Encyclopedia Astronautica
Lox/LH2



j2.jpg
J-2
Credit: © Mark Wade
Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser 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 adopted for the core of the space shuttle, and Centaur stages still fly today.

Liquid oxygen, as normally supplied, is of 99.5 percent purity and is covered in the United States by Military Specification MIL-P-25508. High purity liquid oxygen has a light blue colour and is transparent. It has no characteristic odour. Liquid oxygen does not burn, but will support combustion vigorously. The liquid is stable; however, mixtures of fuel and liquid oxygen are shock-sensitive. Gaseous oxygen can form mixtures with fuel vapours that can be exploded by static electricity, electric spark, or flame. Liquid oxygen is obtained from air by fractional distillation. The 1959 United. States production of high-purity oxygen was estimated at nearly 2 million tonnes. The cost of liquid oxygen, at that time, ex-works, was $ 0.04 per kg. By the 1980's NASA was paying $ 0.08 per kg.

In Russia hydrogen fuelled upper stages were designed and developed by the mid-1970's, but the Russians never seem to have found the extra performance to be worth the extra cost. Europe and China developed liquid oxygen/liquid hydrogen engines for upper stages of the Ariane and Long March launch vehicles.

The equilibrium composition of liquid hydrogen is 99.79 per cent parahydrogen and 0.21 per cent orthohydrogen. The boiling point of this composition is -253 deg C. Liquid hydrogen is transparent and without a characteristic odour. Gaseous hydrogen is colourless. Hydrogen is not toxic but is an extremely flammable material. The flammable limits of gaseous hydrogen in air are 4.0 to 75 volume percent.

Hydrogen is produced from by-product hydrogen from petroleum refining and the partial oxidation of fuel oil. The gaseous hydrogen is purified to 99.999+ per cent, and then liquefied in the presence of paramagnetic metallic oxides. The metallic oxides catalyse the ortho-para transformation of freshly liquefied hydrogen. Freshly liquefied hydrogen which has not been catalysed consists of a 3:1 ortho-para mixture and cannot be stored for any length of time because of the exothermic heat of conversion. The delivered cost of liquid hydrogen in 1960 was approximately $ 2.60 per kg. Large-scale production was expected to reduce the cost to $ 1.00 per kg. In the 1980's NASA was actually paying $ 3.60 per kg.

Oxidizer: LOX. Oxidizer: LOX. Fuel: LH2. Fuel: LH2. Propellant Formulation: LOX/LH2. Propellant Formulation: LOX/Slush LH2. Optimum Oxidizer to Fuel Ratio: 6. Optimum Oxidizer to Fuel Ratio: 4. Temperature of Combustion: 2,985 deg K. Temperature of Combustion: 2,985 deg K. Ratio of Specific Heats: 1.26. Ratio of Specific Heats: 1.26. Density: 0.28 g/cc. Density: 0.33 g/cc. Characteristic velocity c: 2,435 m/s (7,988 ft/sec). Characteristic velocity c: 2,435 m/s (7,988 ft/sec). Isp Shifting: 391 sec. Isp Shifting: 391 sec. Isp Frozen: 388 sec. Isp Frozen: 388 sec. Mol: 10.00 M (32.00 ft). Mol: 10.00 M (32.00 ft). Oxidizer Density: 1.140 g/cc. Oxidizer Density: 1.140 g/cc. Oxidizer Freezing Point: -219 deg C. Oxidizer Freezing Point: -219 deg C. Oxidizer Boiling Point: -183 deg C. Oxidizer Boiling Point: -183 deg C. Fuel Density: 0.071 g/cc. Fuel Density: 0.071 g/cc. Fuel Freezing Point: -259 deg C. Fuel Freezing Point: -259 deg C. Fuel Boiling Point: -253 deg C. Fuel Boiling Point: -253 deg C.

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

More... - Chronology...


Associated Spacecraft
  • 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. More...
  • Centaur C American space tug. 22 launches, (1961) to (1967). Upper stage / space tug - out of production. More...
  • 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. More...
  • LM Langley Lightest American manned lunar lander. Study 1961. Extremely light-weight open-cab lunar module design considered in early Langley studies. More...
  • 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. More...
  • TRW Mars American manned Mars expedition. Study 1963. In 1963 TRW designed a Mars expedition using aerobraking at both Mars and Earth, and a swingby of Venus on return. More...
  • 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. More...
  • Saturn MS-IVB-1 American space tug. Study 1965. Upper stage / space tug - Marshall studies, 1965. Launched by Saturn V. S-IVB structurally strengthened to handle larger payloads, otherwise unchanged More...
  • Saturn MS-IVB-2 American space tug. Study 1965. Upper stage / space tug - Douglas study, 1965. Launched by Saturn V. S-IVB with 315 k high pressure 3000 psia engine, 350,000 pounds propellant More...
  • Saturn MS-IVB-x American space tug. Study 1965. Upper stage / space tug - studied by NASA Marshall in 1965. Launched by Saturn V. S-IVB structurally strengthened to handle larger payloads, otherwise unchanged More...
  • 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. More...
  • 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. More...
  • Saturn MS-IVB-1A American space tug. Study 1966. Upper stage / space tug - Boeing study 1967. Launched by Saturn V. S-IVB with 16.5 foot stretch, 350,000 pounds propellant, standard J-2 engine. More...
  • Centaur D/E American space tug. 56 launches, (1967) to (1983). Upper stage / space tug - out of production. Launched by Atlas Centaur D; Titan 3E. More...
  • Saturn MS-IVB-3B American space tug. Study 1967. Upper stage / space tug - Boeing study, 1967. Launched by Saturn V. S-IVB with 16.5 foot stretch, 350,000 pounds propellant, 1 x 400,000 pound thrust toroidal engine. More...
  • Saturn MS-IVB-4(S)B American space tug. Study 1967. Upper stage / space tug - studied by Boeing in 1967. Standard S-IVB but with structural strength increased from 78% to 217% depending on station, resulting in 11.8% increase in empty weight. More...
  • 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. More...
  • 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. More...
  • N1 Block Sr Russian space tug. Study 1971. Upper stage / space tug - developed 1971-1974 to support manned lunar expedition. Replaced Blok R/Blok S previously under development. More...
  • Space Tug American space tug. Study 1971. 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. More...
  • Centaur I American space tug. 18 launches, (1984) to (1997). Upper stage / space tug - out of production. Launched by Atlas I. More...
  • 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. More...
  • Centaur G Prime American space tug. Cancelled 1987. Upper stage / space tug - out of production. Centaur for Shuttle payload bay. Cancelled after Challenger disaster on safety grounds. More...
  • Centaur G American space tug. 22 launches, (1989) to (1998). Upper stage / space tug - out of production. Centaur for Titan 4 More...
  • LEV American manned lunar lander. Study 1989. The Lunar Excursion Vehicle (LEV) figured in numerous NASA studies of the 1980's and 1990's. More...
  • Centaur II American space tug. 10 launches, (1991) to (1998). Upper stage / space tug - out of production. Launched by Atlas II. More...
  • STCAEM Cryogenic AeroBrake American manned Mars expedition. Study 1991. The STCAEM cryogenic / aerobrake (CAB) concept was used as the NASA reference vehicle. More...
  • Centaur IIA American space tug. 48 launches, (1992) to (2002). Upper stage / space tug - out of production. Launched by Atlas IIA. More...
  • KVRB Russian space tug. Study 1992. Upper stage / space tug - design 1992. High energy upper stage for Proton, never put into production. More...
  • CZ-3A-3 Chinese space tug. 11 launches, (1994) to (2000). Upper stage / space tug - in production. Launched by CZ-3A, CZ-3B, and CZ-3C. More...
  • Centaur C-X American space tug. Study 2001. Upper stage / space tug - in development 1998, not put into production. More...
  • Centaur B-X American space tug. Study 2000. Upper stage / space tug - in development 1998, not put into production. More...
  • Centaur IIIA American space tug. One launch, , 2000. Upper stage / space tug - out of production. Single-engine Centaur for Atlas IIIA. More...
  • Centaur V2 American space tug. Study 2001. Upper stage / space tug - in production. Twin engined Centaur for Atlas V, powered by two Pratt & Whitney RL10A-4-2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. More...
  • Delta 4H - 2 American space tug. Study 2001. Upper stage / space tug - in production. Delta 4 second stage with hydrogen tank increased to 5.1 m diameter. More...
  • Centaur V1 American space tug. One launch, , 2002. Upper stage / space tug - in production. Single engined Centaur for Atlas V, powered by one Pratt & Whitney RL10A-4-2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. More...
  • Centaur IIIB American space tug. One launch, , 2002. Upper stage / space tug - out of production. Dual-engine Centaur for Atlas IIIB. More...
  • Delta 4 - 2 American space tug. One launch, , 2002. Upper stage / space tug - in production. Delta 3 second stage with hydrogen tank stretch. More...
  • Ariane 5 ESC B German space tug. Study 2006. Upper stage / space tug - in production. New upper stage for Ariane 5. More...
  • 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. More...
