor. An improved gas generator igniter that has less foreign object debris potential is also under development and expected to be included in the RS-68A certification program.”

Reference:
Ares V and RS-68B (December 2008)

RS-68B

Propellants: LOX/LH2
Thrust (vac): 797,000 lbf at 414.2 ISP

“NASA has requested three changes for the RS-68B for Ares V. First is to reduce the amount of free hydrogen at engine start to mitigate the potential for fire around the vehicle and need for added thermal protection. The second is to reduce the amount of helium purge gas used by the engine which currently taxes the Cape Canaveral Air Force Station (CCAFS) helium infrastructure from both a flow rate and total usage standpoint for the 3-engine Delta IV Heavy vehicle. The third change is to modify the ablative nozzle to accommodate the duration requirements of the Ares V mission.”

Reference:
Ares V and RS-68B (December 2008)

MB-35

Propellants: LOX/LH2
Thrust (vac): 35,000 lbf at 468 ISP
Mass: 760 lbm
T/W (vac): 46.05
Chamber Pressure: 1,500 psia

Notes: Down-rated version of MB-60 intended to focus better on the changed launch vehicle market.

References:
The MB-60 Cryogenic Upper Stage Engine - A World Class Propulsion System (1.85 MB PDF)
The Concept Design of a Split Flow Liquid Hydrogen Turbopump, Michael A. Arguello (1.88 MB PDF)

Pratt & Whitney / Mitsubishi Heavy Industries MB-60

Propellants: LOX/LH2 Expander Bleed Cycle
O/F Ratio: 5.4
Thrust (vac): 60,000 lbf at 467 ISP
Mass: 1,300 lbm
T/W (vac): 46.15
Chamber Pressure: 1,950 psia
Expansion Ratio: 300

Notes: Joint development between PWR and Mitsubishi Heavy Industries (MHI). Development began in 1998.

(Image of MB-60)

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D
The MB-60 Cryogenic Upper Stage Engine - A World Class Propulsion System (1.85 MB PDF)
The Concept Design of a Split Flow Liquid Hydrogen Turbopump, Michael A. Arguello (1.88 MB PDF)

Advanced Low Cost Engine (ALCE)

Type: Full Flow Staged Combustion (FFSCC)
Propellants
: LOX/LH2
Thrust (sl): 421,000 lbf at 395.4 ISP
Thrust (vac): 486,867 lbf at 457.2 ISP
Chamber Pressure: 4,000 psia
Weight: 4,413 lbs
T/W Ratio (sl):
95.39
T/W Ratio (vac): 110.32
Engine Length: 147 inches
Engine Diameter: 84 inches
Expansion Area Ratio (ε = Ae/At): 70.62

Reference:
Advanced Transportation System Studies – Technical Area 3: Alternate Propulsion Subsystem Concepts (158 kb PDF Excerpt)

Rocketdyne RS-72

Propellants: NTO/MMH
Thrust (vac):
12,454 lbf at 340 ISP
Mass: 303.75~ lbs (calculated from T/W Ratio)
T/W Ratio (vac): 41
Engine Length: 2.286m
Nozzle Diameter: 1.3m
Expansion Area Ratio (ε = Ae/At): 300

Notes: Collaborative effort between PWR and EADS Astrium.

(Image of RS-72)

References:
Project Artemis by Georgia Tech Aerospace Engineering Team 2006 (9.38 MB PDF)

Rocketdyne RS-76

Notes: A ‘clean-sheet’ LOX-RP-1 staged combustion cycle engine with a single turbopump assembly.

References:
Booster Main Engine Selection Criteria for the Liquid Fly-Back Booster
Propulsion System Advances that Enable a Reusable Liquid Fly-Back Booster (LFBB)

Rocketdyne RS-83

Propellants: LOX/LH2
Thrust (sl):
664,000 lbf at 395 ISP
Thrust (vac): 750,000 lbf at 446 ISP
Chamber Pressure: 2,800 psia
Weight: 12,700 lbs
T/W (sl): 52.28
T/W (vac): 59.05
Expansion Area Ratio (ε = Ae/At): 40
Engine Length: 171 inches
Engine Diameter: 115 inches
Throttling: 50-100%

Notes: Developed for the Space Launch Initiative (SLI) Program. Designed for a reusable launch vehicle with an engine lifetime of 100 missions and a maintenance check-up every 50 missions Please note that the specifications are for a prototype engine. These would have changed if it had gone into full-scale development (FSD). As of

(RS-83 Render)

References:
Main Engine Prototype Development for 2nd Generation RLV: RS-83 (12 April 2002) (10.4 MB Powerpoint)
FS-2002-09-141-MSFC Main Engine Candidates for a Second Generation Reusable Launch Vehicle (Sep. 2002) (535~ kb PDF)
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

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
Chamber Pressure: 2,800 psia
Weight: 17,919 lbs
T/W Ratio (sl):
59.38 lbf per lb
T/W Ratio (vac): 63.06 lbf per lb
Expansion Area Ratio (ε = Ae/At): 20
Throttling: 65-100%

Notes: Developed from 1997 to 2004 as a reusable LOX/RP-1 engine with a lifetime of about 100 flights and capable of throttling from 65% to 100% of rated thrust. 10 times the reliability of SSME. Cancelled by NASA in 2005 despite completing it’s Preliminary Design Review (PDR) and making significant inroads on key Critical Design Review (CDR) points.

(3D Render of RS-84)

References:
RS-84 Engine (2.76~ MB PDF)
Reusable Hydrocarbon Rocket Engine Maturity for USAF RBS (February 2012) (4.9 MB PDF)
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

RS-88 (First Program?)

Propellants: LOX/Kerosene
Thrust: 36,000 lbf

References:
Spaceplanes: From Airport to Spaceport by Matthew A. Bentley

RS-88 (Second Program?)

Propellants: LOX/Alcohol (Ethanol)
Thrust (sl): 50,000 lbf

Notes: Selected by Boeing for the CST-100 Launch Escape System (LES) engine.

(RS-88 Test Firing – 3.8 MB JPG)

References:
RS-88 Pad Abort Demonstrator Thrust Testing at NASA Marshall Space Flight Center, November-December 2003

Rocketdyne SE-10

NTO/ Aerozine-50
Thrust (vac): 10,500 lbf at 305 ISP

Notes: Rocketdyne's proposal for the LMDE, program started in 1963. Could deep throttle to 1,080 lbf.

