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Lox/LH2
J-2
J-2
Credit: © Mark Wade
Lox/LH2 propellant.to be used on production space launch vehicles. Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidizer for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was later adopted for the core of the US space shuttle, the European Ariane 5, and the Chinese CZ-5 launch vehicles. It is used in upper stages flown on American, European, Indian, and Chinese boosters. Although extensively developed in Russia, it never reached production for any Russian space launchers.

Specific impulse: 451 s. Specific impulse sea level: 391 s. Location: 2435.

Temperature of Combustion: 2,985 deg K. Ratio of Specific Heats: 1.26. Characteristic velocity c: 2,435 m/s (7,988 ft/sec). Isp Shifting: 391 sec. Isp Frozen: 388 sec. Mol: 10.00 M (32.00 ft).Oxidizer Density: 1.140 g/cc. Oxidizer Freezing Point: -219 deg C.Oxidizer Boiling Point: -183 deg C.Fuel Density: 0.071 g/cc. Fuel Freezing Point: -259 deg C. Fuel Boiling Point: -253 deg C.

Propellant Formulation: LOX/LH2. Optimum Oxidizer to Fuel Ratio: 6. Density: 0.28 g/cc.

Propellant Formulation: LOX/Slush LH2. Optimum Oxidizer to Fuel Ratio: 4. Density: 0.33 g/cc,

Subtopics

5 mlbf Notional LOx/LH2 rocket engine. OOST, ROOST studies 1963.

AEC engine Rocketdyne LOx/LH2 rocket engine. Advanced Expander Cycle Engine. Expander regenerator, pump-fed.

Aerospike Annular Booster Rocketdyne LOx/LH2 rocket engine. Aerospike Typical Annular Booster . Pressure-fed. Thrust from 50,000 to 250,000 lbs at altitude.

AJ23-127 Aerojet LOx/LH2 rocket engine. Booster. The AJ23 was a series of high-performance staged combustion engine designs. None ever made it to production. The -127 featured a gas generator cycle, 56 atm

AJ23-141 Aerojet LOx/LH2 rocket engine. Upper Stage. MIST - Staged Combustion.

AJ23-142 Aerojet LOx/LH2 rocket engine. Booster. ARES - Staged Combustion

AJ23-143 Aerojet LOx/LH2 rocket engine. Booster. Pre-Development, Staged Combustion, 204 atm

AJ23-144 Aerojet LOx/LH2 rocket engine. Booster (Pre-Development, Staged Combustion, 4000 psi). Pre-Development, Staged Combustion, 272 atm

AJ23-145 Aerojet LOx/LH2 rocket engine. Booster (Pre-Development, Staged Combustion, 4000 psi, LOX/RJ-5 or LH2, Single Stage). Pre-Development, Staged Combustion, 272 atm, LOX/RJ-5 or LH2, Single Stage

AJ23-147 Aerojet LOx/LH2 rocket engine. Booster. Gas generator cycle, 55 atm

AJ-60C Aerojet LOx/LH2 rocket engine. Design 2000. Design announced on 3 October 2000 for a new cryogenic upper-stage engine aimed at the very large commercial spacecraft market.

Albatros Carrier Aircraft LOx/LH2 propellant rocket stage. Configuration: delta wing with wingtip vertical stabilizers and canards. Engine type and performance, empty weight estimated.

Albatros Momentum Block LOx/LH2 propellant rocket stage. Unique hydrofoil launch stage for Albatros. Contains 200,000 kg propellants for acceleration by Albatros stage 1 motors to 50 m/s / 180 km/hr launch conditions. Designed by Alekseev Hydrofoil/Ekranoplan OKB.

Albatros Raketoplan LOx/LH2 propellant rocket stage. Configuration: delta wing with wingtip vertical stabilizers.

ALS stage LOx/LH2 propellant rocket stage. .

AMPS-1 Rocketdyne LOx/LH2 rocket engine. Advanced Maneuvering Propulsion System Booster. Pressure-fed.

Angara KVRB LOx/LH2 propellant rocket stage. Planned version for Angara. 5 restarts.

Angara Stage 2 LOx/LH2 propellant rocket stage. Unique configuration with oxidizer in core and fuel in two tanks strapped on in parallel - all of rail-transportable 3.9 m diameter. Built by NPO Energia to Khrunichev design (their own design for Angara and Energia-M were rejected in favor of Khrunichev version). Masses estimated based on engine selected and vehicle performance. Assumed that engine is throttled back to maintain constant 3-G acceleration.

Apollo Direct RM American manned spacecraft module. Study 1961. The retrograde module supplied the velocity increments required during the translunar portion of the mission up to a staging point approximately 1800 m above the lunar surface.

Ares I-2 LOx/LH2 propellant rocket stage. Second stage figures as of summer 2008. Dry mass includes 2500 kg for avionics bay.

Ares Stage 1 LOx/LH2 propellant rocket stage. Core vehicle proposed by NASA for Project Constellation exploration of moon and Mars. It would use shuttle external tank tooling. All masses estimated.

Ares Stage 2 LOx/LH2 propellant rocket stage. Second stage proposed later in design stage by NASA for launch of CEV into low earth orbit. All masses estimated.

Ariane 1-3 H8 LOx/LH2 propellant rocket stage. High energy upper stage for Ariane booster series.

Ariane 2-3 H10 LOx/LH2 propellant rocket stage. .

Ariane 4-3 H10-3 LOx/LH2 propellant rocket stage. .

Ariane 4-3 H10plus LOx/LH2 propellant rocket stage. .

Ariane 5-1 EPC LOx/LH2 propellant rocket stage. 15.2 metric tons increased propellant by moving liquid oxygen bulkhead.

Ariane 5-1 H155 LOx/LH2 propellant rocket stage. Chamber pressure 108 bar; expansion ratio 45.0; propellant mix ratio 5.3.

Ariane 5-2 ESC A LOx/LH2 propellant rocket stage. Uses engine and oxygen tank from Ariane 4 + new liquid hydrogen tank.

Ariane 5-2 ESC B LOx/LH2 propellant rocket stage. New upper stage for Ariane 5. Restartable.

ASE Rocketdyne LOx/LH2 rocket engine. Advanced Space Engine. Staged combustion, pump-fed.

Astro-1 LOx/LH2 propellant rocket stage. Engines 1 x M-1 plus 2 x J-2.

Astro-2 LOx/LH2 propellant rocket stage. Engines 2 x RL-10 plus 1 x J-2.

ATCRE Notional LOx/LH2 rocket engine. Study 1985. Used on Sanger II launch vehicle.

Bono Saucer American manned spaceplane. Study 1963. In 1963 Phil Bono of Douglas Aircraft considered a lenticular configuration for a single-stage-to-orbit reusable booster. This was the largest application found to date for the lenticular concept.

Cargo LV Stage 1 LOx/LH2 propellant rocket stage. Core vehicle proposed by NASA for Project Constellation exploration of moon and Mars. Originally it would use shuttle external tank tooling. This version was proposed by Thiokol prior to Constellation decision. Modification of shuttle external tank. Includes 28.6 metric ton SSME engine pod.

Cargo LV Stage 2 LOx/LH2 propellant rocket stage. Trans-lunar injection stage proposed by NASA for Project Constellation exploration of moon and Mars. It would use shuttle external tank tooling. All masses estimated.

CD Module Notional LOx/LH2 rocket engine. Study 1969. CD Modules - conceptual engines of various thrusts, according to design - were clustered in Martin Marietta Nova designs

Centaur B-X LOx/LH2 propellant rocket stage. Conceptual design, 1998. Not put into production.

Centaur C LOx/LH2 propellant rocket stage. The first high-energy liquid oxygen/liquid hydrogen propellant stage in history. Despite initial development problems, the Centaur is entering its seventh decade of development and production.

Centaur C-X LOx/LH2 propellant rocket stage. Conceptual design - in development 1998, not put into production.

Centaur D/E LOx/LH2 propellant rocket stage. .

Centaur G LOx/LH2 propellant rocket stage. Centaur for Titan 4.

Centaur G Prime LOx/LH2 propellant rocket stage. Centaur for Shuttle payload bay. Cancelled after Challenger disaster on safety grounds.

Centaur G STS LOx/LH2 propellant rocket stage. .

Centaur I LOx/LH2 propellant rocket stage. .

Centaur II LOx/LH2 propellant rocket stage. .