  • CZ-NGLV-HO Chinese space tug. Study 2008. Upper stage / space tug - in development 2004. The upper stage for the Chinese Next Generation Launch Vehicle was a modification of the CZ-3B upper stage. More...
  • 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. More...
  • CEV Schafer American manned spacecraft. Study 2012. Schafer proposed a lightweight 11 metric ton integral CEV, staged from L1. More...

Associated Engines
  • 5 mlbf Notional lox/lh2 rocket engine. 41,361 kN. OOST, ROOST studies 1963. Isp=410s. First flight 1977. More...
  • AEC Rocketdyne lox/lh2 rocket engine. 66.695 kN. Advanced Expander Cycle Engine. Expander regenerator, pump-fed. Isp=481s. More...
  • Aerospike Annular Booster Rocketdyne lox/lh2 rocket engine. 1111.662 kN. Aerospike Typical Annular Booster . Pressure-fed. Thrust from 50,000 to 250,000 lbs at altitude. Isp=450s. More...
  • AJ-60C Aerojet lox/lh2 rocket engine. 280 kN. Design 2000. Isp=470s. Design announced on 3 October 2000 for a new cryogenic upper-stage engine aimed at the very large commercial spacecraft market. More...
  • AMPS-1 Rocketdyne lox/lh2 rocket engine. 80.032 kN. Advanced Maneuvering Propulsion System Booster. Pressure-fed. Isp=468s. More...
  • ASE Rocketdyne lox/lh2 rocket engine. 88.926 kN. Advanced Space Engine. Staged combusion, pump-fed. Isp=473s. More...
  • ATCRE Notional lox/lh2 rocket engine. 1280 kN. Study 1985. Isp=490s. Used on Sanger II launch vehicle. More...
  • CD Module Notional lox/lh2 rocket engine. 7361 kN. Study 1969. Isp=420s. CD Modules - conceptual engines of various thrusts, according to design - were clustered in Martin Marietta Nova designs More...
  • Chamber/single nozzle Notional lox/lh2 rocket engine. 13,231 kN. Study 1963. Isp=455s. 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. More...
  • Cobra Pratt and Whitney lox/lh2 rocket engine. 4500 kN. 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. More...
  • Helios Stage 1 Notional lox/lh2 rocket engine. 1667 kN. Study 1960. Engines for booster stage with Lox tanks only to take nuclear second stage to stratosphere. Isp=400s. Helios A, B, C studies. More...
  • HG-3 Rocketdyne lox/lh2 rocket engine. 1400.7 kN. Study 1967. Isp=451s. 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. More...
  • HG-3-SL Rocketdyne lox/lh2 rocket engine. 1387 kN. Study 1966. Isp=450s. 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. More...
  • HIMES Mitsubishi lox/lh2 rocket engine. 137.3 kN. Design 1999. Isp=452s. Used on H-2 HIMES launch vehicle. More...
  • HM-10 SEP, Ottobrunn lox/lh2 rocket engine. 61.8 kN. Developed 1990's. Engine for potential Ariane 5 upper stage. Isp=443s. More...
  • HM7-A SEP, Ottobrunn lox/lh2 rocket engine. 61.7 kN. Development begun 1973. Out of production. Isp=443s. Used on Ariane 1 launch vehicle. First flight 1979. More...
  • HM7-B SEP, Ottobrunn lox/lh2 rocket engine. 70 kN. Isp=447s. 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. More...
  • HP-1 Notional lox/lh2 rocket engine. 6536 kN. Study 1963. Operational date would have been December 1974. Isp=451s. Used in Martin Nova studies MM 24G, MM 33. More...
  • IPD Rocketdyne lox/lh2 rocket engine. 1100 kN. 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. More...
  • J-2 Rocketdyne lox/lh2 rocket engine. 1033.1 kN. Study 1961. Isp=421s. Used in Saturn IVB stage in Saturn IB and Saturn V, and Saturn II stage in Saturn V. Gas generator, pump-fed. First flight 1966. More...
  • J-2-SL Rocketdyne lox/lh2 rocket engine. 996.7 kN. Study 1966. Sea level version of J-2 with reduced expansion ratio proposed for Saturn II first stage use. Isp=390s. More...
  • J-2S Rocketdyne lox/lh2 rocket engine. 1138.5 kN. Developed 1965-1969. Isp=436s. 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. More...
  • J-2T-200K Rocketdyne lox/lh2 rocket engine. 889.3 kN. Study 1965. Proposed for later versions of Saturn V. Toroidal aerospike plug nozzle version of J-2. Isp=435s. More...
  • J-2T-250K Rocketdyne lox/lh2 rocket engine. 1111.6 kN. Study 1967. Proposed for later versions of Saturn V. Toroidal aerospike plug nozzle version of J-2. Isp=441s. More...
  • J-2X Rocketdyne lox/lh2 rocket engine. 1310 kN. Ares I launch vehicle second stage. In development 2006-2016. Isp=448s. 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. More...
  • L-5.00H Notional lox/lh2 rocket engine. 30,684 kN. Study 1963. Isp=428s. Used on Nova GD-H launch vehicle. More...
  • L-5.25H Notional lox/lh2 rocket engine. 27,350 kN. Study 1963. Isp=410s. Engines for recoverable booster engine package 'half stage' of a 1 1/2 stage arrangement. Used on Nova GD-H launch vehicle. More...
  • L6H Notional lox/lh2 rocket engine. 122,748 kN. Study 1963. Operational date would have been June 1976. Used in booster stage (engines only). Isp=439s. Used on Nova MM 34 launch vehicle. More...
  • LE-5 Mitsubishi lox/lh2 rocket engine. 103 kN. Isp=450s. Used on H-1 launch vehicle. First flight 1986. More...
  • LE-5A Mitsubishi lox/lh2 rocket engine. 121.5 kN. Isp=452s. Used on H-2 launch vehicle. First flight 1994. More...
  • LE-5B Mitsubishi lox/lh2 rocket engine. 137 kN. In production. Isp=447s. 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. More...
  • LE-7 Mitsubishi lox/lh2 rocket engine for H-2 upper stage. 1078 kN. Staged combustion turbopump. No throttle capability. Isp=446s. First flight 1994. More...
  • LE-7A Mitsubishi lox/lh2 rocket engine. 1098 kN. In production. Isp=438s. 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. More...
  • LH2-80k Notional lox/lh2 rocket engine. 355.7 kN. Study 1959. Isp=425s. Used on Nova 4L launch vehicle. More...
  • 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. More...
  • LR87 LH2 Aerojet lox/lh2 rocket engine. 667 kN. 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. More...
  • M-1 Aerojet lox/lh2 rocket engine. 5335.9 kN. Study 1961. Isp=428s. Engine developed 1962-1966 for Uprated Saturn and Nova million-pound payload boosters to support manned Mars missions. Reached component test stage before cancellation. More...
  • MB-35 Rocketdyne lox/lh2 rocket engine. 156 kN. Design 2004. Isp=467s. Mitsubishi / Boeing joint project for an engine for Delta IV cryogenic upper stages. Expander bleed, pump-fed. More...
  • MB-45 Rocketdyne lox/lh2 rocket engine. 200 kN. Design 2004. Isp=467s. Mitsubishi / Boeing joint project for an engine for Delta IV cryogenic upper stages, announced February 2000. More...
  • MB-60 Rocketdyne lox/lh2 rocket engine. 266.7 kN. Design 2004. Isp=467s. Mitsubishi / Boeing joint project for an engine for Delta IV cryogenic upper stages. Expander bleed, pump-fed. More...
  • MBB-ATC500 MBB lox/lh2 rocket engine. 441.3 kN. Study 1969. Isp=460s. Used on Beta launch vehicle. More...
  • Mustard Notional lox/lh2 rocket engine. 2157.4 kN. Study 1968. Isp=405s. Used on Mustard launch vehicle. More...
  • NK-15VM Kuznetsov lox/lh2 rocket engine. 1960 kN. N-1 stage 2 (block B) replacement. Design 1972. Derivative of NK-15 with kerosene replaced by hydrogen. Canceled before hot-tests. More...
  • NK-35 Kuznetsov lox/lh2 rocket engine. 1960 kN. 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. More...
  • Ottobrunn 300N Ottobrunn lox/lh2 rocket engine. 300 N. Upper stages. Developed 2000. Isp=415s - highest value ever achieved in Europe for an engine of such small size. More...
  • P320 Rocketdyne, Friedrichshafen lox/lh2 rocket engine. Development. Launch thrust 129 kN. BORD 1/P320 BOELKOW (Germany)/Rocketdyne Technology. Pressure-fed. More...
  • Plug-Nozzle J-2 Rocketdyne lox/lh2 rocket engine. 6864.6 kN. Study 1993. Plug nozzle version of J-2 proposed for certain Saturn V upgrades in late 1960's. Isp=425s. Used on DC-I launch vehicle. More...
  • Plug-Nozzle SSME Notional lox/lh2 rocket engine. 3728.7 kN. Study 1978. Isp=485s. Used on VTOVL launch vehicle. More...
  • Plug-Nozzle SERV Notional lox/lh2 rocket engine. 31,980.2 kN. Study 1971. Isp=455s. Used on Shuttle SERV launch vehicle. More...