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

P&W Rocketdyne J-2X (Second Program)

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 Area Ratio (ε = Ae/At): 92
Engine Length: 185 inches (4.699 m)
Nozzle Exit Diameter: 120 inches (3.048 m)
Rated Starts: 4
Rated Lifetime: 2,000 seconds

Note: Had to meet much more stringent requirements, including a restart in space after 90~ days cold. In 2007, the operational lifetime was eight restarts and 2,600 seconds – by 2011 this was down to 4 restarts and 2,000 seconds.

(J-2X Render, circa 2007)
(J-2X Render, circa 2007 – Alternate)
(J-2X and RS-68 to scale – 1.58 MB Transparent PNG)

References:
Presentation by Tracy Lamm of P&W Rocketdyne from 26 Feb 2007 (1.8 MB PDF)
J-2X Brochure (350~ kb PDF)

United Technology Center – United Aircraft (UA) Corporation

UA-1203 SRM

TBD

UA-1204 SRM

Diameter: 120 inches
Segments:
4
Rated Thrust: 880,000~ lbf

References:
AFSC Historical Publication Series 63-50-I History of the X-20A Dyna-Soar Volume I (Narrative)

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)
Centaur Mission Planners Guide (August 1971) (5.14 MB PDF)

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)

Morton-Thiokol / Allant Techsystems (ATK)

Thiokol Star 37B (TE-M-364-2) (Burner II)

Propellant Weight: 1,440 lbs

Notes: Referred to as “Standard” motor.

References:
CASD-NAS-73-032 Reusable Centaur Study: Volume II Final Report (15 March 1974)

Thiokol TE-M-364-3

Thrust: x lbf for x seconds at 288.5 ISP
Total Weight: 1,574 lbs
Propellant Weight: 1,440 lbs
Burnout Weight: 134 lbs

Notes: Utilizes the same rocket casing and nozzle as the Surveyor version.

References:
Advanced Planetary Probe Study, Final Technical Report: Volume 2: Spin-Stabilized Spacecraft for the Basic Mission by TRW (27 July 1966)

Thiokol TE-M-364-4

Propellant Weight: 2,300 lbs

Notes: Is the TE-M-364-2 motor with a 14” cylindrical section inserted between the domes of the -2 motor case. Referred to as “Growth” motor.

References:
CASD-NAS-73-032 Reusable Centaur Study: Volume II Final Report (15 March 1974)

Thiokol Star 26B Motor

Chamber Pressure: 623 psia average
Thrust:
7,784 lbf average thrust over 17.8~ seconds at 272.4 ISP
Total Impulse: 142,759 lb-sec
Loaded Weight: 575.6 lbs
Propellant Weight: 524 lbs
Burnout Weight: 50.3 lbs

(Drawing, Specifications and Thrust/Time Graph)

Thiokol C-1 Apollo Common Reaction Control System Engine

Propellants: N2O4 / MMH or Aerozine-50
Thrust (vac): 100 lbf at 301 ISP
Weight: 6.26 lb
T/W Ratio (vac): 15.97
Length: 17.29 inches
Diameter:
7.12 inches

Notes: The C-1 Engine Project was intended to provide a 80-100 lbf pressure fed engine with an ablative or radiatively cooled nozzle capable of meeting the collective requirements of the following applications:

Development began on 8 August 1964, with a six month competitive Definition/Demonstration phase from 5 March 1965 to September 1965 between TRW Systems Group and Reaction Motors Division of the Thiokol Chemical Corporation.

After winning the Definition/Demonstration phase, Reaction Motors Division of the Thiokol Chemical Corporation began the Development phase on 18 October 1965. At the time of the source [January 1967] this phase was to have continued for 21 months, and focused on the Apollo SM and LEM applications to allow a Flight Readiness Demonstration to be completed during the 15th program month and formal qualification to be completed at the end of the 21st program month.

(C-1 Engine Specifications)
(C-1 Engine Schematic)

References:
Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967) (14.1 MB PDF)

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 Area Ratio (ε = Ae/At): 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 Area Ratio (ε = Ae/At): 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)

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 Area Ratio (ε = Ae/At): 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 Ares I / V Five-Segment Motor (RSRMV)

Propellant: PBAN
Maximum Thrust (vac): 3,550,820 lbf
Propellant Weight: 1,373,518 lbs
Length (Motor): 1,868.25 inches
Case Diameter: 146.08 inches
Expansion Area Ratio (ε = Ae/At): 7.22

Notes: Will be used on Space Launch System [SLS] if it is not canceled.

References:
Ares I First Stage Design, Development, Test, and Evaluation (12.1~ MB PDF)

ATK CASTOR IV Strap-On SRM

TBD

ATK CASTOR IVA Strap-On SRM

TBD

Notes: Development/Qualification motors were fired in 1983. By switching to HTPB propellant for this, Delta II performance was improved by 11%.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK CASTOR IVA-XL Strap-On SRM

TBD

Notes: 8-foot extension of Castor IVA. First tested in 1992. A modified version is used on the Japanese H-IIA launch vehicle.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK CASTOR IVB Strap-On SRM

TBD

Notes: First in Castor IVA series to incorporate TVC. Developed for the European Space Agency’s Maxus sounding rockets.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK CASTOR 30 Motor

TBD

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK CASTOR 120 Motor

TBD

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK GEM-40 Strap-On SRM (Ground Ignited Fixed Nozzle)

Notes: Used on Delta II.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK GEM-40 Strap-On SRM (Air Ignited Fixed Nozzle)

Notes: Used on Delta II.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK GEM-40 VN Strap-On SRM (Ground Ignited Vectorable Nozzle)

TBD

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK GEM-46 Strap-On SRM (Ground Ignited Fixed Nozzle)

Notes: Used on Delta III.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK GEM-46 Strap-On SRM (Air Ignited Fixed Nozzle)

Notes: Used on Delta III.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK GEM-46 Strap-On SRM (Ground Ignited Vectorable Nozzle)

Notes: Used on Delta III.