Centaur IIA LOx/LH2 propellant rocket stage. .

Centaur IIIA LOx/LH2 propellant rocket stage. Single-engine Centaur for Atlas IIIA.

Centaur IIIB LOx/LH2 propellant rocket stage. Dual-engine Centaur for Atlas IIIB. The Lockheed Martin manufactured Centaur IIIB upper stage is powered by two Pratt & Whitney RL10A-4-2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. The changes to Centaur for Atlas IIIB are a stretched tank (1.68 m) and the addition of the second engine.

Centaur V1 LOx/LH2 propellant rocket stage. Single-engine Centaur for Atlas V. Centaur is powered by either one or two Pratt & Whitney RL10A-4-2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. For typical, high-energy mission applications, Centaur will be configured with one RL10 engine. Guidance, tank pressurization, and propellant usage controls for both Atlas and Centaur phases are provided by the inertial navigation unit (INU) located on the Centaur forward equipment module.

Centaur V2 LOx/LH2 propellant rocket stage. Dual-engine Centaur for Atlas V. For heavy payload, low earth orbit missions, Centaur will use two RL10 engines to maximize boost phase mission performance.

CEV SAIC American manned spacecraft. Study 2012. SAIC's notional CEV was a Soyuz-shaped aeroshell, enclosing a common pressurized module, and accommodating a crew of four.

CEV Schafer American manned spacecraft. Study 2012. Schafer proposed a lightweight 11 metric ton integral CEV, staged from L1.

Chamber/single nozzle Notional LOx/LH2 rocket engine. Study 1963. Before moving to favored plug nozzle designs, Bono at Douglas considered having multiple combustion chambers exhaust into a single large nozzle to obtained Improved Specific Impulse.

Chang Cheng stage 1 LOx/LH2 propellant rocket stage. All characteristics except dimensions estimated, on assumption that stage used same propulsion systems as Shanghai upper stage.

CLV Stage 2 LOx/LH2 propellant rocket stage. Second stage originally proposed by Thiokol for launch of the CEV into low earth orbit. Also could be used as trans-Mars injection stage on the Cargo LV. Nominal single engine; alternatively 7 RL10-derived engines. All masses estimated.

Cobra Pratt and Whitney LOx/LH2 rocket engine. Design 2003. Proposed as a long-life, moderate-to high-thrust, reusable booster engine that incorporated a safe, low-cost, low-risk, LH2/LOX single burner, using a fuel-rich, staged combustion cycle.

CZ H-18 Chinese space tug. 11 launches, (1994) to (2000). Upper stage / space tug - in production. Launched by CZ-3A, CZ-3B, and CZ-3C.

CZ H-18 stage LOx/LH2 propellant rocket stage. .

CZ-NGLV-500 LOx/LH2 propellant rocket stage. From top to bottom the 5-m Chinese new generation launch vehicle consists of a 117.3 cubic meter liquid oxygen tank, an intertank section, a 350.7 cubic meter liquid hydrogen tank, and an engine section with two gimbaled LOX /LH2 engines of 660 kN vacuum thrust each. The hydrogen tank is pressurized using hydrogen bled from the engine and helium is used to pressurize the oxygen tank.

CZ-NGLV-HO LOx/LH2 propellant rocket stage. The upper stage for the Chinese Next Generation Launch Vehicle is a modification of the CZ-3B upper stage. The stage uses a version of the LOx/LH2 YF-75 engine, simplified for improved reliability. The stage is of hammerhead form, with the upper LH2 tank with a diameter of 5 m, and the lower liquid oxygen tank with a diameter of 3.35 m. The total propellant is 22,900 kg with a burn time of over 600 seconds. Empty mass has not yet been released and is estimated.

CZ-YF-73 LOx/LH2 propellant rocket stage. .

DAC Helios ISI-1 LOx/LH2 propellant rocket stage. .

DAC Helios-1 LOx/LH2 propellant rocket stage. .

Delta 3-2 LOx/LH2 propellant rocket stage. The upgraded cryogenic second-stage Pratt & Whitney RL10B-2 engine is based on the 30-year heritage of the reliable RL10 engine. It incorporates an extendable exit cone for increased specific impulse (Isp) and payload capability.

Delta 4-2 LOx/LH2 propellant rocket stage. Delta 3 second stage with hydrogen tank stretch.

Delta 4H-2 LOx/LH2 propellant rocket stage. Delta 4 second stage with hydrogen tank increased to 5.1 m diameter.

Delta RS-68 LOx/LH2 propellant rocket stage. Low cost expendable stage using lower performance engine. Used in Delta 4, Boeing EELV. Engine can be throttled to 60%.

Energia Core LOx/LH2 propellant rocket stage. .

Energia EUS LOx/LH2 propellant rocket stage. Upper stage designed for use with Energia to boost payloads to geosynchronous, lunar, or planetary trajectories.

Energia M-1 Core stage of Energia-M. LOx/LH2 propellant rocket stage.

GSLV-3 stage LOx/LH2 propellant rocket stage. The stage finally reached hardware status as a joint Russian-Indian development for India's GSLV booster.

H-2-1 LOx/LH2 propellant rocket stage. .

H-2A LRB LOx/LH2 propellant rocket stage. Two-engine version of H-2A-1 used as strap-on booster.

H-2A-1 LOx/LH2 propellant rocket stage. Lower cost version of H-2 first stage. Can be throttled to 72% thrust.

Helios A-1 LOx/LH2 propellant rocket stage. Booster stage with LOx tanks only to take nuclear second stage to stratosphere.

Helios B-1 LOx/LH2 propellant rocket stage. Booster stage with LOx tanks only to take nuclear second stage to stratosphere.

Helios C-1 LOx/LH2 propellant rocket stage. Booster stage with LOx tanks only to take nuclear second stage to stratosphere.

Helios Stage 1 Notional LOx/LH2 rocket engine. Study 1960. Engines for booster stage with LOx tanks only to take nuclear second stage to stratosphere. Helios A, B, C studies.

HG-3 Rocketdyne LOx/LH2 rocket engine. Study 1967. High-performance high-pressure chamber engine developed from J-2. Considered for upgrades to Saturn V launch vehicle upper stages. Technology led to Space Shuttle Main Engines.

HG-3-SL Rocketdyne LOx/LH2 rocket engine. Study 1966. High-performance high-pressure chamber engine developed from the J-2, fitted with lower-expansion nozzle for sea level use on Saturn INT-17. Technology led to Space Shuttle Main Engines.

HIMES LOx/LH2 propellant rocket stage. .

HIMES engine Mitsubishi LOx/LH2 rocket engine. Design 1999. Used on H-2 HIMES launch vehicle.

HM-10 SEP, Ottobrunn LOx/LH2 rocket engine. Developed 1990's. Engine for potential Ariane 5 upper stage.

HM7-A SEP, Ottobrunn LOx/LH2 rocket engine. Development begun 1973. Out of production. Used on Ariane 1 launch vehicle. First flight 1979.

HM7-B SEP, Ottobrunn LOx/LH2 rocket engine. Increased performance version of the HM-7 engine for the Ariane 2 and 3. Combustion chamber pressure raised from 30 to 35 bar and nozzle extended. First flight 1984.

HP-1 Notional LOx/LH2 rocket engine. Study 1963. Operational date would have been December 1974. Used in Martin Nova studies MM 24G, MM 33.

Hyperion Booster LOx/LH2 propellant rocket stage. .

Hyperion SSTO stage LOx/LH2 propellant rocket stage. All values estimated based on drawing, statement that 5 x mass of SASSTO, payload performance, and 300 m/s sled velocity augmentation.

Interim HOTOL stage LOx/LH2 propellant rocket stage. .

IPD Rocketdyne LOx/LH2 rocket engine. Development ended 2006. Integrated Powerhead Demonstrator, end goal was flight-rated, full-flow, hydrogen-fueled, staged combustion rocket engine in the 1.1-million-newton thrust class.

J-2 Rocketdyne LOx/LH2 rocket engine. Used in the Saturn IVB stage in Saturn IB and Saturn V, and the Saturn II stage in the Saturn V. First flight 1966. Sea level versions with reduced expansion ratio were proposed for Saturn II first stage use. Upgraded toroidal aerospike versions (J-2T-200K and J-2T-250K) were developed for upgrades to Saturn upper stages. The modestly improved J-2S was tested and provided basis for X-33 linear aerospike engine thirty years later. After 30 years the J-2 was resurrected again for use in boosting NASA's new Orion manned capsule to orbit. In the event, NASA was unable to resist 'improving' the J-2S, and by early 2007 the engine for the second stage of the Ares 1 Crew Launch Vehicle was the redesignated and substantially different J-2X.