  • Plug-Nozzle Pegasus Notional lox/lh2 rocket engine. 23,928 kN. Study 1966. Isp=459s. Used on Pegasus VTOVL launch vehicle. More...
  • PW 1000000 lb LH2 Pratt and Whitney lox/lh2 rocket engine. 4457 kN. 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. Isp=425s. More...
  • RBCC Rocketdyne lox/lh2 rocket engine. Launch thrust 111.158 kN. Isp>400s. Rocket Based Combined-Cycle A5 Development Engine; integrated rocket, air-augmented rocket, ramjet, and sramjet propulsion elements into a single flowpath. More...
  • RD-0128 Kosberg lox/lh2 rocket engine. 98 kN. upper stage. Design concept 1996-. Concept for a cryogenic engine for upper stages. One single chamber with bell nozzle, separate turbopumps. Isp=474s. More...
  • RD-0133 Kosberg lox/lh2 rocket engine. 98 kN. upper stage. Design concept 1996-. Concept for a cryogenic engine for upper stages. Four chambers with bell nozzles, common turbopump. Isp=467s. More...
  • RD-0132 Kosberg lox/lh2 rocket engine. 98 kN. 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. Isp=469s. More...
  • RD-0122 Kosberg lox/lh2 rocket engine. 2313 kN. Energia-M core stage. Planned for Angara central stage. Developed 1990-. Isp=460s. Upgrade of RD-0120 engine for Energia-M launcher with increased thrust. Prototype from RD-0120 hardware. More...
  • RD-0131 Kosberg lox/lh2 rocket engine. 98 kN. upper stage. Design concept 1996-. Concept for a cryogenic engine for upper stages. Single annular chamber with expansion-deflection nozzle, common turbopump. Isp=467s. More...
  • RD-0146 Kosberg lox/lh2 rocket engine. 98.1 kN. Centaur upper stage (Atlas); high performance upper stages for Onega, Proton, Angara launch vehicles. Design concept 1998-. Isp=463s. More...
  • RD-0126 Kosberg lox/lh2 rocket engine. 39.2 kN. Space tugs or upper stage for Onega or Yastreb versions of Soyuz. Isp=476s. Single annular chamber with expansion-deflection nozzle, separate turbopumps. Design concept 1993. Hot-tests in 1998. More...
  • RD-0120M Kosberg lox/lh2 rocket engine. 1961 kN. Energia-M core stage. Development ended 1993. Isp=455s. 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. More...
  • RD-0120 Kosberg lox/lh2 rocket engine. 1961 kN. Energia core stage. Design 1987. Isp=455s. First operational Russian cryogenic engine system, built to the same overall performance specifications as America's SSME, but using superior Russian technology. More...
  • RD-0126A Kosberg lox/lh2 rocket engine. 98 kN. Upper stages. Design concept 1996-. Concept for a cryogenic engine for upper stages. Single annular chamber with expansion-deflection nozzle, common turbopump. Isp=476s. More...
  • RD-0126E Kosberg lox/lh2 rocket engine. 39.2 kN. Upper stages. Design concept 1998-. Concept for a cryogenic engine for upper stages. Single annular chamber with straight expansion nozzle, common turbopump. Isp=472s. More...
  • RD-135 Glushko lox/lh2 rocket engine. upper stage. Developed -1976. Experimental cryogenic engine. (Ref. May be not correct.) More...
  • RD-56M Isayev lox/lh2 rocket engine. 73.580 kN. Proton and Angara upper stage KVRB, 12KRB upper stage for GSLV (India). In development. Isp=461s. First flight 2001. More...
  • RD-57M Lyulka lox/lh2 rocket engine. 397 kN. Vulkan Blok V. Development ended 1976. Isp=461s. Version with extendible nozzle. Length 4.06 / 2.61 m. Specific impulse 461 / 448 sec. Area ratio 170 / 87.6. More...
  • RD-57 Lyulka lox/lh2 rocket engine. 392 kN. 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. Isp=456s. More...
  • RD-56 Isayev lox/lh2 rocket engine. 69.6 kN. 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. Isp=462s. More...
  • RD-54 Lyulka lox/lh2 rocket engine. 392 kN. N1 concept stage III. Developed 1960-75. Isp=440s. More...
  • RD-57A-1 Lyulka lox/lh2 rocket engine. 395 kN. Developed 1995-98. Isp=460s. New version of RD-57M for SSTO-demonstrator proposed by Aerojet. Optimized nozzle contour for performance increase, new chamber material for weight reduction. More...
  • RL-10A-4 Pratt and Whitney lox/lh2 rocket engine. 92.5 kN. Out of production. Isp=449s. Centaur stage for Atlas IIA, Atlas IIAS. First flight 1992. More...
  • RL-10C-X Pratt and Whitney lox/lh2 rocket engine. 110.8 kN. Design concept 1994. Isp=450s. More...
  • RL-10C Pratt and Whitney lox/lh2 rocket engine. 155.7 kN. In Production. Used in Delta 3 - 2. Isp=450s. First flight 1998. More...
  • RL-10B-X Pratt and Whitney lox/lh2 rocket engine. 93.4 kN. Design concept 1994. Isp=470s. More...
  • RL-10B-2 Pratt and Whitney lox/lh2 rocket engine. 110 kN. In production. Isp=462s. Used on Delta 3 , Delta IV launch vehicles. First flight 1998. Extendable exit cone for increased specific impulse; electromechanical actuators replace hydraulic systems. More...
  • RL-10A-5KA Pratt and Whitney lox/lh2 rocket engine. 100.488 kN. Kistler proposal. Design 1992. Isp=398s. Throttleable to 30% of thrust, sea level version of RL10 with extendable nozzle for high altitude operation. More...
  • RL-10A-5 Pratt and Whitney lox/lh2 rocket engine. 64.7 kN. Isp=373s. 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. More...
  • RL-10A-4-2 Pratt and Whitney lox/lh2 rocket engine. 99.1 kN. In production. Isp=451s. Used on Atlas IIIB launch vehicle. First flight 2002. Two engines; electro-mechanical thrust vector control actuators replaced earlier hydraulically actuated system. More...
  • RL-10A-3A Pratt and Whitney lox/lh2 rocket engine. 73.4 kN. Isp=444s. Used on Centaur stage atop Atlas G, Atlas I, Atlas II, Titan 4. First flight 1984. More...
  • RL-10A-3 Pratt and Whitney lox/lh2 rocket engine. 65.6 kN. Study 1968. Isp=444s. First flight 1967. More...
  • RL-10A-1 Pratt and Whitney lox/lh2 rocket engine. 66.7 kN. Isp=425s. Version used on Atlas Centaur LV-3C, and proposed for various early Saturn launch vehicle designs. First flight 1961. More...
  • RL-10 Pratt and Whitney lox/lh2 rocket engine. 66.7 kN. Isp=410s. Early version as proposed for Nova A, Nova B, Saturn B-1, Saturn C-2, Saturn C-3, Saturn I. First flight 1961. More...
  • RL-10A-4-1 Pratt and Whitney lox/lh2 rocket engine. 99.1 kN. Out of production. Isp=451s. 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. More...
  • RL-50 Pratt and Whitney lox/lh2 rocket engine. 290 kN. Development. Isp=472s. Advanced, high-performance upper-stage rocket engine proposed by Pratt & Whitney for both domestic and international launch vehicles. More...
  • RL-60 Pratt and Whitney lox/lh2 rocket engine. 289.1 kN. Design. Isp=470s. 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. More...
  • RM-1500H Rocketdyne lox/lh2 rocket engine. 6.660 kN. Space Shuttle Orbiter Auxiliary Propulsion. Pressure-fed. Isp=400s. More...
  • RS-2200 Rocketdyne lox/lh2 rocket engine. 2201 kN. Development cancelled 1999. Isp=455s. Linear Aerospike Engine developed for use on the Lockheed Reusable Launch Vehicle, the production follow-on to the X-33. More...
  • RS-2100 Rocketdyne lox/lh2 rocket engine. Launch thrust 2047.6 kN. Next Generation Launch Vehicle Booster. Full flow staged combustion, pump-fed. Thrust and specific impulse values are at sea level. More...
  • RS-52 Rocketdyne lox/lh2 rocket engine. 0.107 kN. 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. Isp=405s. More...
  • 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. More...
  • RS-68B Rocketdyne lox/lh2 rocket engine. Design concept -2004. Upgrade (details not specificed) to basic RS-68 for Delta IV Heavy growth versions. More...
  • RS-68 Rocketdyne lox/lh2 rocket engine. 3312 kN. In production. Isp=420s. First new large liquid-fueled rocket engine developed in America in more than 25 years. Powered the Delta IV booster. First flight 2002. More...
  • RS-71 Rocketdyne lox/lh2 rocket engine. 31.126 kN. Development ended 1999. Linear Aerospike SR-71 Experiment. Pressure-fed. Isp=430s. More...
  • RS-74 Rocketdyne lox/lh2 rocket engine. Launch thrust 1112 kN. Next Generation Launch Vehicle Booster. Full flow staged combustion, pump-fed. Thrust and specific impulse values are at sea level. More...