References:
ATK May 2008 Solid Motor Catalog (6.85 MB PDF)

ATK GEM-60 Strap-On SRM (Vectorable Nozzle)

Propellant: HTPB
Motor Diameter:
60 inches
Motor Length: 518 inches
Total Length: 53 feet
Burn Time: 90.8 seconds
Chamber Pressure: 818 psia average
Total Impulse: 17,950,000 lbf-sec
Average Thrust: 197,539 lbf over burn time
Nozzle Exit Diameter: 43.12 inches
Expansion Ratio: 11
Thrust (sl): 280,000 lbf
Total Loaded Mass: 74,158 lb
Propellant Mass: 65,471 lb
Case Mass: 3,578 lb
Nozzle Mass: 2187 lb
Other Mass: 2922 lb
Burnout Mass: 8346 lbf

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)
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs
(2006) – Appendix D

ATK Solid Rocket Motor Upgrade (SRMU)

Notes: Developed for Titan IVB.

References:
GEM-60 Press Brochure (400 kb PDF)

Boeing

Boeing Xenon Ion Propulsion System (XIPS)

Size: 25 cm diameter
Propellant: Electric/Xenon
Thrust (vac): 79 millinewtons at 3,400 ISP consuming 2.2 kW (Low Power Station Keeping Mode)
                      165 millinewtons at 3,500 ISP consuming 4.5 kW (High Power Orbit Raising Mode)

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs
(2006) – Appendix D

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

McDonnell Douglas

ROOST Primary Engine

Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (vac): 1,160,000 lbf at 410 ISP
Thrust (sl): 1,000,000 lbf at 355 ISP
Weight: 9,000~ lbs
T/W Ratio (vac): 128.88
T/W Ratio (sl): 111.11
Chamber Pressure: 1,000 psia
Overall Length: 175 inches
Nozzle Exit Diameter: 116 inches
Expansion Area Ratio (ε = Ae/At): 16

Notes: Paper engine designed for the Douglas Reusable One-Stage Orbital Space Truck (ROOST) study in 1962.

(Illustration/Specs of ROOST Primary Engine)

ROOST Secondary Engine

Propellants: N2O4/MMH
O/F Ratio: 2.1
Thrust (vac): 10,000 lbf at 315 ISP
Weight: 120~ lbs
T/W Ratio (vac): 83.33
Chamber Pressure: 200 psia
Overall Length: 52.3 inches
Nozzle Exit Diameter: 22.6 inches
Expansion Area Ratio (ε = Ae/At): 30

Notes: Paper engine designed for the Douglas Reusable One-Stage Orbital Space Truck (ROOST) study in 1962.

(Illustration/Specs of ROOST Secondary Engine)

TRW Systems / Northrup Grumman

MIRA 150A / Back-up Surveyor Vernier Engine

Notes: Pintle-Injector Engine.

References:
AIAA 2000-3871:
TRW Pintle Engine Heritage and Performance Characteristics (1.33 MB PDF)

URSA-25R [Universal Rocket For Space Applications]

Propellants: N2O4/MMH or N2O4/Aerozine-50
Thrust (vac):
25 lbf

Notes: Pintle-Injector Engine.

Notes: Planned applications for the URSA family of engines included Gemini, Apollo, Dyna-Soar, Manned Orbiting Laboratory, and the Multi-Mission Bipropellant Propulsion System (MMBPS).

References:
AIAA 2000-3871:
TRW Pintle Engine Heritage and Performance Characteristics (1.33 MB PDF)

URSA-100R [Universal Rocket For Space Applications]

Propellants: N2O4/MMH or N2O4/Aerozine-50
Thrust (vac):
100 lbf

Notes: Pintle-Injector Engine.

Notes: Planned applications for the URSA family of engines included Gemini, Apollo, Dyna-Soar, Manned Orbiting Laboratory, and the Multi-Mission Bipropellant Propulsion System (MMBPS).

References:
AIAA 2000-3871:
TRW Pintle Engine Heritage and Performance Characteristics (1.33 MB PDF)

URSA-200R [Universal Rocket For Space Applications]

Propellants: N2O4/MMH or N2O4/Aerozine-50
Thrust (vac):
200 lbf

Notes: Pintle-Injector Engine.

Notes: Planned applications for the URSA family of engines included Gemini, Apollo, Dyna-Soar, Manned Orbiting Laboratory, and the Multi-Mission Bipropellant Propulsion System (MMBPS).

References:
AIAA 2000-3871:
TRW Pintle Engine Heritage and Performance Characteristics (1.33 MB PDF)

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 Area Ratio (ε = Ae/At):
47.5

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.

(LMDE Drawing)

References Looking For:
Characteristics of the TRW Lunar Module Descent Engine (31 January 1969) (Photo of Cover)

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)
NASA Technical Note TN D-7143
Apollo Experience Report – Descent Propulsion System (2.08~ MB PDF)

1965 Minimum Cost Design Space Launch Vehicle Program [Engine]

Propellants: N2O4/UDMH
Thrust: 250,000 lbf

Notes: 44 separate 250K hot fire tests were conducted (including steady-state tests of 66, 83 and 98 seconds duration), demonstrating dynamic combustion stability via “bomb” testing and evaluating performance and ablative chamber durability. All firings were performed at AFRPL, Edwards AFB, CA from Oct 1968 to Jan 1970.

References:
AIAA 2000-3871:
TRW Pintle Engine Heritage and Performance Characteristics (1.33 MB PDF)

TRW TR-106 Pintle Injector Engine

Propellants: LOX/LH2
Thrust (sl): 650,000~ lbf

Notes: Starting in the late 1990s, Northrop Grumman Space Technology (then TRW) undertook, using company funds, to design and build a large engine that could operate on either LOx/LH2 or LOx/RP-1. The engine was expected to replace solid propellant booster strap-ons with liquid propellant stages having on-command throttling shutdown and even restart. Liquid propellants were considered safer and more environmentally friendly. The engine also was envisioned for powering the first stage of expendable, or fly-back, boosters.

The engine designed and demonstrated in this effort was designated the TR-106. It had a planned sea-level thrust of 650,000 lb and was to be either pressure fed or operated with gas-generator-driven turbopumps in the propellant lines. The center pintle injector incorporated in the TR-106 engine can operate equally well using LOx with RP-1, ethanol, propane, methane, or LH2. This basic injector technology has a 40-year history of producing high-performance and totally stable combustion without baffles or quarter-wave acoustic chambers in engines with thrust ranging from 100 lb to 650,000 lb.