J-2S Rocketdyne LOx/LH2 rocket engine. Developed 1965-1969. J-2 version proposed for Saturn follow-on vehicles, using results of the J-2X technology program. The engine was simplified while offering improved performance.

J-2-SL Rocketdyne LOx/LH2 rocket engine. Study 1966. Sea level version of J-2 with reduced expansion ratio proposed for Saturn II first stage use.

J-2T-200K Rocketdyne LOx/LH2 rocket engine. Study 1965. Proposed for later versions of Saturn V. Toroidal aerospike plug nozzle version of J-2.

J-2T-250K Rocketdyne LOx/LH2 rocket engine. Study 1967. Proposed for later versions of Saturn V. Toroidal aerospike plug nozzle version of J-2.

J-2X Rocketdyne LOx/LH2 rocket engine. Ares I launch vehicle second stage. In development 2006-2016. Began as an update to the J-2 engine of the 1960s, but final design was all-new, 20% more thrust, but double the weight.

Jarvis-2 LOx/LH2 propellant rocket stage. .

KVRB Russian space tug. Study 1992. Upper stage / space tug - design 1992. High energy upper stage for Proton, never put into production.

L-5.00H Notional LOx/LH2 rocket engine. Study 1963. Used on Nova GD-H launch vehicle.

L-5.25H Notional LOx/LH2 rocket engine. Study 1963. Engines for recoverable booster engine package 'half stage' of a 1 1/2 stage arrangement. Used on Nova GD-H launch vehicle.

L6H Notional LOx/LH2 rocket engine. Study 1963. Operational date would have been June 1976. Used in booster stage (engines only). Used on Nova MM 34 launch vehicle.

LE-5 Mitsubishi LOx/LH2 propellant rocket stage. Upper stage on the H-1 launch vehicle. First flight 1986. The LE-5A, a simplified, lower cost version, used an expander bleed cycle turbopump.

LE-5 engine Mitsubishi LOx/LH2 rocket engine. Used on H-1 launch vehicle. First flight 1986.

LE-5A Mitsubishi LOx/LH2 rocket engine. Used on H-2 launch vehicle. First flight 1994.

LE-5B LOx/LH2 propellant rocket stage. Lower cost version of H-2 second stage.

LE-5B engine Mitsubishi LOx/LH2 rocket engine. In production. Improved model of the LE-5A for second stage of the H-II rocket; used hydrogen for the cooling of the thrust chamber, then as the gas to drive the turbine. First flight 2001.

LE-5EC LOx/LH2 propellant rocket stage. .

LE-7 Mitsubishi LOx/LH2 rocket engine for H-2 upper stages. Staged combustion turbopump. No throttle capability. First flight 1994.

LE-7A Mitsubishi LOx/LH2 rocket engine. In production. Improved model of the original LE-7 for the first stage of the H-II rocket with a two stage combustion cycle system. First flight 2001.

LEV American manned lunar lander. Study 1989. The Lunar Excursion Vehicle (LEV) figured in numerous NASA studies of the 1980's and 1990's.

LH2 2000/3000 lbf thrust Aerojet LOx/LH2 rocket engine. Development begun July 1946. Engine R&D, hydrogen production

LH2-80k Notional LOx/LH2 rocket engine. Study 1959. Used on Nova 4L launch vehicle.

LLV American lunar logistics spacecraft. Study 1966. Many versions of new Lunar Logistic Vehicles (LLV's) using several possible candidate propellants were studied by NASA and its contractors in the mid-1960's for post-Apollo lunar base support.

LLV L-I American manned spacecraft module. Study 1966. Lunar Orbit Insertion stage for placing LLV into lunar orbit. Propulsion 2 x RL10-A3 with N2O4/MMH thrusters for orientation, midcourse, and ullage. Lunar orbit insertion of Lunar Logistics Vehicle lander and payload.

LLV L-II American manned spacecraft module. Study 1966. Landing stage for delivery of up to 13,400 kg payload from lunar orbit to lunar surface. Propulsion 2 x RL10-A3 with N2O4/MMH thrusters for orientation, midcourse, and ullage. Delivery of lunar base elements from lunar orbit to lunar surface.

LM Langley Lighter American manned lunar lander. Study 1961. This early open-cab Langley design used cryogenic propellants. The cryogenic design was estimated to gross 3,284 kg - to be compared with the 15,000 kg / 2 man LM design that eventually was selected.

LM Langley Lightest American manned lunar lander. Study 1961. Extremely light-weight open-cab lunar module design considered in early Langley studies.

LR129 Pratt and Whitney LOx/LH2 rocket engine. Engine developed for boost/glide aerospace craft; later modified into unsuccessful competitor for Space Shuttle main engine.

LR87 LH2 Aerojet LOx/LH2 rocket engine. Development ended 1961. Version of the Titan engine, and first large LOx/LH2 engine fired in the world. 52 static tests. But NASA selected Rocketdyne instead to develop the J-2 engine for Saturn from scratch.

M-1 Aerojet LOx/LH2 rocket engine. Study 1961. Engine developed 1962-1966 for Uprated Saturn and Nova million-pound payload boosters to support manned Mars missions. Reached component test stage before cancellation.

Magnum Core LOx/LH2 propellant rocket stage. Alternative configurations used 2 to 3 RS-68 engines.

Mars Oz Australian manned Mars expedition. Study 2006. 2001 design study by the Mars Society Australia that incorporated many innovative elements to produce a minimum-mass non-nuclear Mars expedition concept.

MB-35 Rocketdyne LOx/LH2 rocket engine. Design 2004. Mitsubishi / Boeing joint project for an engine for Delta IV cryogenic upper stages. Expander bleed, pump-fed.

MB-45 Rocketdyne LOx/LH2 rocket engine. Design 2004. Mitsubishi / Boeing joint project for an engine for Delta IV cryogenic upper stages, announced February 2000.

MB-60 Rocketdyne LOx/LH2 rocket engine. Design 2004. Mitsubishi / Boeing joint project for an engine for Delta IV cryogenic upper stages. Expander bleed, pump-fed.

MBB-ATC500 MBB LOx/LH2 rocket engine. Study 1969. Used on Beta launch vehicle.

McDonnell-Douglas ILRV Drop Tanks LOx/LH2 propellant rocket drop tank for ILRV design.

McDonnell-Douglas ILRV stage LOx/LH2 propellant rocket stage. High-pressure Lox/Lh2 engine.

MLLV Core LOx/LH2 propellant rocket stage. Boeing study, 1969.

Mustard The British Aircraft Corporation "Multi-Unit Space Transport And Recovery Device" design of 1964-1965 was a winged two-stage-to-orbit reusable space shuttle using the 'triamese' concept. The three components of the design were lifting bodies with a configuration similar to the American HL-10 vehicle. BAC sought to reduce development cost by use of two boosters nearly identical to the orbiter vehicle.

Mustard 1 LOx/LH2 propellant rocket stage. .

Mustard 2 LOx/LH2 propellant rocket stage. .

N1 Block R LOx/LH2 propellant rocket stage. Designed 1965-1971 as replacement for N-1 Blok D. Cancelled 1971 in favor of Blok Sr; revived and developed in 1974-1976. First static test Oct 12 1976. Two stages tested 1976-1977. Strangely never replaced Blok D on Proton.

N1 Block S LOx/LH2 propellant rocket stage. Designed 1965-1971 as replacement for N-1 Blok G. Cancelled in 1971 in favor of development of single stage, Block Sr.

N1 Block Sr LOx/LH2 propellant rocket stage. Upper stage developed 1971-1974 to support manned lunar expedition. Replaced Blok R/Blok S previously under development. Capable of five restarts and 11 days of flight. Could insert 24 metric tons into lunar orbit or 20 metric tons into geosynchronous orbit.

N1 Block V-II LOx/LH2 propellant rocket stage. N1 improvement study, 1965. LOx/LH2 replacement for Block B second stage.