  • RS-800 Rocketdyne lox/lh2 rocket engine. 4110 kN. Design concept -2004. New high-thrust cryogenic engine for Delta IV Heavy growth versions. More...
  • RS-XXX Rocketdyne lox/lh2 rocket engine. 8230 kN. Design concept -2004. New high-thrust cryogenic engine concept for Next Generation Delta with 7 m diameter modules. More...
  • Sea Dragon-2 Aerojet lox/lh2 rocket engine. 62,270 kN. Design, 1962. Truax pressure fed design. Diameter of extended nozzle 30 m. Specific impulse estimated from booster performance figures. Isp=320s. More...
  • SPW-2000 SNECMA, Pratt and Whitney lox/lh2 rocket engine. 230.4 kN. Design 2000. New upper-stage cryogenic engine for the upgraded Ariane-5, the Atlas-5, and other new vehicles. More...
  • SSME Plus Notional lox/lh2 rocket engine. 3728.7 kN. VTOHL studies, 1978. Isp=467s. More...
  • SSME Rocketdyne lox/lh2 rocket engine. 2278 kN. In production. Isp=453s. 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. More...
  • STME Rocketdyne lox/lh2 rocket engine. 2890 kN. Cancelled 1984. Isp=430s. 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. More...
  • Toroid FD Notional lox/lh2 rocket engine. 20,015 kN. Study 1963. Operational date would have been December 1976. Engines for recoverable stage. Isp=455s. Used on Nova MM T10RR-2 launch vehicle. More...
  • Toroidal 400k Notional lox/lh2 rocket engine. 1778 kN. Study 1967. Isp=447s. Used on Saturn V-3B launch vehicle. More...
  • TR-106 TRW lox/lh2 rocket engine. 2892 kN. Development. Innovative TRW 650K Low Cost Pintle Engine, test fired at NASA's test center in October 2000. More...
  • Truax LH2 Aerojet lox/lh2 rocket engine. 147.1 kN. Test 1962. Used in Sea Horse-2. Isp=425s. More...
  • Vinci Snecma, Ottobrunn lox/lh2 rocket engine. 180 kN. Upper Stages. In development. Isp=467s. Advanced expander cycle cryogenic propellant rocket engine with the capability of five in-space restarts. First hot-fire tests 2005. First flight 2010. More...
  • Vulcain SEP, Ottobrunn lox/lh2 rocket engine. 1075 kN. In production. Isp=431s. Powered the cryogenic core stage of Ariane 5. First flight 1996. More...
  • Vulcain 2 SEP, Ottobrunn lox/lh2 rocket engine. 1350 kN. In development. Isp=434s. New generator cycle rocket engine for an Ariane 5 core stage upgrade. Thrust increased more than 30% from Vulcain 1. First flight 2002. More...
  • X-8 Rocketdyne lox/lh2 rocket engine. Launch thrust 400.1 kN. Booster applications. Gas generator, pump-fed. Thrust and specific impulse values are at sea level. More...
  • XRS-2200 Rocketdyne lox/lh2 rocket engine. 1192 kN. Development ended 1999. Isp=439s. Linear aerospike engine for X-33 SSTO technology demonstrator. Based on J-2S engine developed for improved Saturn launch vehicles in the 1960's. More...
  • YF-50t CAALPT lox/lh2 rocket engine. 700 kN. In development. Isp=432s. New Lox/LH2 engine for next generation Chinese launch vehicles. It is an indigenous development based on Chinese experience with the YF-73 and YF-75 upper stage engines. More...
  • YF-73 Beijing Wan Yuan lox/lh2 rocket engine. 11 kN. In development. Gas-generator turbopump. Gimballed engine. Isp=425s. Used on CZ-3 launch vehicle. First flight 1984. More...
  • YF-75 Beijing Wan Yuan lox/lh2 rocket engine. 78.5 kN. In development. Gas-generator turbopump. Gimballed engine. Isp=440s. First flight 1994. More...

Associated Stages
  • Albatros Raketoplan Lox/LH2 propellant rocket stage. Loaded/empty mass 320,000/82,000 kg. Thrust 1,960.00 kN. Vacuum specific impulse 455 seconds. Configuration: delta wing with wingtip vertical stabilizers. More...
  • Albatros Carrier Aircraft Lox/LH2 propellant rocket stage. Loaded/empty mass 1,250,000/210,000 kg. Thrust 7,840.00 kN. Vacuum specific impulse 455 seconds. Configuration: delta wing with wingtip vertical stabilizers and canards. Engine type and performance, empty weight estimated. More...
  • Albatros Momentum Block Lox/LH2 propellant rocket stage. Loaded/empty mass 2,000,000/1,800,000 kg. 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 Alexeyev Hydrofoil/Ekranoplan OKB. More...
  • ALS Lox/LH2 propellant rocket stage. Loaded/empty mass 780,000/60,000 kg. Thrust 15,477.16 kN. Vacuum specific impulse 435 seconds. More...
  • Angara KVRB Lox/LH2 propellant rocket stage. Loaded/empty mass 23,300/3,500 kg. Thrust 73.50 kN. Vacuum specific impulse 461 seconds. Planned version for Angara. 5 restarts. More...
  • Angara Stage 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 75,000/9,000 kg. Thrust 1,960.00 kN. Vacuum specific impulse 455 seconds. Unique configuration with oxidizer in core and fuel in two tanks strapped on in parallel - all of rail-transportable 3.9 m diameter. Buildt 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. More...
  • Ares I-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 158,500/20,500 kg. Thrust 1,304.00 kN. Vacuum specific impulse 448 seconds. Second stage figures as of summer 2008. Dry mass includes 2500 kg for avionics bay. More...
  • Ares Stage 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 787,700/64,200 kg. Thrust 8,705.99 kN. Vacuum specific impulse 453 seconds. Core vehicle proposed by NASA for Project Constellation exploration of moon and Mars. It would use shuttle external tank tooling. All masses estimated. More...
  • Ares Stage 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 172,000/13,200 kg. Thrust 1,113.00 kN. Vacuum specific impulse 465 seconds. Second stage proposed later in design stage by NASA for launch of CEV into low earth orbit. All masses estimated. More...
  • Ariane H155 Lox/LH2 propellant rocket stage. Loaded/empty mass 170,800/12,700 kg. Thrust 1,114.00 kN. Vacuum specific impulse 430 seconds. Chamber pressure 108 bar; expansion ratio 45.0; propellant mix ratio 5.3. More...
  • Ariane H8 Lox/LH2 propellant rocket stage. Loaded/empty mass 9,687/1,457 kg. Thrust 61.67 kN. Vacuum specific impulse 443 seconds. High energy upper stage for Ariane booster series. More...
  • Ariane H10plus Lox/LH2 propellant rocket stage. Loaded/empty mass 12,800/1,740 kg. Thrust 62.70 kN. Vacuum specific impulse 446 seconds. More...
  • Ariane H10-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 12,310/1,570 kg. Thrust 62.70 kN. Vacuum specific impulse 446 seconds. More...
  • Ariane 5 ESC B Lox/LH2 propellant rocket stage. Loaded/empty mass 27,500/3,400 kg. Thrust 153.90 kN. Vacuum specific impulse 467 seconds. New upper stage for Ariane 5. More...
  • Ariane 5 EPC Lox/LH2 propellant rocket stage. Loaded/empty mass 186,000/12,700 kg. Thrust 1,114.00 kN. Vacuum specific impulse 434 seconds. 15.2 tonnes increased propellant by moving liquid oxygen bulkhead. More...
  • Ariane 5 ESC A Lox/LH2 propellant rocket stage. Loaded/empty mass 16,500/2,100 kg. Thrust 64.70 kN. Vacuum specific impulse 446 seconds. Uses engine and oxygen tank from Ariane 4 + new liquid hydrogen tank. More...
  • Ariane H10 Lox/LH2 propellant rocket stage. Loaded/empty mass 12,000/1,600 kg. Thrust 62.70 kN. Vacuum specific impulse 446 seconds. More...
  • Astro-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 302,183/32,558 kg. Thrust 8,820.00 kN. Vacuum specific impulse 410 seconds. Engines 1 x M-1 plus 2 x J-2 More...
  • Astro-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 89,290/14,000 kg. Thrust 1,170.00 kN. Vacuum specific impulse 425 seconds. Engines 2 x RL-10 plus 1 x J-2 More...
  • Beta Lox/LH2 propellant rocket stage. Loaded/empty mass 450,000/40,000 kg. Thrust 5,736.00 kN. Vacuum specific impulse 460 seconds. More...
  • Cargo LV Stage 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,093,330/88,449 kg. Thrust 10,441.06 kN. Vacuum specific impulse 452 seconds. 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 tonne SSME engine pod. More...
  • Cargo LV Stage 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 227,036/19,343 kg. Thrust 2,442.07 kN. Vacuum specific impulse 451 seconds. 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. More...
  • Centaur II Lox/LH2 propellant rocket stage. Loaded/empty mass 18,833/2,053 kg. Thrust 146.80 kN. Vacuum specific impulse 444 seconds. More...