The concept was originally developed at Space Technology Laboratories (STL) in 1960 as a 20:1 throttling injector for a 500-lb-thrust space-maneuvering thruster using dinitrogen tetroxide (N2O4) with monopropellant hydrazine (N2H4). That throttling design was scaled up to 10,000 lb thrust with 10:1 throttling using N2O4 with N2H4 50-UDMH 50 (Aerozine-50) and was used in the lunar module descent engine (LMDE) for the Apollo program. A fixed-thrust version of the engine was used on the second stage of Thor-Delta in the 1980s, where it accomplished more than 65 completely successful launches.

The same basic pintle injector geometry has been tested at thrusts of 40,000 and 250,000 lb operating with N2O4/A-50; at 50,000 lb with LOx/RP-1; and at 40,000 and 650,000 lb with LOx/LH2.

(TR-106 Image)

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

TRW TR-107

Propellants: LOX/RP-1 Kerosene
Thrust: 1,000,000~ lbf

Notes: Northrop Grumman has carried out the detailed design of a 1-million-lb-thrust booster rocket engine utilizing LOX/HC propellants as part of NGLT under NASA’s SLI. The authorization to proceed on this design was awarded in March 2003. The primary goal for the TR-107 engine program was to continue development of an engine that would increase the safety, reliability, and affordability of next-generation reusable space launch and transportation vehicles

(TR-107 Render)
(TR-107 Render II and Details)

References:
FS-2002-09-141-MSFC
Main Engine Candidates for a Second Generation Reusable Launch Vehicle (Sep. 2002) (535~ kb PDF)
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

TRW TR-201

Propellant: NTO / Aerozine-50
Thrust: 9,900 lbf
Dry Weight: 298 lbs

Notes: Simplified, low cost derivative of the LMDE used as second stage on the Delta 2914 and Delta 3914 launch vehicles.

References:
AIAA 2000-3871:
TRW Pintle Engine Heritage and Performance Characteristics (1.33 MB PDF)
The Development of Propulsion Technology for U.S. Space-launch Vehicles: 1926 – 1991 by J. D. Hunley

TRW Low Cost Pintle Engine (LCPE)

Propellants: LOX/LH2
Thrust: 650,000 lbf

References:
AIAA 2000-3871:
TRW Pintle Engine Heritage and Performance Characteristics (1.33 MB PDF)

Northrup Gruppman TR-40

LOX/LH2 Split Expander Cycle
Thrust (Vac):
40,000 lbf

Notes: USET-funded.

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

Messerschmitt-Bölkow-Blohm (MBB) / DASA / EADS

Rocketdyne-Bölkow BORD

LOX/LH2

Notes: Name was an abbreviation of Bölkow-Rocketdyne and this was a jointly financed program that featured the first milled channel cooling jacket by Rocketdyne. The maximum chamber pressure reached during 12 tests was 283 bar (4,104.56 PSI).

References:
History of Liquid Propellant Rocket Engines by George Paul Sutton

Unknown Test Fired Concepts

0.4K H2O2 (HTP) Aerospike

8K LOX/RP-1 Aerospike

10K N2O2 (Nitric Oxide) / Aerozine-50 Aerospike

40K Hydrolox Aerospike

250K Hydrolox Aerospike

250K Hydrolox Linear Spike

125K Hydrolox Linear Spike

NASA/AEC NERVA Program

KIKI-B4D

Chamber Temperature: 1890 to 2130 K
Fuel Exit Temperature: 2222K
Vacuum ISP: 780
Time at Full Power: 1 minute
Fuel Type: UC2/Graphite
Core Power: 914 MW(t)
Thrust: 45,861 lbf (204 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

KIWI-B4E

Chamber Temperature: 1890 to 5100K
Fuel Exit Temperature: 2389K
Vacuum ISP: 820
Time at Full Power: 82.5 minute
Fuel Type: UC2/Graphite
Core Power: 914 MW(t)
Thrust: 45,861 lbf (204 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

NRX-A2

Chamber Temperature: 2090 K
Fuel Exit Temperature: >2200K
Vacuum ISP: 775
Time at Full Power: 3.4 minute
Fuel Type: UC2/Graphite
Core Power: 1100 MW(t)
Thrust: 55078 lbf (245 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

NRX-A3

Chamber Temperature: 2244 K
Fuel Exit Temperature: >2400K
Vacuum ISP: 820
Time at Full Power: 16.3 (or 18.3) minute
Fuel Type: UC2/Graphite
Core Power: 1100 MW(t)
Thrust: 55078 lbf (245 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

NRX-A5

Chamber Temperature: 2280 to 2333K
Fuel Exit Temperature: >2400K (820 vac ISP)
Time at Full Power: 29.6 minute
Fuel Type: UC2/Graphite
Core Power: 1100 MW(t)
Thrust: 55078 lbf (245 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

PHOEBUS-1A

Chamber Temperature: 2366 K
Fuel Exit Temperature: 2478 (835 vac ISP)
Time at Full Power: 10.5 minute
Fuel Type: UC2/Graphite
Core Power: 1340 MW(t)
Thrust: 66,993 lbf (298 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

NRX-A4 (NRX-E8T)

Chamber Temperature: 2264 to 2290K
Fuel Exit Temperature: >2400K (820 vac ISP)
Time at Full Power: 28.5 minute
Fuel Type: UC2/Graphite
Core Power: 1100 MW(t)
Thrust: 55,078 lbf (245 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

PHOEBUS-1B

Chamber Temperature: 2222 to 2290 K
Fuel Exit Temperature: 2445 (828 vac ISP)
Time at Full Power: 30 minutes
Fuel Type: UC2/Graphite
Core Power: 1340 MW(t)
Thrust: 66,993 lbf (298 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

NRX-A8

Chamber Temperature: 2300 to 2405 K
Fuel Exit Temperature: 2308 (805 vac ISP)
Time at Full Power: 62.7 minute
Fuel Type: UC2/Graphite
Core Power: 1100 MW(t)
Thrust: 55078 lbf (245 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

PEWEE-1

Chamber Temperature: 1835 K
Fuel Exit Temperature: 2550K (845 vac ISP) and 2750K (890 vac ISP)
Time at Full Power: 43 minute
Fuel Type: UC2/Graphite
Core Power: 500 MW(t)
Thrust: 24,954 lbf (111 kN)
Core Dimensions (52 inches long? 20 inches wide)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)
To the End of the Solar System: The Story of the Nuclear Rocket by James A. Dewar

XE-PRIME (XE’)

Fuel Type: UC2/Graphite
Chamber Temperature:
2,270K
Chamber Pressure: 559.84 PSI (3.86 MPa)
Fuel Exit Temperature: >2400K
ISP: 710.74 sea level, 820 Vacuum at 35.9 kg/sec
Time at Full Power: 7.8 minutes
Core Power: 1,140 MW(t)
Turbopump Power: 5.1 MW
Turbopump Speed: 22,270 RPM
Thrust: 55,078 lbf (245 kN)

(Nuclear Turbopumps)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)
Space Nuclear Power by Joseph A. Angelo, Jr and David Buden.