N1 Block V-III LOx/LH2 propellant rocket stage. N1 improvement study, 1965. LOx/LH2 replacement for Block V third stage. Pursued into 1966 and later, but later efforts concentrated on Block S, R, and SR cryogenic stages.

NASA Mars Expedition 1971 American manned Mars expedition. Study 1971. Final NASA Mars expedition before the 1980's. The spacecraft would use shuttle hardware, including SSME engines in the rocket stages.

NASA Mars Flyby 1965 American manned Mars flyby. Study 1965. Mars flyby mission designed by NASA Huntsville in 1965 to use existing Apollo hardware, allowing a manned flyby of Mars by 1975.

NK-15VM Kuznetsov LOx/LH2 rocket engine. N-1 stage 2 (block B) replacement. Design 1972. Derivative of NK-15 with kerosene replaced by hydrogen. Canceled before hot-tests.

NK-35 Kuznetsov LOx/LH2 rocket engine. Design 1972. Derivative of the NK-15 with kerosene replaced by hydrogen. The engine was canceled before hot-tests. Proposed for the UR-700M Mars booster in 1972, but this was not approved either.

NLS Core LOx/LH2 propellant rocket stage. .

NLS HLV stage LOx/LH2 propellant rocket stage. .

NLS Semistage LOx/LH2 propellant rocket stage. .

Nova 4 J-2 LOx/LH2 propellant rocket stage. Nova third stage.

Nova 59-4-3 LOx/LH2 propellant rocket stage. Empty Mass Estimated.

Nova 59-4-4 LOx/LH2 propellant rocket stage. Empty Mass Estimated.

Nova 60-8-2 LOx/LH2 propellant rocket stage. Mass estimated based on total LV weight. J-2-powered version of this stage also proposed.

Nova 60-8-3 LOx/LH2 propellant rocket stage. Mass estimated based on total LV weight.

Nova 9L-3 LOx/LH2 propellant rocket stage. Masses estimated based on total vehicle thrust, performance, and stage volumes.

Nova 9L-4 LOx/LH2 propellant rocket stage. Masses estimated based on total vehicle thrust, performance, and stage volumes.

Nova A-2 LOx/LH2 propellant rocket stage. .

Nova A-3 LOx/LH2 propellant rocket stage. .

Nova B-2 LOx/LH2 propellant rocket stage. .

Nova B-3 LOx/LH2 propellant rocket stage. .

Nova DAC 2-1 LOx/LH2 propellant rocket stage. Operational date would have been July 1977. Recoverable stage. 10% plug nozzle.

Nova DAC 2-2 LOx/LH2 propellant rocket stage. Operational date would have been July 1977. Recoverable stage.

Nova DAC ISI-1 LOx/LH2 propellant rocket stage. Operational date would have been July 1977. Recoverable stage. 10% plug nozzle.

Nova DAC ISI-2 LOx/LH2 propellant rocket stage. Operational date would have been July 1977. Recoverable stage.

Nova DAC-2 LOx/LH2 propellant rocket stage. Operational date would have been July 1977. Recoverable stage.

Nova GD-B-2 LOx/LH2 propellant rocket stage. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions.

Nova GD-E-2 LOx/LH2 propellant rocket stage. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions.

Nova GD-F-2 LOx/LH2 propellant rocket stage. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions.

Nova GD-H-0 LOx/LH2 propellant rocket stage. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. Recoverable booster engine package 'half stage' of a 1 1/2 stage arrangement. Separation at 2,980 m/s at 87,800 m altitude; splashdown under 4 46 m diameter parachutes 1,000 km downrange.

Nova GD-H-1 LOx/LH2 propellant rocket stage. .

Nova GD-J-2 LOx/LH2 propellant rocket stage. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions.

Nova MM 14A-2 LOx/LH2 propellant rocket stage. Operational date would have been April 1973.

Nova MM 14B-2 LOx/LH2 propellant rocket stage. Operational date would have been February 1973.

Nova MM 1B-2 LOx/LH2 propellant rocket stage. Operational date would have been February 1973.

Nova MM 1C-2 LOx/LH2 propellant rocket stage. Operational date would have been February 1973.

Nova MM 24G-2 LOx/LH2 propellant rocket stage. Operational date would have been December 1974.

Nova MM 33-1 LOx/LH2 propellant rocket stage. Operational date would have been April 1975. SSTO - payload 1,042,000 lbs.

Nova MM 34-0 LOx/LH2 propellant rocket stage. Operational date would have been June 1976. Booster stage (engines only).

Nova MM 34-1 LOx/LH2 propellant rocket stage. Operational date would have been June 1976. Sustainer stage (required 4-engine booster stage).

Nova MM S10E-1 stage LOx/LH2 propellant rocket stage. Operational date would have been October 1977. SSTO; expendable.

Nova MM S10E-2 stage LOx/LH2 propellant rocket stage. Operational date would have been November 1977. SSTO; expendable; payload 1,283,000 lbs.

Nova MM S10R-1 stage LOx/LH2 propellant rocket stage. Operational date would have been June 1978. SSTO; recoverable.

Nova MM S10R-2 stage LOx/LH2 propellant rocket stage. Operational date would have been July 1978. SSTO; recoverable; payload 842,000 lbs.

Nova MM T10EE-1-1 LOx/LH2 propellant rocket stage. Operational date would have been November 1976. Expendable stage.

Nova MM T10EE-1-2 LOx/LH2 propellant rocket stage. Operational date would have been November 1976. Expendable stage.

Nova MM T10RE-1-1 LOx/LH2 propellant rocket stage. Operational date would have been January 1977. Recoverable stage.

Nova MM T10RR-2-1 LOx/LH2 propellant rocket stage. Operational date would have been December 1976. Recoverable stage.

Nova MM T10RR-2-2 LOx/LH2 propellant rocket stage. Operational date would have been December 1976. Recoverable stage.

Nova MM T10RR-3-1 LOx/LH2 propellant rocket stage. Operational date would have been July 1977. Recoverable stage. 10% plug nozzle.

Nova MM T10RR-3-2 LOx/LH2 propellant rocket stage. Operational date would have been July 1977. Recoverable stage.

Nova MM-1 LOx/LH2 propellant rocket stage. Operational date would have been December 1974.

Nova NASA-3 LOx/LH2 propellant rocket stage. .

OOST ISI stage LOx/LH2 propellant rocket stage. .

OOST stage Bono's earliest design for an expendable single-stage-to-orbit LH2/LOx booster. The baseline version used conventional engines.

Ottobrunn 300N Ottobrunn LOx/LH2 rocket engine. 300 N. Upper stages. Developed 2000.- highest value ever achieved in Europe for an engine of such small size.

OTV American space tug. Studied 1985-1989. The Orbital Transfer Vehicle was reusable space tug, powered by LOx/LH2 engines and equipped with an aerobrake allowing it to be returned for refueling and reuse at an orbiting space station.

P320 Rocketdyne, Friedrichshafen LOx/LH2 rocket engine. BORD 1/P320 BOELKOW (Germany)/Rocketdyne Technology. Pressure-fed.

Pegasus Tanks xLOx/LH2 propellant rocket drop tank. . Four tanks jettisoned at 130 seconds after liftoff; two at 250 seconds, last two at orbital insertion, 360 seconds after liftoff.

Pegasus VTOVL stage LOx/LH2 propellant rocket stage. Empty mass includes 29,600 kg of propellants used for re-entry cooling of plug nozzle and rocket soft landing at landing field.

Plug-Nozzle J-2 Rocketdyne LOx/LH2 rocket engine. Study 1993. Plug nozzle version of J-2 proposed for certain Saturn V upgrades in late 1960's. Used on DC-I launch vehicle.

Plug-Nozzle Pegasus Notional LOx/LH2 rocket engine. Study 1966. Used on Pegasus VTOVL launch vehicle.

Plug-Nozzle Rombus Notional LOx/LH2 rocket engine. Study 1964. Used on Rombus launch vehicle.

Plug-Nozzle SASSTO Notional LOx/LH2 rocket engine. Study 1967. Used on SASSTO launch vehicle.

Plug-Nozzle SERV Notional LOx/LH2 rocket engine. Study 1971. Used on Shuttle SERV launch vehicle.

Plug-Nozzle SSME Notional LOx/LH2 rocket engine. Study 1978. Used on VTOVL launch vehicle.