  • Centaur C Lox/LH2 propellant rocket stage. Loaded/empty mass 15,600/1,996 kg. Thrust 133.45 kN. Vacuum specific impulse 425 seconds. The first high-energy liquid oxygen/liquid hydrogen propellant stage in history. Despite initial development problems, the Centaur is entering its sixth decade of development and production. More...
  • Centaur IIIB Lox/LH2 propellant rocket stage. Loaded/empty mass 22,960/2,130 kg. Thrust 198.32 kN. Vacuum specific impulse 451 seconds. 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. More...
  • Centaur V2 Lox/LH2 propellant rocket stage. Loaded/empty mass 23,050/2,250 kg. Thrust 198.40 kN. Vacuum specific impulse 451 seconds. 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. More...
  • Centaur IIIA Lox/LH2 propellant rocket stage. Loaded/empty mass 18,710/1,905 kg. Thrust 99.16 kN. Vacuum specific impulse 451 seconds. Single-engine Centaur for Atlas IIIA. More...
  • Centaur IIA Lox/LH2 propellant rocket stage. Loaded/empty mass 19,073/2,293 kg. Thrust 185.01 kN. Vacuum specific impulse 449 seconds. More...
  • Centaur V1 Lox/LH2 propellant rocket stage. Loaded/empty mass 22,825/2,026 kg. Thrust 99.19 kN. Vacuum specific impulse 451 seconds. 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. More...
  • Centaur G STS Lox/LH2 propellant rocket stage. Loaded/empty mass 16,327/2,600 kg. Thrust 146.80 kN. Vacuum specific impulse 444 seconds. More...
  • Centaur G Prime Lox/LH2 propellant rocket stage. Loaded/empty mass 19,501/3,000 kg. Thrust 146.80 kN. Vacuum specific impulse 444 seconds. Centaur for Shuttle payload bay. Cancelled after Challenger disaster on safety grounds. More...
  • Centaur G Lox/LH2 propellant rocket stage. Loaded/empty mass 23,880/2,775 kg. Thrust 146.80 kN. Vacuum specific impulse 444 seconds. Centaur for Titan 4 More...
  • Centaur C-X Lox/LH2 propellant rocket stage. Loaded/empty mass 19,138/2,358 kg. Thrust 110.80 kN. Vacuum specific impulse 450 seconds. Conceptual design. Not put into production. More...
  • Centaur B-X Lox/LH2 propellant rocket stage. Loaded/empty mass 19,138/2,358 kg. Thrust 186.80 kN. Vacuum specific impulse 470 seconds. Conceptual design. Not put into production. More...
  • Centaur D/E Lox/LH2 propellant rocket stage. Loaded/empty mass 16,258/2,631 kg. Thrust 131.22 kN. Vacuum specific impulse 444 seconds. More...
  • Centaur I Lox/LH2 propellant rocket stage. Loaded/empty mass 15,600/1,700 kg. Thrust 146.80 kN. Vacuum specific impulse 444 seconds. More...
  • Chang Cheng 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 57,000/30,000 kg. Thrust 490.00 kN. Vacuum specific impulse 230 seconds. All characteristics except dimensions estimated, on assumption that stage used same propulsion systems as Shanghai upper stage. More...
  • CLV Stage 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 160,000/16,000 kg. Thrust 2,333.00 kN. Vacuum specific impulse 465 seconds. 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. More...
  • CZ H-18 Lox/LH2 propellant rocket stage. Loaded/empty mass 21,000/2,800 kg. Thrust 156.00 kN. Vacuum specific impulse 440 seconds. More...
  • CZ-NGLV-500 Lox/LH2 propellant rocket stage. Loaded/empty mass 175,000/17,000 kg. Thrust 1,399.99 kN. Vacuum specific impulse 432 seconds. 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 gimballed LOX /LH2 engines of 660 kN vacuum thrust each. The hydrogen tank is pressurised using hydrogen bled from the engine and helium is used to pressurise the oxygen tank. More...
  • CZ-NGLV-HO Lox/LH2 propellant rocket stage. Loaded/empty mass 26,000/3,100 kg. Thrust 156.00 kN. Vacuum specific impulse 448 seconds. 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. More...
  • CZ-YF-73 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,500/2,000 kg. Thrust 44.10 kN. Vacuum specific impulse 425 seconds. More...
  • DAC Helios ISI-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,333,000/120,000 kg. Thrust 49,807.00 kN. Vacuum specific impulse 455 seconds. More...
  • DAC Helios-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,660,000/149,000 kg. Thrust 52,926.00 kN. Vacuum specific impulse 410 seconds. More...
  • DC-I Lox/LH2 propellant rocket stage. Loaded/empty mass 470,000/36,000 kg. Thrust 6,860.00 kN. Vacuum specific impulse 425 seconds. More...
  • DC-X Lox/LH2 propellant rocket stage. Loaded/empty mass 16,320/7,200 kg. Thrust 262.80 kN. Vacuum specific impulse 373 seconds. More...
  • DC-Y Lox/LH2 propellant rocket stage. Loaded/empty mass 84,000/6,000 kg. Thrust 1,051.20 kN. Vacuum specific impulse 373 seconds. More...
  • Delta 3-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 19,300/2,476 kg. Thrust 110.03 kN. Vacuum specific impulse 462 seconds. 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. More...
  • Delta 4 - 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 24,170/2,850 kg. Thrust 110.05 kN. Vacuum specific impulse 462 seconds. Delta 3 second stage with hydrogen tank stretch. More...
  • Delta 4H - 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 30,710/3,490 kg. Thrust 110.05 kN. Vacuum specific impulse 462 seconds. Delta 4 second stage with hydrogen tank increased to 5.1 m diameter. More...
  • Delta RS-68 Lox/LH2 propellant rocket stage. Loaded/empty mass 226,400/26,760 kg. Thrust 3,312.76 kN. Vacuum specific impulse 420 seconds. Low cost expendable stage using lower performance engine. Used in Delta 4, Boeing EELV. Engine can be throttled to 60%. More...
  • Energia Core Lox/LH2 propellant rocket stage. Loaded/empty mass 905,000/85,000 kg. Thrust 7,848.12 kN. Vacuum specific impulse 453 seconds. More...
  • Energia EUS Lox/LH2 propellant rocket stage. Loaded/empty mass 77,000/7,000 kg. Thrust 1,962.03 kN. Vacuum specific impulse 455 seconds. More...
  • Energia M Lox/LH2 propellant rocket stage. Loaded/empty mass 272,000/28,000 kg. Thrust 1,960.00 kN. Vacuum specific impulse 455 seconds. More...
  • GSLV-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 14,600/2,200 kg. Thrust 75.05 kN. Vacuum specific impulse 460 seconds. The stage finally reached hardware status as a joint Russian-Indian development for India's GSLV booster. More...
  • H-2-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 98,100/11,900 kg. Thrust 1,078.00 kN. Vacuum specific impulse 446 seconds. More...
  • H-2A LRB Lox/LH2 propellant rocket stage. Loaded/empty mass 117,000/17,800 kg. Thrust 2,196.00 kN. Vacuum specific impulse 440 seconds. Two-engine version of H-2A-1 used as strap-on booster. More...
  • H-2A-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 113,600/13,600 kg. Thrust 1,098.00 kN. Vacuum specific impulse 440 seconds. Lower cost version of H-2 first stage. Can be throttled to 72% thrust. More...
  • Helios C-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 395,000/18,000 kg. Thrust 12,450.00 kN. Vacuum specific impulse 400 seconds. Booster stage with Lox tanks only to take nuclear second stage to stratosphere. More...
  • Helios A-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 200,000/10,000 kg. Thrust 6,660.00 kN. Vacuum specific impulse 400 seconds. Booster stage with Lox tanks only to take nuclear second stage to stratosphere. More...
  • Helios B-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 236,000/14,000 kg. Thrust 7,786.00 kN. Vacuum specific impulse 400 seconds. Booster stage with Lox tanks only to take nuclear second stage to stratosphere. More...
  • HIMES Lox/LH2 propellant rocket stage. Loaded/empty mass 13,600/3,100 kg. Thrust 274.00 kN. Vacuum specific impulse 452 seconds. More...
  • Hyperion Booster Lox/LH2 propellant rocket stage. Loaded/empty mass 394,625/18,144 kg. Thrust 13,700.00 kN. Vacuum specific impulse 457 seconds. More...
  • Hyperion SSTO Lox/LH2 propellant rocket stage. Loaded/empty mass 450,000/44,000 kg. Thrust 7,840.00 kN. Vacuum specific impulse 459 seconds. All values estimated based on drawing, statement that 5 x mass of SASSTO, payload performance, and 300 m/s sled velocity augmentation. More...
  • Interim HOTOL Lox/LH2 propellant rocket stage. Loaded/empty mass 250,000/33,100 kg. Thrust 7,840.00 kN. Vacuum specific impulse 455 seconds. More...
  • Jarvis-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 145,000/15,000 kg. Thrust 1,031.98 kN. Vacuum specific impulse 425 seconds. More...
  • LE-5EC Lox/LH2 propellant rocket stage. Loaded/empty mass 16,700/2,700 kg. Thrust 121.50 kN. Vacuum specific impulse 452 seconds. More...