Nuclear Furnace 1 (NF-1)

Fuel Exit Temperature: 2450K (830 vac ISP)
Time at Full Power: 109 minute
Fuel Type: Composite/Carbide

Notes: Used as a cheap, efficient way to test promising fuel elements outside of a full up rocket engine.

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

PHOEBUS-2A


Design

Actual Testing

Chamber Temperature:

2500K

2256K

Fuel Exit Temperature:

2550K (840 vac ISP)

2308K (805 vac ISP)

Time at Full Power:


12.5 minutes

Fuel Type:

UC2/Graphite

UC2/Graphite

Core Power:

5000 MW(t)

4,100 MW(t)

Thrust:

250,212 lbf (1,113 kN)

205,251 lbf (913 kN)

References:
Rover/NERVA-Derived Near-Term Nuclear Propulsion: FY92 Final Review (2.27~ MB PDF)

NERVA I (“Iron Horse” Baseline)

Propellants: Nuclear/LH2 (1,110~ MWt core)
Core Dimensions: approx 35 x 52 inches

Notes: This was the early baseline for NERVA I.

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 at 41.9 kg/sec
Weight: 27,728 lbm [1] or 25,803 lbm [2]
T/W Ratio (vac): 2.7 [1] or 2.9 [2]
Core Dimensions: approx 35 x 52 inches
Chamber Pressure: 450 psia
Chamber Temperature: 4,250 deg R
Turbopump Speed: 23,920 RPM
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)
(NERVA I Pressure Vessel Information)
Alternate Nozzle Design Concepts (19 July 1968) (330~ kb PDF)
(Nuclear Turbopumps)
(NERVA I Nozzle)

References:
[1] SD 71-466-4:
Nuclear Flight System Definition Study. Phase III Final Report, Volume II, Part C. (April 1971)
[2] NASA TM 105252:
An Overview of Tested and Analyzed NTP Concepts (September 1991)
Space Nuclear Power by Joseph A. Angelo, Jr and David Buden.

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 239 lb/sec.
Weight: 28,530 lbs (25,540 lbm for bare engine, 1,940 lbm for internal 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
Expansion Area Ratio (ε = Ae/At): 100

Notes: This was intended to be the “advanced” engine, based upon the PHOEBUS line of ground tested reactors, in particular PHOEBUS 2A. This design 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...with the exception of a manned Mars mission.

(NERVA II Specifications)
(NERVA II Primary Propulsion Module Drawing)
(NERVA II Dimensional Envelope)
(NERVA II Start-up/Shut-down Graphs)
(NERVA II Artwork with projected range of specifications – November 1968)

References:
Boeing Integrated Manned Interplanetary Spacecraft Concept Definition – Volume IV (January 1968)

Los Alamos Small Nuclear Engine (SNE) for the Space Shuttle (STS)

Propellant: Slush LH2 / Nuclear Full Flow Topping Cycle
Thrust (vac): 16,186 lbf (72 kN) at 874.8 ISP at 8.5 kg/sec
Chamber Temperature: 2,695K
Chamber Pressure: 449.61 PSIA (3.1 MPa)
Core Power: 365 MW(t)
Core Length: 37 inches
Core Diameter: 37 inches
Weight: 5,600 lbs
T/W Ratio (vac): 2.89
Turbopump Power: 0.9 MW
Turbopump Speed: 46,950 RPM
Engine Length: 123 inches (Folded Nozzle), 174 inches (Unfolded Nozzle)
Nozzle Diameter: inches
Expansion Area Ratio (ε = Ae/At): 100

Notes: The design constraints for this system were the Space Shuttle’s 15x60 foot cargo bay and 65,000 lb payload capacity. LANL used the PEWEE engine as the basis for their design. Due to the constraints of the cargo bay, the radiation shield would have been internal to save a few inches; consisting of a six inch slab of borated zirconium hydride with holes drilled into it to allow slush LH2 to flow into the core.

The core itself would have used PEWEE type zirconium tubes and used a 6 inch thick beryllium reflector to promote criticality. To save almost two feet in length, a two part nozzle was chosen. The first part was a fixed regeneratively cooled nozzle. The second part was a swing-in place graphite or molybdenum nozzle extension.

LANL planned for one reactor and seven engine tests, starting in 1976 and ending in 1982. One test would last for an hour at 875 ISP, another for two hours at 860 ISP. LANL thought that a second generation SNE would have an ISP over 900 seconds and might possibly have dual mode operation – it would provide injection thrust, then switch over to providing 10-25 kW(e) for the payload for up to five years. Thoughts were that a third generation system could reach over 1,000 ISP.

McDonnell Douglas’s Reusable Nuclear Stage (RNS) would have had a mass of 17,783 kg, of which 12,814 kg would be useable propellant, allowing for a burn time of 1,500 seconds. If a 18.3m long propellant module (PM) massing 23,181 kg, of which 21,265 kg was useable propellant was attached to the RNS, a burn time of 4,000 seconds would be achieved.

(SNE Shuttle Stage Drawing)
(SNE Drawing)
(SNE Key Points)
(SNE Fuel Element)
(SNE Reactor Core)
(SNE Diagram and State Points)
(SNE Mass Estimates)
(SNE-powered Stage for Shuttle by McDonnell Douglas)
(Nuclear Turbopumps)

References:
To the End of the Solar System: The Story of the Nuclear Rocket by James A. Dewar
Space Nuclear Power by Joseph A. Angelo, Jr and David Buden.