Project 921 LV-2 LOx/LH2 propellant rocket stage. Additionally 4 vernier LOx/LH2 engines with a total thrust of 4600 kgf and a storable engine package for stage propellant ullage and restart.

Proton KM-4 LOx/LH2 propellant rocket stage. Planned version for Proton. Never developed.

PW 1000000 lb LH2 Pratt and Whitney LOx/LH2 rocket engine. Study 1988. Part of launch vehicle proposed by Martin as alternative to NLS. All figures estimated based on 1,000,000 lb thrust single engine.

RBCC Rocketdyne LOx/LH2 rocket engine. Isp>400s. Rocket Based Combined-Cycle A5 Development Engine; integrated rocket, air-augmented rocket, ramjet, and scramjet propulsion elements into a single flow path.

RD-0120 Kosberg engine used in the Energia core stage. In 1987 it became the first operational Russian LOx/LH2 engine system, built to the same overall performance specifications as America's SSME, but using Russian technology.

RD-0120M Kosberg LOx/LH2 rocket engine. Energia-M core stage. Development ended 1993. From 1987 KBKhA worked on upgrading the 11D122 (RD-0120) engine for Energia-M launcher, including the possibility to throttle the engine down to 28% thrust.

RD-0122 Kosberg LOx/LH2 rocket engine. Energia-M core stage. Planned for Angara central stage. Developed 1990-. Upgrade of RD-0120 engine for Energia-M launcher with increased thrust. Prototype from RD-0120 hardware.

RD-0126 Kosberg LOx/LH2 rocket engine. Space tugs or upper stage for Onega or Yastreb versions of Soyuz. Single annular chamber with expansion-deflection nozzle, separate turbopumps. Design concept 1993. Hot-tests in 1998.

RD-0126A Kosberg LOx/LH2 rocket engine. Upper stages. Design concept 1996-. Concept for a cryogenic engine for upper stages. Single annular chamber with expansion-deflection nozzle, common turbopump.

RD-0126E Kosberg LOx/LH2 rocket engine. Upper stages. Design concept 1998-. Concept for a cryogenic engine for upper stages. Single annular chamber with straight expansion nozzle, common turbopump.

RD-0128 Kosberg LOx/LH2 rocket engine. upper stage. Design concept 1996-. Concept for a cryogenic engine for upper stages. One single chamber with bell nozzle, separate turbopumps.

RD-0131 Kosberg LOx/LH2 rocket engine. upper stage. Design concept 1996-. Concept for a cryogenic engine for upper stages. Single annular chamber with expansion-deflection nozzle, common turbopump.

RD-0132 Kosberg LOx/LH2 rocket engine. upper stage. Design concept 1996-. Concept for a cryogenic engine for upper stages. Derived from RD-0131, but four chambers with bell nozzles, common turbopump.

RD-0133 Kosberg LOx/LH2 rocket engine. upper stage. Design concept 1996-. Concept for a cryogenic engine for upper stages. Four chambers with bell nozzles, common turbopump.

RD-0146 Kosberg LOx/LH2 rocket engine. Centaur upper stage (Atlas); high performance upper stages for Onega, Proton, Angara launch vehicles. Design concept 1998-.

RD-135 Glushko LOx/LH2 rocket engine. upper stage. Developed -1976. Experimental cryogenic engine. (Ref. May be not correct.)

RD-54 Lyulka LOx/LH2 rocket engine. N1 concept stage III. Developed 1960-75.

RD-56 Isayev LOx/LH2 rocket engine. N1 block R. Development ended 1971. Oxygen-hydrogen engine for cryogenic upper stage. Developed but never flown. Design sold to India in 1990's for GSLV.

RD-56M Isayev LOx/LH2 rocket engine. Originally developed for Proton and Angara upper stage KVRB with planned first flight 1995. Finally flown on 12KRB upper stage for India's GSLV. First flight 2001.

RD-57 Lyulka LOx/LH2 rocket engine. N1 Block S (N-1M). Study 1965. One to have been used in N1 Block S. In fixed chamber version, 3 to 6 to have been used in N1 Block V-III. Engine system includes roll control thruster with 1.29 kN thrust.

RD-57A-1 Lyulka LOx/LH2 rocket engine. Developed 1995-98. New version of RD-57M for SSTO-demonstrator proposed by Aerojet. Optimized nozzle contour for performance increase, new chamber material for weight reduction.

RD-57M Lyulka LOx/LH2 rocket engine. Vulkan Blok V. Development ended 1976. Version with extendible nozzle. Length 4.06 / 2.61 m. Specific impulse 461 / 448 sec. Area ratio 170 / 87.6.

RL-10 Pratt and Whitney LOx/LH2 rocket engine family. First flight 1961. Originally planned for use in Centaur upper stage for Atlas, but earliest successful flights in Saturn IV stage for Saturn I. Throttleable version designed for direct-landing Apollo mission, but cancelled. Sea-level version used in DC-X SSTO test vehicle. Numerous developed versions used in Atlas, Atlas V, Delta IV upper stages. Only production American upper-stage LOx/LH2 engine. Specifications are for early version as proposed for Nova A, Nova B, Saturn B-1, Saturn C-2, Saturn C-3, Saturn I. First flight 1961.

RL-10A-1 Pratt and Whitney LOx/LH2 rocket engine. Version used on Atlas Centaur LV-3C, and proposed for various early Saturn launch vehicle designs. First flight 1961.

RL-10A-3 Pratt and Whitney LOx/LH2 rocket engine. First flight 1967.

RL-10A-3A Pratt and Whitney LOx/LH2 rocket engine. Used on Centaur stage atop Atlas G, Atlas I, Atlas II, Titan 4. First flight 1984.

RL-10A-4 Pratt and Whitney LOx/LH2 rocket engine. Out of production. Centaur stage for Atlas IIA, Atlas IIAS. First flight 1992.

RL-10A-4-1 Pratt and Whitney LOx/LH2 rocket engine. Out of production. Used on Atlas IIIA launch vehicle. First flight 2000. Version with one of engines removed; remaining engine re-positioned to center-mount; new electro-mechanical gimbals.

RL-10A-4-2 Pratt and Whitney LOx/LH2 rocket engine. In production. Used on Atlas IIIB launch vehicle. First flight 2002. Two engines; electro-mechanical thrust vector control actuators replaced earlier hydraulically actuated system.

RL-10A-5 Pratt and Whitney LOx/LH2 rocket engine. Throttleable to 30% of thrust, sea level version of RL10. Four engines were built and were used on the DC-X and the upgraded DC-XA VTOVL SSTO launch vehicle demonstrators. First flight 1993.

RL-10A-5KA Pratt and Whitney LOx/LH2 rocket engine. Kistler proposal. Design 1992. Throttleable to 30% of thrust, sea level version of RL10 with extendable nozzle for high altitude operation.

RL-10B-2 Pratt and Whitney LOx/LH2 rocket engine. In production. Used on Delta 3 , Delta IV launch vehicles. First flight 1998. Extendable exit cone for increased specific impulse; electromechanical actuators replace hydraulic systems.

RL-10B-X Pratt and Whitney LOx/LH2 rocket engine. Design concept 1994.

RL-10C Pratt and Whitney LOx/LH2 rocket engine. In Production. Used in Delta 3 - 2. First flight 1998.

RL-10C-X Pratt and Whitney LOx/LH2 rocket engine. Design concept 1994.

RL-50 Pratt and Whitney LOx/LH2 rocket engine. Development. Advanced, high-performance upper-stage rocket engine proposed by Pratt & Whitney for both domestic and international launch vehicles.

RL-60 Pratt and Whitney LOx/LH2 rocket engine. Design. Upper stage engine to have been developed by Pratt and Whitney with several international partners. Same dimensions as the RL-10, but over twice the thrust.

RM-1500H Rocketdyne LOx/LH2 rocket engine. Space Shuttle Orbiter Auxiliary Propulsion. Pressure-fed.

Rombus core LOx/LH2 propellant rocket stage. 36 x plug-nozzle engines (20 atm chamber pressure, 7:1 mixture ratio).

Rombus Tank LOx/LH2 propellant rocket drop tank. . Eight of these liquid hydrogen tanks would be mounted around the core of Rombus and stage in pairs at 130 seconds, 196 seconds, and 300 seconds after launch.

ROOST ISI stage LOx/LH2 propellant rocket stage. .

ROOST stage LOx/LH2 propellant rocket stage. .