  • LE-5 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,600/1,800 kg. Thrust 102.90 kN. Vacuum specific impulse 450 seconds. More...
  • LE-5B Lox/LH2 propellant rocket stage. Loaded/empty mass 19,600/3,000 kg. Thrust 137.00 kN. Vacuum specific impulse 447 seconds. Lower cost version of H-2 second stage. More...
  • Magnum Core Lox/LH2 propellant rocket stage. Loaded/empty mass 830,000/70,000 kg. Thrust 9,112.38 kN. Vacuum specific impulse 453 seconds. Alternative configurations used 2 to 3 RS-68 engines More...
  • McDonnell-Douglas ILRV Lox/LH2 propellant rocket stage. Loaded/empty mass 215,000/95,254 kg. Thrust 16,242.82 kN. High-pressure Lox/Lh2 engine. More...
  • MLLV Lox/LH2 propellant rocket stage. Loaded/empty mass 5,352,400/317,400 kg. Thrust 71,171.70 kN. Vacuum specific impulse 455 seconds. Boeing study, 1969. More...
  • Mustard 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 141,043/24,036 kg. Thrust 2,150.00 kN. Vacuum specific impulse 405 seconds. More...
  • Mustard 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 142,184/24,943 kg. Thrust 2,150.00 kN. Vacuum specific impulse 405 seconds. More...
  • N1 Block V-II Lox/LH2 propellant rocket stage. Loaded/empty mass 805,000/115,000 kg. Thrust 7,840.00 kN. Vacuum specific impulse 440 seconds. N1 improvement study, 1965. Lox/LH2 replacement for Block B second stage. More...
  • N1 Block V-III Lox/LH2 propellant rocket stage. Loaded/empty mass 325,000/35,000 kg. Thrust 2,350.00 kN. Vacuum specific impulse 440 seconds. 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. More...
  • N1 Block S Lox/LH2 propellant rocket stage. Loaded/empty mass 58,000/8,000 kg. Thrust 392.00 kN. Vacuum specific impulse 440 seconds. Designed 1965-1971 as replacement for N-1 Blok G. Cancelled in 1971 in favor of development of single stage, Block Sr. More...
  • N1 Block R Lox/LH2 propellant rocket stage. Loaded/empty mass 23,000/4,300 kg. Thrust 73.50 kN. Vacuum specific impulse 440 seconds. 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. More...
  • N1 Block Sr Lox/LH2 propellant rocket stage. Loaded/empty mass 77,900/11,500 kg. Thrust 147.88 kN. Vacuum specific impulse 441 seconds. 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 tonnes into lunar orbit or 20 tonnes into geosynch orbit. More...
  • NLS Core Lox/LH2 propellant rocket stage. Loaded/empty mass 815,732/44,757 kg. Thrust 7,150.00 kN. Vacuum specific impulse 430 seconds. More...
  • NLS HLV Lox/LH2 propellant rocket stage. Loaded/empty mass 833,732/62,757 kg. Thrust 14,310.00 kN. Vacuum specific impulse 425 seconds. More...
  • NLS Semistage Lox/LH2 propellant rocket stage. Loaded/empty mass 36,000/36,000 kg. Thrust 14,310.00 kN. Vacuum specific impulse 425 seconds. More...
  • Nova MM 1C-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,041,000/163,000 kg. Thrust 20,015.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been February 1973 More...
  • Nova MM S10E-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,302,000/639,000 kg. Thrust 160,672.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been November 1977. SSTO; expendable; payload 1,283,000 lbs. More...
  • Nova MM S10E-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,317,000/635,000 kg. Thrust 160,672.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been October 1977. SSTO; expendable. More...
  • Nova MM 34-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,126,000/544,000 kg. Thrust 30,685.00 kN. Vacuum specific impulse 439 seconds. Operational date would have been June 1976. Sustainer stage (required 4-engine booster stage). More...
  • Nova MM 34-0 Lox/LH2 propellant rocket stage. Loaded/empty mass 227,000/227,000 kg. Thrust 122,749.00 kN. Vacuum specific impulse 439 seconds. Operational date would have been June 1976. Booster stage (engines only). More...
  • Nova MM S10R-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,470,000/816,000 kg. Thrust 160,672.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been June 1978. SSTO; recoverable. More...
  • Nova MM 24G-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,043,000/136,000 kg. Thrust 10,228.00 kN. Vacuum specific impulse 451 seconds. Operational date would have been December 1974. More...
  • Nova 9L-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 363,000/36,000 kg. Thrust 5,334.00 kN. Vacuum specific impulse 420 seconds. Masses estimated based on total vehicle thrust, performance, and stage volumes. More...
  • Nova MM 1B-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,361,000/122,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been February 1973 More...
  • Nova MM 14B-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,721,000/245,000 kg. Thrust 26,683.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been February 1973 More...
  • Nova MM 33-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,489,000/626,000 kg. Thrust 156,876.00 kN. Vacuum specific impulse 443 seconds. Operational date would have been April 1975. SSTO - payload 1,042,000 lbs. More...
  • Nova MM S10R-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 10,502,000/839,000 kg. Thrust 160,672.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been July 1978. SSTO; recoverable; payload 842,000 lbs. More...
  • Nova MM T10EE-1-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,125,000/136,000 kg. Thrust 11,032.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been November 1976. Expendable stage. More...
  • Nova MM T10RE-1-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 4,979,000/454,000 kg. Thrust 96,379.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been January 1977. Recoverable stage. More...
  • Nova MM T10RR-2-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 9,089,000/680,000 kg. Thrust 176,686.00 kN. Vacuum specific impulse 304 seconds. Operational date would have been December 1976. Recoverable stage. More...
  • Nova MM T10RR-2-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,041,000/181,000 kg. Thrust 20,015.00 kN. Vacuum specific impulse 455 seconds. Operational date would have been December 1976. Recoverable stage. More...
  • Nova MM T10RR-3-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 5,517,000/680,000 kg. Thrust 109,000.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been July 1977. Recoverable stage. 10% plug nozzle. More...
  • Nova MM T10RR-3-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,229,000/136,000 kg. Thrust 12,052.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been July 1977. Recoverable stage. More...
  • Nova MM-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 5,040,000/317,000 kg. Thrust 95,261.00 kN. Vacuum specific impulse 451 seconds. Operational date would have been December 1974. More...
  • Nova NASA-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 227,000/21,000 kg. Thrust 2,667.00 kN. Vacuum specific impulse 420 seconds. More...
  • Nova MM 14A-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 3,401,000/295,000 kg. Thrust 33,352.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been April 1973 More...
  • Nova MM T10EE-1-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 4,944,000/317,000 kg. Thrust 96,379.00 kN. Vacuum specific impulse 454 seconds. Operational date would have been November 1976. Expendable stage. More...
  • Nova A-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 40,000/4,000 kg. Thrust 313.00 kN. Vacuum specific impulse 427 seconds. More...
  • Nova 4 J-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 771,000/63,000 kg. Thrust 3,559.00 kN. Vacuum specific impulse 420 seconds. Nova third stage. More...
  • Nova 59-4-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 168,000/18,000 kg. Thrust 1,421.00 kN. Vacuum specific impulse 420 seconds. Empty Mass Estimated More...
  • Nova 59-4-4 Lox/LH2 propellant rocket stage. Loaded/empty mass 68,000/9,000 kg. Thrust 353.00 kN. Vacuum specific impulse 420 seconds. Empty Mass Estimated More...
  • Nova 60-8-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 680,000/54,000 kg. Thrust 10,669.00 kN. Vacuum specific impulse 428 seconds. Mass estimated based on total LV weight. J-2-powered version of this stage also proposed. More...
  • Nova 60-8-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 227,000/23,000 kg. Thrust 892.00 kN. Vacuum specific impulse 425 seconds. Mass estimated based on total LV weight. More...
  • Nova A-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 403,000/29,000 kg. Thrust 5,334.00 kN. Vacuum specific impulse 420 seconds. More...
  • Nova 9L-4 Lox/LH2 propellant rocket stage. Loaded/empty mass 118,000/14,000 kg. Thrust 1,775.00 kN. Vacuum specific impulse 420 seconds. Masses estimated based on total vehicle thrust, performance, and stage volumes. More...
  • Nova GD-J-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,270,000/91,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 428 seconds. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova B-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 614,000/38,000 kg. Thrust 10,228.00 kN. Vacuum specific impulse 420 seconds. More...
  • Nova B-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 63,000/5,000 kg. Thrust 451.00 kN. Vacuum specific impulse 427 seconds. More...
  • Nova GD-F-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,497,000/91,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 428 seconds. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova GD-H-0 Lox/LH2 propellant rocket stage. Loaded/empty mass 295,000/295,000 kg. Thrust 109,402.00 kN. Vacuum specific impulse 410 seconds. 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. More...
  • Nova DAC 2-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 4,943,000/311,000 kg. Thrust 102,293.00 kN. Vacuum specific impulse 410 seconds. Operational date would have been July 1977. Recoverable stage. 10% plug nozzle. More...