Miscellaneous Consortium Engines

NASA Advanced OTV Propulsion Technology Program Goals for the 1990s (circa 1983)

Propellants: LOX/LH2
O/F Ratio: 6.0
Thrust (vac): 10,000 to 25,000 lbf design point range at 520 ISP
Throttle Ratio: 30:1
Weight: 360 pounds
T/W Ratio: 27.7 to 69.4
Engine Length (stowed): 40 inches
Service Life:
      500 cycles and/or 20 hours of operation before overhaul
      100 cycles and/or 4 hours of operation without service

Notes: Studies began in 1981 with contracts to Aerojet, Pratt & Whitney and Rocketdyne to define propulsion concepts for an Advanced Orbit Transfer Vehicle. The three concepts are described below in tabular form. Aerojet’s selection of the 3K thrust level was to facilitate a multiple engine installation, which reflected Aerojet’s approach to reliability for a man-rated OTV.


Aerojet

Pratt & Whitney

Rocketdyne

Thrust

3,000 lbf at 482 ISP

15,000 lbf at 486 ISP

15,000 lbf at 492 ISP

Cycle

Expander H2-O2

Expander H2

Expander H2

Chamber Pressure

2,000 psia

1,500 psia

2,000 psia

Overall Length

77 inches

40 inches (nozzle retracted)
120 inches (nozzle deployed)

40 inches (nozzle retracted)
154 inches (nozzle deployed)

Engine Weight

126 pounds

450 pounds

435 pounds

Expansion Area Ratio
(ε = Ae/At)

1,200

640

1,300

Nozzle Exit Diameter

30 inches

64 inches

78 inches

Drawings

Aerojet Concept Drawing

P&W Concept Drawing

Rocketdyne Concept Drawing

Notes:


Later became Advanced Expander Cycle Engine (AEE) after further development.


(Overall Comparison Viewgraph)

References:
NASA TM 83419: Advanced Propulsion Concepts for Orbital Transfer Vehicles by L.P. Cooper (June 1983) (682 KB PDF)
NASA TM 83624: Propulsion Issues for Advanced Orbit Transfer Vehicles by L.P. Cooper (February 1983)
NASA TM 87069 Status of Advanced Orbital Transfer Propulsion by L.P. Cooper (October 1985)

NASA MC-1 (Fastrac) Engine

Type: Gas Generator, Pump-Fed
Propellants
: LOX/RP-1
O/F Ratio: 2.17
Thrust (vac):
   59,958 lbf @ 295 seconds (Bantam)
   63,939 lbf @ 314 seconds (X-34)
Weight: Unknown
T/W Ratio: Unknown
Chamber Pressure: 652 psia
Expansion Area Ratio (ε = Ae/At)
   15:1 (Bantam)
   30:1 (X-34)

Notes: Developed originally for the X-34. Later became the SpaceX Merlin engine.

Reference:
Development Status of the NASA MC-1 (Fastrac) Engine (661 kb PDF)

USAF Maneuvering Space Propulsion System (MSPS)

Thrust (vac): 45,000 lbf at 460 ISP with LOX/LH2
                      45,000 lbf at 470 ISP with LF2/LH2

Mass: 640 lbs with LOX/LH2
            470 lbs with LF2/LH2

T/W Ratio (vac): 70.31 with LOX/LH2
                             95.74 with LF2/LH2

References:
TM-67-1013-2 Preliminary Design of a Cryogenic Planetary Propulsion Module (Image of coversheet and relevant page)

1970 Space Shuttle Booster Engine Baseline

Propellants: LOX/LH2
O/F Ratio:
6.0
Thrust (sl):
400,000 lbf at 383 ISP
Thrust (vac): 462,000 lbf at 442 ISP
Weight: 5,200 lb
T/W Ratio (sl): 76.92
T/W Ratio (vac): 88.84
Chamber Pressure: 3,000 psia
Engine Length:
160 inches
Nozzle Diameter:
75 inches
Expansion Area Ratio (ε = Ae/At):
53

(Drawing and Specifications)

References:
NASA TM X-52876: Space Transportation System Technology Symposium: Volume IV – Propulsion (July 1970)

1970 Space Shuttle Orbiter Engine Baseline

Propellants: LOX/LH2
O/F Ratio:
6.0
Thrust (vac):
480,000 lbf at 459 ISP
Weight: 7,000 lb
T/W Ratio (vac): 68.57
Chamber Pressure: 3,000 psia
Engine Length:
260 inches
Nozzle Diameter:
146 inches
Expansion Area Ratio (ε = Ae/At):
200

(Drawing and Specifications)
(Possible Rocketdyne Mockup Model)

References:
NASA TM X-52876: Space Transportation System Technology Symposium: Volume IV – Propulsion (July 1970)

“Ultimate” Orbital Transfer Vehicle (OTV) Engine

Ready: Mid-1990s
Propellants:
LOX/LH2
Cycle: Expander
Thrust (sl):
15,000 lbf at “Greater than 480 ISP” [481.4 ISP calculated in Rocket Propulsion Analysis]
Engine Length: 40 inches (nozzle stowed)
Chamber Pressure: 2,000 psia
Expansion Area Ratio (ε = Ae/At):
1300 (or more)

Notes: Capable of 20 hours of service-free life, deep 30:1 throttling and has an advanced health monitoring system. This was envisioned circa 1985 in a briefing by a Rocketdyne engineer.

(Technological Roadmap to “Ultimate” OTV Engine)

References:
Advanced OTV Engine Concepts by A.T. Zachary

Space Transportation Booster Engine (STBE)

Ready: Late 1980s to Early 1990s
Propellants:
LOX/JP-4
O/F Ratio:
2.8
Thrust (sl):
1,500,000 to 2,000,000 lbf at 289 ISP
Weight: 16,340 to 24,160 lbs
T/W Ratio (sl): 82.78 to 91.79
Chamber Pressure: 2,000 psia
Engine Length:
199 to 226 inches
Nozzle Diameter:
116 to 131 inches
Expansion Area Ratio (ε = Ae/At):
25

(Drawing/Specification/Reference)

Space Transportation Main Engine Model 481 (STME-481)

Propellants: LOX/LH2
O/F Ratio: 6
Thrust (sl): 397,000 lbf at 380.4 ISP (Stowed)
Thrust (vac):
      468,000 lbf at 449 ISP (Stowed)
      481,000 lbf at 461 ISP (Extended)
Weight: 7,142 lbs
T/W Ratio (sl): 55.58
T/W Ratio (vac):
      65.52 (Stowed)
      67.34 (Extended)
Chamber Pressure: 3,006 PSIA
Engine Length:
      139 inches (Stowed)
      219 inches (Extended)
Nozzle Diameter:
      76.2 inches (Stowed)
      126.3 inches (Extended)
Expansion Area Ratio (ε = Ae/At):
      55 (Stowed)
      150 (Extended)

Notes: This version featured an extensible nozzle like RL10B.