RS-2100 Rocketdyne LOx/LH2 rocket engine. Next Generation Launch Vehicle Booster. Full flow staged combustion, pump-fed. Thrust and specific impulse values are at sea level.

RS-2200 Rocketdyne LOx/LH2 rocket engine. Development cancelled 1999. Linear Aerospike Engine developed for use on the Lockheed Reusable Launch Vehicle, the production follow-on to the X-33.

RS-52 Rocketdyne LOx/LH2 rocket engine. Oxygen/Hydrogen Space Station Thruster. Pressure-fed. Technology was developed with 0.1 lb thrust resistojet by using electrically heated waste for space station propulsion.

RS-68 Rocketdyne LOx/LH2 rocket engine. In production. First new large liquid-fueled rocket engine developed in America in more than 25 years. Powered the Delta IV booster. First flight 2002.

RS-68 Regen Rocketdyne LOx/LH2 rocket engine. Design concept -2004. Upgrade to basic RS-68 for Delta IV Heavy growth versions. Would use a regeneratively-cooled expansion nozzle, allowing it to run hotter, with higher thrust and specific impulse.

RS-68B Rocketdyne LOx/LH2 rocket engine. Design concept -2004. Upgrade (details not specified) to basic RS-68 for Delta IV Heavy growth versions.

RS-71 Rocketdyne LOx/LH2 rocket engine. Development ended 1999. Linear Aerospike SR-71 Experiment. Pressure-fed.

RS-74 Rocketdyne LOx/LH2 rocket engine. Next Generation Launch Vehicle Booster. Full flow staged combustion, pump-fed. Thrust and specific impulse values are at sea level.

RS-800 Rocketdyne LOx/LH2 rocket engine. Design concept -2004. New high-thrust cryogenic engine for Delta IV Heavy growth versions.

RS-XXX Rocketdyne LOx/LH2 rocket engine. Design concept -2004. New high-thrust cryogenic engine concept for Next Generation Delta with 7 m diameter modules.

Sanger I-1 LOx/LH2 propellant rocket stage. The first stage took the second stage to 50 km altitude and 4000 m/s separation conditions. The first stage would then land 500 km from the launch point. Takeoff speed was 300 m/s; and landing speed 80 m/s. A two-man crew piloted the booster.

Sanger I-2 LOx/LH2 propellant rocket stage. The second stage would reach a 300 km earth orbit and a top speed of 8000 m/s. The glider had a landing speed of 90 m/s. Aside from the two-man crew, a five metric ton payload could be delivered into orbit.

Sanger II-2 LOx/LH2 propellant rocket stage. 6000 kg to LEO.

SASSTO stage LOx/LH2 propellant rocket stage. Recoverable S-IVB with plug nozzle engine. 36 x plug-nozzle engines (102 atm chamber pressure, 6:1 mixture ratio).

Saturn II LOx/LH2 propellant rocket stage. Configuration as flown.

Saturn II C-5A LOx/LH2 propellant rocket stage. Final common second stage design for Saturn C-3, C-4 and C-5 (November 1961). Developed into Saturn V second stage.

Saturn II-INT-17 LOx/LH2 propellant rocket stage. Saturn II modified with reduced expansion ratio HG-3 high pressure engines for use a first stage (sea level launch).

Saturn II-SL LOx/LH2 propellant rocket stage. Saturn II modified with reduced expansion ratio J-2 engines for use a first stage (sea level launch). Requires solid rocket motor augmentation to get off the ground.

Saturn IV LOx/LH2 propellant rocket stage. Configuration as flown.

Saturn IVB LOx/LH2 propellant rocket stage. Configuration as flown on Saturn V.

Saturn IVB (S-IB) LOx/LH2 propellant rocket stage. Saturn IB version of S-IVB stage. Due to lower payload, 300 kg saving in structure compared to Saturn V version. Due to deletion of restart requirement, 700 kg saving in propulsion system (primarily reduction in helium for restart).

Saturn IVB (S-V) LOx/LH2 propellant rocket stage. Saturn V version of S-IVB stage for use with upper stage.

Saturn IVB C-3B LOx/LH2 propellant rocket stage. Final common third stage design for Saturn C-3B (November 1961).

Saturn IVB C-5A LOx/LH2 propellant rocket stage. Final common third stage design for Saturn C-4 and C-5 (November 1961). Developed into Saturn V second stage. After development started, decision taken to boost performance by increasing diameter to 6.61 m and increasing propellant load.

Saturn IVB-A LOx/LH2 propellant rocket stage. Douglas Studies, 1965: S-IVB with 215k lbf J-1 (actual final model had 230k J-1).

Saturn MS-II-1 LOx/LH2 propellant rocket stage. Basic Saturn II with 41 inch stretch of hydrogen tank.

Saturn MS-II-1A LOx/LH2 propellant rocket stage. Basic Saturn II with 187 inch stretch of propellant tanks, 1.2 million pound propellant capacity, and 7 J-2 engines.

Saturn MS-II-1-J-2T-200K LOx/LH2 propellant rocket stage. Basic Saturn II with 41 inch stretch of hydrogen tank, uprated J-2T 200k engines with 10 second increase in specific impulse.

Saturn MS-II-1-J-2T-250K LOx/LH2 propellant rocket stage. Basic Saturn II with 41 inch stretch of hydrogen tank, uprated J-2T 250k engines with 25% improvement in thrust and 16 second increase in specific impulse.

Saturn MS-II-2 LOx/LH2 propellant rocket stage. Basic Saturn II with 187 inch stretch of propellant tanks and high chamber pressure SSME-type engines with 65% increase in thrust and 26 second improvement in specific impulse.

Saturn MS-II-3B LOx/LH2 propellant rocket stage. S-II with 15.5 foot stretch, 1.29 million pounds propellant, 7 x 400,000 lb thrust toroidal engines.

Saturn MS-II-4(S)B LOx/LH2 propellant rocket stage. Standard S-II but with structural strength increased from 86% to 502% depending on station, resulting in 8.6% increase in empty weight.

Saturn MS-IVB-1 LOx/LH2 propellant rocket stage. Marshall studies, 1965: S-IVB structurally strengthened to handle larger payloads, otherwise unchanged.

Saturn MS-IVB-1A LOx/LH2 propellant rocket stage. S-IVB with 16.5 foot stretch, 350,000 pounds propellant, standard J-2 engine.

Saturn MS-IVB-2 LOx/LH2 propellant rocket stage. Douglas Studies, 1965: S-IVB with 315 k high pressure 3000 psia engine, 350,000 pounds propellant.

Saturn MS-IVB-3B LOx/LH2 propellant rocket stage. S-IVB with 16.5 foot stretch, 350,000 pounds propellant, 1 400,000 pound thrust toroidal engine.

Saturn MS-IVB-4(S)B LOx/LH2 propellant rocket stage. Standard S-IVB but with structural strength increased from 78% to 217% depending on station, resulting in 11.8% increase in empty weight.

Saturn MS-IVB-x LOx/LH2 propellant rocket stage. Marshall studies, 1965: S-IVB structurally strengthened to handle larger payloads, otherwise unchanged.

Saturn S-II LOx/LH2 propellant rocket stage. Early design version for use with Saturn I first stage.

Saturn S-II-4 LOx/LH2 propellant rocket stage. Version for Saturn C-4.

Saturn S-II-8 LOx/LH2 propellant rocket stage. Version for Saturn C-8.

Saturn S-II-C3 LOx/LH2 propellant rocket stage. Version for Saturn C-3.

Saturn S-IVC LOx/LH2 propellant rocket stage. .

Sea Dragon-2 LOx/LH2 propellant rocket stage. Length with extendible nozzle deployed 87 m. Diameter of extended nozzle 30 m. Total mass, specific impulse estimated from booster performance figures.

Sea Dragon-2 engine Aerojet LOx/LH2 rocket engine. Design, 1962. Truax pressure fed design. Diameter of extended nozzle 30 m. Specific impulse estimated from booster performance figures.

Sea Horse-2 LOx/LH2 propellant rocket stage. .

Shuttle DC-3-1 LOx/LH2 propellant rocket stage. Faget Configuration.

Shuttle DC-3-2 LOx/LH2 propellant rocket stage. Faget Configuration - Cross Range 323 km.

Shuttle FR-3-1 LOx/LH2 propellant rocket stage. Trapezoidal lifting body configuration.