  • Nova GD-E-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 4,535,000/363,000 kg. Thrust 26,683.00 kN. Vacuum specific impulse 428 seconds. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova GD-B-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 998,000/75,000 kg. Thrust 10,669.00 kN. Vacuum specific impulse 428 seconds. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova DAC-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,268,000/143,000 kg. Thrust 26,683.00 kN. Vacuum specific impulse 426 seconds. Operational date would have been July 1977. Recoverable stage. More...
  • Nova DAC ISI-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,696,000/107,000 kg. Thrust 33,754.00 kN. Vacuum specific impulse 455 seconds. Operational date would have been July 1977. Recoverable stage. More...
  • Nova DAC ISI-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 3,084,000/142,000 kg. Thrust 70,146.00 kN. Vacuum specific impulse 350 seconds. Operational date would have been July 1977. Recoverable stage. 10% plug nozzle. More...
  • Nova DAC 2-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,633,000/113,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 426 seconds. Operational date would have been July 1977. Recoverable stage. More...
  • Nova GD-H-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 8,526,000/172,000 kg. Thrust 30,685.00 kN. Vacuum specific impulse 428 seconds. More...
  • OOST Lox/LH2 propellant rocket stage. Loaded/empty mass 7,982,000/431,000 kg. Thrust 123,191.00 kN. Vacuum specific impulse 410 seconds. More...
  • OOST ISI Lox/LH2 propellant rocket stage. Loaded/empty mass 5,125,000/292,000 kg. Thrust 85,386.00 kN. Vacuum specific impulse 455 seconds. More...
  • Pegasus Tanks xLox/LH2 propellant rocket drop tank. Loaded/empty mass 20,000/3,700 kg. Vacuum specific impulse 459 seconds.Four tanks jettisoned at 130 seconds after liftoff; two at 250 seconds, last two at orbital insertion, 360 seconds after liftoff. More...
  • Pegasus VTOVL Lox/LH2 propellant rocket stage. Loaded/empty mass 1,250,000/148,000 kg. Thrust 23,947.00 kN. Vacuum specific impulse 459 seconds. Empty mass includes 29,600 kg of propellants used for re-entry cooling of plug nozzle and rocket soft landing at landing field. More...
  • Project 921-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 57,000/7,000 kg. Thrust 490.00 kN. Vacuum specific impulse 440 seconds. Additionally 4 vernier Lox/LH2 engines with a total thrust of 4600 kgf and a storable engine package for stage propellant ullage and restart. More...
  • Proton KM-4 Lox/LH2 propellant rocket stage. Loaded/empty mass 19,500/3,000 kg. Thrust 73.58 kN. Vacuum specific impulse 461 seconds. Planned version for Proton. Never developed. More...
  • Rombus Lox/LH2 propellant rocket stage. Loaded/empty mass 5,102,041/306,175 kg. Thrust 101,900.00 kN. Vacuum specific impulse 455 seconds. 36 x plug-nozzle engines (20 atm chamber pressure, 7:1 mixture ratio) More...
  • Rombus Tank Lox/LH2 propellant rocket drop tank. Loaded/empty mass 107,501/18,143 kg. Vacuum specific impulse 455 seconds. 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. More...
  • ROOST Lox/LH2 propellant rocket stage. Loaded/empty mass 10,898,000/608,000 kg. Thrust 165,447.00 kN. Vacuum specific impulse 410 seconds. More...
  • ROOST ISI Lox/LH2 propellant rocket stage. Loaded/empty mass 6,218,000/435,000 kg. Thrust 101,842.00 kN. Vacuum specific impulse 455 seconds. More...
  • Sanger I-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 79,000/10,000 kg. Thrust 885.00 kN. Vacuum specific impulse 430 seconds. 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. More...
  • Sanger I-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 21,000/5,000 kg. Thrust 295.00 kN. Vacuum specific impulse 430 seconds. 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 tonne payload could be delivered into orbit. More...
  • Sanger II-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 112,000/32,600 kg. Thrust 1,280.00 kN. Vacuum specific impulse 490 seconds. 6000 kg to LEO More...
  • SASSTO Lox/LH2 propellant rocket stage. Loaded/empty mass 97,976/6,668 kg. Thrust 1,558.10 kN. Vacuum specific impulse 464 seconds. Recoverable S-IVB with plug nozzle engine. 36 x plug-nozzle engines (102 atm chamber pressure, 6:1 mixture ratio) More...
  • Saturn S-II-C3 Lox/LH2 propellant rocket stage. Loaded/empty mass 204,044/24,938 kg. Thrust 3,557.31 kN. Vacuum specific impulse 420 seconds. Version for Saturn C-3. More...
  • Saturn MS-II-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 600,788/59,844 kg. Thrust 7,003.48 kN. Vacuum specific impulse 451 seconds. 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. More...
  • Saturn MS-II-3B Lox/LH2 propellant rocket stage. Loaded/empty mass 643,200/53,500 kg. Thrust 12,455.40 kN. Vacuum specific impulse 447 seconds. S-II with 15.5 foot stretch, 1.29 million pounds propellant, 7 x 400,000 lb thrust toroidal engines. More...
  • Saturn MS-IVB-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 120,500/13,900 kg. Thrust 1,031.60 kN. Vacuum specific impulse 421 seconds. Marshall studies, 1965: S-IVB structurally strengthened to handle larger payloads, otherwise unchanged More...
  • Saturn MS-IVB-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 176,600/17,800 kg. Thrust 1,401.30 kN. Vacuum specific impulse 447 seconds. Douglas Studies, 1965: S-IVB with 315 k high pressure 3000 psia engine, 350,000 pounds propellant More...
  • Saturn MS-IVB-3B Lox/LH2 propellant rocket stage. Loaded/empty mass 179,200/20,400 kg. Thrust 1,778.90 kN. Vacuum specific impulse 447 seconds. S-IVB with 16.5 foot stretch, 350,000 pounds propellant, 1 400,000 pound thrust toroidal engine. More...
  • Saturn MS-IVB-4(S)B Lox/LH2 propellant rocket stage. Loaded/empty mass 118,400/14,100 kg. Thrust 1,031.60 kN. Vacuum specific impulse 421 seconds. Standard S-IVB but with structural strength increased from 78% to 217% depending on station, resulting in 11.8% increase in empty weight. More...
  • Saturn MS-IVB-x Lox/LH2 propellant rocket stage. Loaded/empty mass 195,000/36,300 kg. Thrust 912.00 kN. Vacuum specific impulse 421 seconds. Marshall studies, 1965: S-IVB structurally strengthened to handle larger payloads, otherwise unchanged More...
  • Saturn S-II Lox/LH2 propellant rocket stage. Loaded/empty mass 100,000/14,000 kg. Thrust 3,557.31 kN. Vacuum specific impulse 420 seconds. Early design version for use with Saturn I first stage. More...
  • Saturn MS-II-1-J-2T-250K Lox/LH2 propellant rocket stage. Loaded/empty mass 521,447/49,877 kg. Thrust 5,558.31 kN. Vacuum specific impulse 441 seconds. 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. More...
  • Saturn S-II-8 Lox/LH2 propellant rocket stage. Loaded/empty mass 770,835/63,480 kg. Thrust 8,265.26 kN. Vacuum specific impulse 425 seconds. Version for Saturn C-8. More...
  • Saturn MS-II-4(S)B Lox/LH2 propellant rocket stage. Loaded/empty mass 494,100/42,300 kg. Thrust 5,169.00 kN. Vacuum specific impulse 421 seconds. Standard S-II but with structural strength increased from 86% to 502% depending on station, resulting in 8.6% increase in empty weight. More...
  • Saturn S-IVC Lox/LH2 propellant rocket stage. Loaded/empty mass 204,044/22,671 kg. Thrust 1,031.98 kN. Vacuum specific impulse 425 seconds. More...
  • Saturn S-II-4 Lox/LH2 propellant rocket stage. Loaded/empty mass 294,731/24,938 kg. Thrust 3,557.31 kN. Vacuum specific impulse 420 seconds. Version for Saturn C-4. More...
  • Saturn II C-5A Lox/LH2 propellant rocket stage. Loaded/empty mass 384,057/31,740 kg. Thrust 4,446.65 kN. Vacuum specific impulse 420 seconds. Final common second stage design for Saturn C-3, C-4 and C-5 (November 1961). Developed into Saturn V second stage. More...
  • Saturn MS-II-1A Lox/LH2 propellant rocket stage. Loaded/empty mass 600,800/60,000 kg. Thrust 6,380.94 kN. Vacuum specific impulse 421 seconds. Basic Saturn II with 187 inch stretch of propellant tanks, 1.2 million pound propellant capacity, and 7 J-2 engines. More...
  • Saturn MS-II-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 521,447/49,877 kg. Thrust 5,165.79 kN. Vacuum specific impulse 425 seconds. Basic Saturn II with 41 inch stretch of hydrogen tank. More...
  • Saturn II Lox/LH2 propellant rocket stage. Loaded/empty mass 490,778/39,048 kg. Thrust 5,165.79 kN. Vacuum specific impulse 421 seconds. Configuration as flown. More...