(Drawing/Specification/Reference)

Space Transportation Main Engine (STME)

Ready: 1990s
Cycle: Gas Generator, Pump-Fed
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 Area Ratio (ε = Ae/At):
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. Could operate at two thrust levels – 100% and 70%. Cost goals were $3.5M per engine by the 500th unit. [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) (alternate version PDF)

SpaceX

Merlin

Thrust: 75,000 lbf
Chamber Pressure: 850 psia

References:
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

OKB-x Saturn

RD-57 (РД-57, Ракетный Двигатель-57)

OKB-1 Korolev

RD-56 (РД-56, Ракетный Двигатель-56)



RD-100 (РД-100, Ракетный Двигатель-100)

Flow Rate: 57.8 kg/sec alcohol, 74 kg/sec LOX
Chamber Pressure: 1.62 MPa
Chamber Temperature: 2,300 C
Exhaust Velocity: 2,130 m/sec
Thrust (sl): 267 kN

(RD-100 Drawing)

References:
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

RD-101 (РД-101, Ракетный Двигатель-101)

Flow Rate: 70.6 kg/sec alcohol, 102.3 kg/sec LOX
Thrust (sl): 363 kN at 214 ISP
Thrust (vac): 402 kN at 240 ISP
Chamber Pressure: 2.16 MPa
Minimum Nozzle Diameter: 400mm
Nozzle Mouth Diameter: 740mm

(RD-101 Drawing)
(RD-101 Plumbing Diagram)

References:
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

RD-103 (РД-103, Ракетный Двигатель-103)

Flow Rate: 80.5 kg/sec alcohol, 115.9 kg/sec LOX
Thrust (sl): 432 kN at 244 ISP
Thrust (vac): 500 kN at 251 ISP
Chamber Pressure: 2.44 MPa
Minimum Nozzle Diameter: 400mm
Nozzle Mouth Diameter: 810mm

(RD-103 Drawing)

References:
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

RD-105 (РД-105, Ракетный Двигатель-105)

Flow Rate: 55 kg/sec fuel, 149 kg/sec LOX
O/F Ratio: 2.7
Thrust (sl): 55,000 kgf at 260 ISP
Thrust (vac): 64,000 kgf at 302 ISP
Throat Diameter:
278.8mm
Exit Diameter: 1050mm

(RD-105/106 Chamber Dimensions)

References:
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

RD-106 (РД-106, Ракетный Двигатель-106)

Flow Rate: 55 kg/sec fuel, 149 kg/sec LOX
O/F Ratio: 2.7
Thrust (sl): 53,000 kgf at 250 ISP
Thrust (vac): 66,000 kgf at 310 ISP
Throat Diameter: 278.8mm
Exit Diameter: 1260mm

(RD-105/106 Chamber Dimensions)

References:
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

RD-110 (РД-110, Ракетный Двигатель-110)

Flow Rate: 131.3 kg/sec fuel, 347.9 kg/sec LOX
O/F Ratio: 2.65
Thrust (sl): 120,000 kgf at 244 ISP
Thrust (vac): 140,000 kgf at 285 ISP
Weight (wet): 2,100 kg
T/W (sl): 57.14 (wet engine)
T/W (vac): 66.66 (wet engine)

(RD-110 Chamber Dimensions)

References:
The Germans and the Development of Rocket Engines in the USSR by Olaf H. Przybilski (1.03 MB PDF)

RO-95

Propellants: LOX/LH2
Thrust: 10,000 kgf at 475 ISP

Notes: Development began in 1988, with ground fire testing planned for 1991-92. Canceled while still in preliminary design.

References:
AIAA 2006-4904 The First Russian LOX-LH2 Expander Cycle LRE: RD0146 by V.Rachuk and N.Titkov

Chemiautomatics Design Bureau (CADB) RD-0146

Propellants: LOX/LH2 Full Expander
O/F Ratio: 6
Thrust (vac): 22,000 lbf (10,000 kgf) at 451 ISP
Weight: 534 lbs

Notes: Intended for the Proton and Angara upper stages.

References:
AIAA 2006-4904 The First Russian LOX-LH2 Expander Cycle LRE: RD0146 by V.Rachuk and N.Titkov
A Review of United States Air Force and Department of Defense Aerospace Propulsion Needs (2006) – Appendix D

RD-0120

OKB-276 Kuznetsov / NK Engines Company

Kuznetsov NK-9 / НК-9 (8Д517)



Kuznetsov NK-15 / НК-15 (11Д51)





N-1 First Stage Blok A (блока А)

Kuznetsov NK-15V / НК-15В (11Д52)



Notes: Used in N-1 Second Stage Blok B (блока Б)

Kuznetsov NK-19 / НК-19 (11Д53)



Kuznetsov NK-21 / НК-21 (11Д54)



Kuznetsov NK-33 / НК-33 (11Д111)

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 Area Ratio (ε = Ae/At): 27
Chamber Pressure: 2,113.19 PSI (145.70 bar)

Notes: Developed for the N-1F Stage 1 (Blok A).

Kuznetsov NK-43 / НК-43 (11Д112)

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 Area Ratio (ε = Ae/At): 70

Notes: Developed for the N-1F Stage 2 (Blok B).

OKB-456 Glushko / NPO Energomash

RD-270 (РД-270, Ракетный Двигатель-270) (8Д420)

Propellants: N2O4/UDMH
O/F Ratio: 2.67
Thrust (sl): 1,410,001~ lbf (639,566 kgf) [6272 kN] at 301 ISP
Thrust (vac): 1,509,142~ lbf (684,535.49 kgf) [6713 kN] at 322 ISP
Weight: 7,429.58 lbs (3,370 kg) dry
Chamber Pressure: 3,785~ psia (26.1 MPa)
T/W Ratio (sl): 189.78
T/W Ratio (vac): 203.12
Engine Length:
4.85 meters
Engine Diameter:
3.3 meters
Expansion Area Ratio (ε = Ae/At):
Unknown
Throttle Control:
95-105%

Notes: Development began in June 1962, eventually selected for UR-700 rocket. First engine test fire was made in October 1967. Work suspended in August 1969, ultimately canceled on 31 December 1970 along with all other work on the UR-700. If development had continued, it would have been flight qualified sometime in 1972.