Shuttle FR-3-2 LOx/LH2 propellant rocket stage. Trapezoidal lifting body configuration. Cross range 2419 km.

Shuttle H33-1 LOx/LH2 propellant rocket stage. Swept wing configuration.

Shuttle H33-2 LOx/LH2 propellant rocket stage. Delta wing configuration with drop tanks - Cross Range 1,774 km.

Shuttle HCR-1 LOx/LH2 propellant rocket stage. Swept wing configuration.

Shuttle HCR-2 LOx/LH2 propellant rocket stage. Delta winged configuration.

Shuttle LCR-1 LOx/LH2 propellant rocket stage. Swept winged configuration.

Shuttle LCR-2 LOx/LH2 propellant rocket stage. Faget Straight Wing Configuration.

Shuttle LRB stage LOx/LH2 propellant rocket stage. .

Shuttle LS A-1 LOx/LH2 propellant rocket stage. Delta winged configuration.

Shuttle LS A-2 LOx/LH2 propellant rocket stage. High-fineness lifting-body configuration. Cross Range 2,419 km.

Shuttle LS200-1 LOx/LH2 propellant rocket stage. High-fineness lifting-body configuration. Cross Range 2,419 km.

Shuttle MDC-A-1 LOx/LH2 propellant rocket stage. Delta winged configuration.

Shuttle MDC-A-2 LOx/LH2 propellant rocket stage. HL-10 lifting body configuration.

Shuttle NAR A-1 LOx/LH2 propellant rocket stage. Faget Straight Wing Configuration.

Shuttle NAR A-2 LOx/LH2 propellant rocket stage. Faget Straight Wing Configuration - Cross Range 2,000 km.

Shuttle Orbiter LOx/LH2 propellant rocket stage. .

Shuttle R134C-1 LOx/LH2 propellant rocket stage. Delta winged configuration.

Shuttle R134C-2 LOx/LH2 propellant rocket stage. Delta winged configuration.

Shuttle R134G-1 LOx/LH2 propellant rocket stage. Delta winged configuration.

Shuttle R134G-2 LOx/LH2 propellant rocket stage. Straight winged configuration.

Shuttle SERV-1 LOx/LH2 propellant rocket stage. Single stage to orbit, ballistic reentry.

Shuttle Super Lightweight Tank LOx/LH2 propellant rocket stage. The Super Lightweight Tank used 2195 Aluminum-Lithium alloy as the main structural material in place of the 2219 aluminum alloy of the original design. This saved 3,500 kg in empty mass, increasing shuttle payload by the same amount. This change was made in anticipation of Shuttle-Mir and Shuttle-ISS flights to high inclination 51.6 degree orbits. The tank was first used on STS-91 in June 1998.

Shuttle Tank LOx/LH2 propellant rocket drop tank. . Original version.

SLS Stage A LOx/LH2 propellant rocket stage. Smallest LOx/LH2 stage planned for SLS series. Empty mass estimated. Sized for rail transport within USA.

SLS Stage B LOx/LH2 propellant rocket stage. Translunar injection stage for Project Lunex. Masses estimated based on optimum apportioning of B+C stage total masses.

SLS Stage C LOx/LH2 propellant rocket stage. Launch vehicle core stage for Project Lunex. Masses estimated based on optimum apportioning of B+C stage total masses. Thrust, engines estimated based on requirements.

Space Tug The original Boeing Space Tug design of the early 1970's was sized to be flown either in a single shuttle mission or as a Saturn V payload. Optimum mass was found to be 20.6 metric tons regardless.

Spacemaster-1 LOx/LH2 propellant rocket stage. Unique Catamaran configuration.

Spacemaster-2 LOx/LH2 propellant rocket stage. Delta Winged, Cross Range 2,742 km.

SPW-2000 SNECMA, Pratt and Whitney LOx/LH2 rocket engine. Design 2000. New upper-stage cryogenic engine for the upgraded Ariane-5, the Atlas-5, and other new vehicles.

SSME Rocketdyne LOx/LH2 rocket engine. In production. Space Shuttle Main Engines; only high-pressure closed-cycle reusable cryogenic rocket engine ever flown. . Three mounted in the base of the American space shuttle. First flight 1981.

SSME Demonstrator Booster Rocketdyne LOx/LH2 rocket engine. Pressure-fed.

SSME Plus Notional LOx/LH2 rocket engine. VTOHL studies, 1978.

SSME Study Notional LOx/LH2 rocket engine. Study 1967.

Starlifter LOx/LH2 propellant rocket stage. .

Starlifter Tank LOx/LH2 rocket stage. Mass 257,778 kg (568,303 lb).

STCAEM Cryogenic Aerobrake American manned Mars expedition. Study 1991. The STCAEM cryogenic / aerobrake (CAB) concept was used as the NASA reference vehicle.

STME Rocketdyne LOx/LH2 rocket engine. Cancelled 1984. Space Transportation Main Engine. Rocketdyne was teamed with Aerojet and Pratt & Whitney on the STME, which was to have powered the next generation of large launch vehicles.

Titan 5 stage LOx/LH2 propellant rocket stage. Part of launch vehicle proposed by Martin as alternative to NLS. All figures estimated based on 1,000,000 lb thrust single engine.

Titan C-2 LOx/LH2 propellant rocket stage. Engine developed 1958-1960, but launch vehicle cancelled 1961.

Toroid FD Notional LOx/LH2 rocket engine. Study 1963. Operational date would have been December 1976. Engines for recoverable stage. Used on Nova MM T10RR-2 launch vehicle.

Toroidal 400k Notional LOx/LH2 rocket engine. Study 1967. Used on Saturn V-3B launch vehicle.

Toroidal 560k Notional LOx/LH2 rocket engine. Design concept 1990's.

TR-106 TRW LOx/LH2 rocket engine. Development. Innovative TRW 650K Low Cost Pintle Engine, test fired at NASA's test center in October 2000.

Truax LH2 Aerojet LOx/LH2 rocket engine. Test 1962. Used in Sea Horse-2.

TRW Mars American manned Mars expedition. Study 1963. In 1963 TRW designed a Mars expedition using aerobraking at both Mars and Earth, and a swing-by of Venus on return.

UR-700M-3 LOx/LH2 propellant rocket stage. Total mass, length, estimated based on empty mass, total vehicle mass. Engine specific impulse estimated based on performance requirements.

Venturestar American SSTO winged orbital launch vehicle. Production reusable single-stage-to-orbit launch vehicle using technology developed in X-33 test bed.

Vinci Snecma, Ottobrunn LOx/LH2 rocket engine. Upper Stages. In development. Advanced expander cycle cryogenic propellant rocket engine with the capability of five in-space restarts. First hot-fire tests 2005. First flight 2010.

VTOHL 45t American SSTO winged orbital launch vehicle. Vertical Takeoff Horizontal Landing (winged).

VTOHL 9t American SSTO winged orbital launch vehicle. Vertical Takeoff Horizontal Landing (winged).

VTOVL 150t LOx/LH2 propellant rocket stage. .

Vulcain SEP, Ottobrunn LOx/LH2 rocket engine. In production. Powered the cryogenic core stage of Ariane 5. First flight 1996. Upgraded versions developed and proposed for later Ariane 5 versions.

Vulcain 2 SEP, Ottobrunn LOx/LH2 rocket engine. In development. New generator cycle rocket engine for an Ariane 5 core stage upgrade. Thrust increased more than 30% from Vulcain 1. First flight 2002.

Vulkan 1 LOx/LH2 propellant rocket stage. Original version of Energia core as used on Vulkan booster, with in-line upper stages and payloads. Developed 1974-1976; cancelled when Energia / Buran development begun.

Vulkan Blok V LOx/LH2 propellant rocket stage. Upper stage design by KB Saturn for manned lunar expedition, large geosynchronous platform launch.

X-8 Rocketdyne LOx/LH2 rocket engine. Booster applications. Gas generator, pump-fed. Thrust and specific impulse values are at sea level.

XRS-2200 Rocketdyne LOx/LH2 rocket engine. Development ended 1999. Linear aerospike engine for X-33 SSTO technology demonstrator. Based on J-2S engine developed for improved Saturn launch vehicles in the 1960's.

YF-73 Beijing Wan Yuan LOx/LH2 rocket engine. In development. Gas-generator turbopump. Gimbaled engine. Used on CZ-3 launch vehicle. First flight 1984.

YF-75 Beijing Wan Yuan LOx/LH2 rocket engine. In development. Gas-generator turbopump. Gimbaled engine. First flight 1994.

YF-77 CAALPT LOx/LH2 rocket engine for next generation Chinese launch vehicle series. It was an indigenous development based on Chinese experience with the YF-73 and YF-75 upper stage engines.

Engines: RD-57, RD-57M, LR87 LH2, Sea Dragon-2 engine, Truax LH2, AJ23-142, J-2T-200K, AJ23-141, NK-15VM, HG-3-SL, J-2-SL, HG-3, J-2T-250K, Aerospike Annular Booster, AMPS-1, J-2S, RL-10A-3, AJ23-143, AJ23-147, AJ23-127, SSME Demonstrator Booster, RM-1500H, NK-35, HM7-A, AEC engine, HM7-B, STME, AJ23-144, AJ23-145, RD-0120M, PW 1000000 lb LH2, RD-0122, RL-10A-5KA, Plug-Nozzle J-2, RL-10B-X, RL-10C-X, RD-56M, RD-0120TD, LE-5B engine, LE-7A, RD-57A-1, RD-0126A, RD-0128, RD-0132, RD-0133, RD-0750, RD-0126E, RD-0146, XRS-2200, HIMES engine, MB-35, MB-45, MB-60, RS-74, SPW-2000, IPD, RS-68 Regen, RS-68B, J-2X, 5 mlbf, AJ-60C, ASE, ATCRE, CD Module, Chamber/single nozzle, Cobra, Helios Stage 1, HM-10, HP-1, J-2, LE-5 engine, LE-7, LE-5A, L-5.00H, L-5.25H, L6H, LH2 2000/3000 lbf thrust, LH2-80k, LR129, M-1, MBB-ATC500, Ottobrunn 300N, P320, Plug-Nozzle Pegasus, Plug-Nozzle Rombus, Plug-Nozzle SASSTO, Plug-Nozzle SERV, Plug-Nozzle SSME, RL-60, RL-10A-1, RL-10A-4-1, RL-10A-4-2, RL-10B-2, RL-10A-4, RL-10A-5, RL-10A-3A, RL-10C, RBCC, RD-0120, RD-0126, RD-0131, RD-135, RD-54, RD-56, RD-701, RD-704, RL-10, RL-50, RS-2100, RS-2200, RS-52, RS-68, RS-71, RS-800, RS-XXX, SSME, SSME Plus, SSME Study, Toroid FD, Toroidal 400k, Toroidal 560k, TR-106, Vulcain 2, Vinci, Vulcain, X-8, YF-50t, YF-73, YF-75. Spacecraft: Apollo Direct RM, LM Langley Lighter, LM Langley Lightest, Bono Saucer, TRW Mars, NASA Mars Flyby 1965, LLV L-I, LLV L-II, NASA Mars Expedition 1971, Space Tug, OTV, LLV, LEV, STCAEM Cryogenic Aerobrake, KVRB, Mars Oz, CEV SAIC, CEV Schafer, CZ H-18. Launch Vehicles: Mustard, VTOHL 45t, VTOHL 9t, Venturestar. Stages: CZ-YF-73, N1 Block R, Jarvis-2, Sea Horse-2, Saturn II, Saturn IV, Pegasus Tanks, Sanger I-2, Starlifter, Titan C-2, N1 Block S, SLS Stage A, Angara Stage 2, Sanger I-1, Astro-2, Shuttle Orbiter, Rombus Tank, Sanger II-2, Ares I-2, SLS Stage B, Ares Stage 2, CZ-NGLV-500, Cargo LV Stage 2, Starlifter Tank, Astro-1, Albatros Raketoplan, Shuttle Tank, Ares Stage 1, Vulkan 1, Magnum Core, McDonnell-Douglas ILRV Drop Tanks, Cargo LV Stage 1, Albatros Carrier Aircraft, Albatros Momentum Block, UR-700M-3, Shuttle SERV-1, VTOVL 150t, Sea Dragon-2, Hyperion Booster, Nova 59-4-3, Nova 59-4-4, Nova NASA-3, Centaur C, Helios A-1, Helios B-1, Helios C-1, Nova 4 J-2, Nova 60-8-2, Nova 60-8-3, Nova 9L-3, Nova 9L-4, Nova A-2, Nova A-3, Nova B-2, Nova B-3, Saturn IVB, Saturn S-II, Saturn S-II-4, Saturn S-II-8, Saturn S-II-C3, Saturn II C-5A, Saturn IVB C-3B, Saturn IVB C-5A, SLS Stage C, Saturn S-IVC, DAC Helios ISI-1, DAC Helios-1, Nova DAC 2-1, Nova DAC 2-2, Nova DAC ISI-1, Nova DAC ISI-2, Nova DAC-2, Nova GD-B-2, Nova GD-E-2, Nova GD-F-2, Nova GD-H-0, Nova GD-H-1, Nova GD-J-2, Nova MM 14A-2, Nova MM 14B-2, Nova MM 1B-2, Nova MM 1C-2, Nova MM 24G-2, Nova MM 33-1, Nova MM 34-0, Nova MM 34-1, Nova MM S10E-1 stage, Nova MM S10E-2 stage, Nova MM S10R-1 stage, Nova MM S10R-2 stage, Nova MM T10EE-1-1, Nova MM T10EE-1-2, Nova MM T10RE-1-1, Nova MM T10RR-2-1, Nova MM T10RR-2-2, Nova MM T10RR-3-1, Nova MM T10RR-3-2, Nova MM-1, OOST ISI stage, OOST stage, ROOST ISI stage, ROOST stage, Rombus core, N1 Block V-II, N1 Block V-III, Saturn IVB-A, Saturn MS-II-1, Saturn MS-II-1-J-2T-200K, Saturn MS-IVB-1, Saturn MS-IVB-2, Saturn MS-IVB-x, Pegasus VTOVL stage, Saturn II-INT-17, Saturn II-SL, Saturn MS-II-1A, Saturn MS-II-2, SASSTO stage, Saturn IVB (S-IB), Saturn MS-II-1-J-2T-250K, Saturn MS-II-3B, Saturn MS-IVB-1A, Saturn MS-IVB-3B, Saturn MS-IVB-4(S)B, Spacemaster-1, Spacemaster-2, Centaur D/E, Hyperion SSTO stage, Mustard 1, Mustard 2, Saturn IVB (S-V), Saturn MS-II-4(S)B, MLLV Core, Shuttle FR-3-1, Shuttle FR-3-2, Shuttle HCR-1, Shuttle HCR-2, Shuttle LCR-1, Shuttle LCR-2, Shuttle LS A-1, Shuttle LS A-2, Shuttle MDC-A-1, Shuttle MDC-A-2, Shuttle NAR A-1, Shuttle NAR A-2, Shuttle DC-3-1, Shuttle DC-3-2, Shuttle R134C-1, Shuttle R134C-2, Shuttle R134G-1, Shuttle R134G-2, Shuttle H33-1, Shuttle H33-2, Shuttle LS200-1, Ariane 1-3 H8, N1 Block Sr, Energia Core, Energia EUS, Vulkan Blok V, Ariane 2-3 H10, Shuttle LRB stage, Ariane 4-3 H10plus, Ariane 4-3 H10-3, ALS stage, Chang Cheng stage 1, Titan 5 stage, Interim HOTOL stage, NLS Core, NLS HLV stage, NLS Semistage, LE-5, Project 921 LV-2, Proton KM-4, Energia M-1, Ariane 5-1 H155, CZ H-18 stage, Centaur B-X, Centaur C-X, H-2-1, HIMES, LE-5EC, Delta 3-2, Ariane 5-1 EPC, Ariane 5-2 ESC A, Centaur IIIA, Centaur G, CLV Stage 2, Ariane 5-2 ESC B, Angara KVRB, Centaur V1, Centaur V2, Centaur IIA, Centaur IIIB, Centaur I, Centaur II, CZ-NGLV-HO, Centaur G STS, Centaur G Prime, Delta 4-2, Delta 4H-2, Delta RS-68, GSLV-3 stage, H-2A-1, H-2A LRB, LE-5B, McDonnell-Douglas ILRV stage, Shuttle Super Lightweight Tank.

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