References:

Russian Language Website (http://www.lpre.de/energomash/RD-270/index.htm)
NPO Energomash Website; Engine Listing (http://www.npoenergomash.ru/engines/)

There apparently was a written scholarly article about the RD-270:

К ИСТОРИИ РАЗРАБОТКИ ЖИДКОСТНОГО РАКЕТНОГО ДВИГАТЕЛЯ РД-270 ДЛЯ РАКЕТЫ-НОСИТЕЛЯ УР-700
© В.С.Судаков, Р.Н.Котельникова
© Государственный музей истории космонавтики им. К.Э. Циолковского, г. Калуга
Секция "История ракетно-космической науки и техники"
2002 г.

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 Area Ratio (ε = Ae/At):
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)

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

Ready (Flight Certified): 24 May 2000 (Atlas III First Flight)
Propellants: LOX/RP-1
O/F Ratio: 2.72
Throttle Range: 47% to 100%
Thrust (sl): 859,802 lbf (390,000 kgf) at 311.3 ISP
Thrust (vac): 932,555 lbf (423,000 kgf) at 337.8 ISP
Weight: 12,081 lb (5,520 kg)
T/W (sl): 71.16
T/W (vac): 77.18
Chamber Pressure: 3,721~ psia (25.66 MPa)
Engine Length: 140.945 inches (3.58 meters)
Engine Diameter:
118.11~ inches (3 meters)
Chamber Exit Diameter: 56.29 inches (1.43 meters)
Expansion Area Ratio (ε = Ae/At):
36.87

Notes: Development was begun in 1995 through RD AMROSS, a 50/50 joint venture between NPO Energomash and Pratt & Whitney after Pratt won the race to acquire rights to the RD-170 (Aerojet and Rocketdyne were the other competitors).

Developed as a two-chamber derivative of the 4-chamber RD-170/171 engine; with 70% common hardware to the RD-170 with the remaining 30% hardware being scaled up RD-170 components. Due to strong RD-170 heritage, it took only 42 months to develop and qualify the RD-180. A complete engine takes 18 months to assemble from scratch, with a further two months required for hot fire testing and preparation for final shipment.

In May 2004, detailed RD-180 engineering/manufacturing documentation (225,000+ pages) was delivered to PWR by NPO Energomash, enabling PWR to begin development of an indigenous US manufacturing capability of RD-180, though to date, all engines used so far in the Atlas V program have been built in Russia.

An interesting point to note is that most of NPO Energomash’s structural calculations and analyses were conducted by hand; whereas PWR created over 48 Finite Element Models of various components during analysis of the RD-180.

(RD-180 Pixel Drawing)
(RD-180 Labelled Front/Side Drawing)
(RD-180 Production Schedule)
(RD-180 CAD Render)

References:
International Liquid Rocket Cooperation: The Case of the RD-180 Engine (October 2001) (2.55 MB PDF)
AIAA-2004-3998
RD-180 Engine Production and Flight Experience (602 kb PDF)
AIAA 2006-4361 U.S. Engineering and Operational Capability for Atlas V RD-180

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 Area Ratio (ε = Ae/At):
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)

Miscellaneous Russian

RD-0120

RD-701 Tripropellant

Conceptual Theoretical Engines

75K Nuclear Vapor Thermal Reactor (NVTR)

Propellants: Nuclear/LH2
Thrust (vac): 75,000 lbf at 997.8 ISP
Weight: 6,830 kg total mass (with shield)
T/W Ratio (vac): 4.98
Chamber Pressure: 1,500 PSIA
Chamber Temperature: 5,580 deg R
Expansion Area Ratio (ε = Ae/At): 500

Notes: Uses modified NERVA geometries, with the solid fuel replaced with 85%> enriched Uranium Tetrafluoride (UF4) vapor over Carbon-Carbon Composite fuel elements (C-C).

(Basic Operating Parameters)

Reference:
Conceptual Design of a Vapor Core Reactor Rocket Engine for Space Propulsion (April 1996)

2.8K “Nuclear Lightbulb” Engine with Cold Beryllium Reflectors

Propellant: U233 Plasma
Buffer Vortex: Neon Fluid

# of Cavities: One
Cavity Power: 189 MW(t)
Cavity Power Density: 566 MW(t) / m3
Power Transferred to Propellant: 168 MW(t)
Reflective Liner Heat Load: 21 MW(t)
Beryllium Reflector Heat Load: 7.5 MW(t)
Cavity Length: 1.83 meters
Cavity Diameter: 0.49 meters
Cavity Volume: 0.34 m3
Cavity Pressure: 7,353 PSI (50.7 MPa)

Thrust (Vac): 2,800 lbf (12.5 kN) at 1,835 ISP (1,800 m/s) at 0.68 kg/sec
T/W Ratio (Vac): 0.18
Propellant Exit Temperature: 6700K
Effective Fuel Radiating Temperature: 8300K

Mass, Total: 7,000 kg
        Engine: 4,620 kg
        Moderator: 1,860 kg
        Pressure Vessel: 1,830 kg
        Nozzles and Pumps: 930 kg
        Space Radiator: 2,380 kg

Notes: Fissioning plasma is confined within a transparent (quartz) cell and is kept away from the cell's walls by the swirling flow of a tangentially injected buffer gas (e.g., argon).

Here, energy transfer from the plasma to the propellant is by radiation heat transfer alone, since the two streams cannot physically mix. Even though the quartz cell effectively confines the nuclear fuel, radiation heat transfer is limited to selected regions of the electromagnetic spectrum, established by the transparency of the cell material and its resistance to nuclear radiation-induced "darkening".

This concept has the common engineering problem of promoting effective radiation heat transfer to the flowing propellant, while avoiding thermally induced destruction of the cavity materials. Since neutrons and gamma rays from the fissioning plasma may not be totally absorbed by the propellant stream, nuclear radiation heating of the cavity walls must also be effectively handled.

(Diagram of Concept)

References:
Space Nuclear Power by Joseph A. Angelo, Jr and David Buden.