  • Saturn IVB-A Lox/LH2 propellant rocket stage. Loaded/empty mass 118,800/12,900 kg. Thrust 956.10 kN. Vacuum specific impulse 421 seconds. Douglas Studies, 1965: S-IVB with 215k lbf J-1 (actual final model had 230k J-1) More...
  • Saturn MS-IVB-1A Lox/LH2 propellant rocket stage. Loaded/empty mass 179,200/20,400 kg. Thrust 1,031.60 kN. Vacuum specific impulse 421 seconds. S-IVB with 16.5 foot stretch, 350,000 pounds propellant, standard J-2 engine. More...
  • Saturn IVB C-5A Lox/LH2 propellant rocket stage. Loaded/empty mass 102,249/11,563 kg. Thrust 889.33 kN. Vacuum specific impulse 420 seconds. 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. More...
  • Saturn IVB C-3B Lox/LH2 propellant rocket stage. Loaded/empty mass 52,144/6,801 kg. Thrust 889.33 kN. Vacuum specific impulse 420 seconds. Final common third stage design for Saturn C-3B (November 1961). More...
  • Saturn IVB (S-V) Lox/LH2 propellant rocket stage. Loaded/empty mass 119,900/13,300 kg. Thrust 1,031.60 kN. Vacuum specific impulse 421 seconds. Saturn V version of S-IVB stage for use with upper stage. More...
  • Saturn IVB (S-IB) Lox/LH2 propellant rocket stage. Loaded/empty mass 118,800/12,900 kg. Thrust 1,031.60 kN. Vacuum specific impulse 421 seconds. Saturn IB version of S-IVB stage. Due to lower payload 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). More...
  • Saturn IVB Lox/LH2 propellant rocket stage. Loaded/empty mass 119,920/13,311 kg. Thrust 1,031.98 kN. Vacuum specific impulse 425 seconds. Configuration as flown on Saturn V. More...
  • Saturn IV Lox/LH2 propellant rocket stage. Loaded/empty mass 50,576/5,217 kg. Thrust 400.35 kN. Vacuum specific impulse 410 seconds. Configuration as flown. More...
  • Saturn II-SL Lox/LH2 propellant rocket stage. Loaded/empty mass 490,952/44,240 kg. Thrust 5,500.77 kN. Vacuum specific impulse 390 seconds. Saturn II modifed 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. More...
  • Saturn II-INT-17 Lox/LH2 propellant rocket stage. Loaded/empty mass 495,000/48,000 kg. Thrust 9,713.40 kN. Vacuum specific impulse 450 seconds. Saturn II modifed with reduced expansion ratio HG-3 high pressure engines for use a first stage (sea level launch). More...
  • Saturn MS-II-1-J-2T-200K Lox/LH2 propellant rocket stage. Loaded/empty mass 521,447/49,877 kg. Thrust 4,446.65 kN. Vacuum specific impulse 435 seconds. Basic Saturn II with 41 inch stretch of hydrogen tank, uprated J-2T 200k engines with 10 second increase in specific impulse. More...
  • Sea Dragon-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 5,900,000/530,000 kg. Thrust 62,270.00 kN. Vacuum specific impulse 320 seconds. Length with extendible nozzle deployed 87 m. Diameter of extended nozzle 30 m. Total mass, specific impulse estimated from booster performance figures. More...
  • Sea Horse-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 15,000/2,000 kg. Thrust 147.00 kN. Vacuum specific impulse 425 seconds. More...
  • Shuttle MDC-A-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,242,630/220,254 kg. Thrust 18,956.00 kN. Vacuum specific impulse 442 seconds. Delta winged configuration. More...
  • Shuttle Super Lightweight Tank Lox/LH2 propellant rocket stage. Loaded/empty mass 748,000/27,000 kg. Thrust 6,834.30 kN. Vacuum specific impulse 452.5 seconds. The Super Lightweight Tank used 2195 Aluminium-Lithium alloy as the main structural material in place of the 2219 aluminium 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. More...
  • Shuttle SERV-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,040,816/226,757 kg. Thrust 31,980.52 kN. Vacuum specific impulse 455 seconds. Single stage to orbit, ballistic reentry. More...
  • Shuttle R134G-2 Lox/LH2 propellant rocket stage. Loaded mass 380,000 kg. Thrust 4,804.13 kN. Vacuum specific impulse 459 seconds.Straight winged configuration. More...
  • Shuttle R134G-1 Lox/LH2 propellant rocket stage. Loaded mass 1,600,000 kg. Thrust 28,824.77 kN. Vacuum specific impulse 442 seconds.Delta winged configuration. More...
  • Shuttle H33-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,489,717/224,431 kg. Thrust 24,973.40 kN. Vacuum specific impulse 442 seconds. Swept wing configuration. More...
  • Shuttle R134C-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 424,449/138,910 kg. Thrust 4,804.13 kN. Vacuum specific impulse 459 seconds. Delta winged configuration. More...
  • Shuttle Tank Lox/LH2 propellant rocket drop tank. Loaded/empty mass 750,975/29,930 kg. Vacuum specific impulse 455 seconds. Original version. More...
  • Shuttle R134C-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,764,039/351,538 kg. Thrust 28,824.77 kN. Vacuum specific impulse 442 seconds. Delta winged configuration. More...
  • Shuttle Orbiter Lox/LH2 propellant rocket stage. Loaded/empty mass 99,318/99,117 kg. Thrust 6,834.30 kN. Vacuum specific impulse 455 seconds. More...
  • Shuttle NAR A-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 395,011/121,542 kg. Thrust 5,256.89 kN. Vacuum specific impulse 459 seconds. Faget Straight Wing Configuration - Cross Range 2,000 km More...
  • Shuttle NAR A-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,641,723/273,469 kg. Thrust 28,130.19 kN. Vacuum specific impulse 442 seconds. Faget Straight Wing Configuration More...
  • Shuttle MDC-A-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 335,601/73,442 kg. Thrust 3,070.40 kN. Vacuum specific impulse 459 seconds. HL-10 lifting body configuration More...
  • Shuttle FR-3-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 388,189/130,159 kg. Thrust 4,549.05 kN. Vacuum specific impulse 459 seconds. Trapezoidal lifting body configuration. Cross range 2419 km. More...
  • Shuttle HCR-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,634,467/304,535 kg. Thrust 29,135.63 kN. Vacuum specific impulse 442 seconds. Swept wing configuration. More...
  • Shuttle DC-3-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 799,537/131,519 kg. Thrust 10,290.00 kN. Vacuum specific impulse 442 seconds. Faget Configuration More...
  • Shuttle LS200-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,730,803/133,514 kg. Thrust 27,422.00 kN. Vacuum specific impulse 455 seconds. High-fineness lifting-body configuration. Cross Range 2,419 km More...
  • Shuttle FR-3-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,169,691/234,467 kg. Thrust 30,088.43 kN. Vacuum specific impulse 442 seconds. Trapezoidal lifting body configuration. More...
  • Shuttle H33-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 474,199/100,153 kg. Thrust 7,079.45 kN. Vacuum specific impulse 459 seconds. Delta wing configuration with drop tanks - Cross Range 1,774 km More...
  • Shuttle HCR-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 342,948/129,329 kg. Thrust 4,719.64 kN. Vacuum specific impulse 459 seconds. Delta winged configuration. More...
  • Shuttle LCR-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,512,381/286,621 kg. Thrust 27,054.52 kN. Vacuum specific impulse 442 seconds. Swept wingd configuration. More...
  • Shuttle LCR-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 322,449/120,816 kg. Thrust 4,719.64 kN. Vacuum specific impulse 459 seconds. Faget Straight Wing Configuration More...
  • Shuttle LRB Lox/LH2 propellant rocket stage. Loaded/empty mass 350,000/52,000 kg. Thrust 10,318.11 kN. Vacuum specific impulse 435 seconds. More...
  • Shuttle LS A-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,225,668/162,494 kg. Thrust 26,076.59 kN. Vacuum specific impulse 442 seconds. Delta winged configuration. More...
  • Shuttle LS A-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 406,344/104,891 kg. Thrust 6,823.56 kN. Vacuum specific impulse 459 seconds. High-fineness lifting-body configuration. Cross Range 2,419 km More...
  • Shuttle DC-3-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 199,238/54,422 kg. Thrust 2,940.00 kN. Vacuum specific impulse 459 seconds. Faget Configuration - Cross Range 323 km More...
  • SLS Stage A Lox/LH2 propellant rocket stage. Loaded/empty mass 59,000/6,000 kg. Thrust 889.33 kN. Vacuum specific impulse 424 seconds. Smallest Lox/LH2 stage planned for SLS series. Empty mass estimated. Sized for rail transport within USA. More...
  • SLS Stage B Lox/LH2 propellant rocket stage. Loaded/empty mass 160,000/11,200 kg. Thrust 1,778.65 kN. Vacuum specific impulse 424 seconds. Tranlunar injection stage for Project Lunex. Masses estimated based on optimum apportioning of B+C stage total masses. More...
  • SLS Stage C Lox/LH2 propellant rocket stage. Loaded/empty mass 825,000/58,000 kg. Thrust 10,673.32 kN. Vacuum specific impulse 428 seconds. 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. More...
  • Spacemaster-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 362,812/80,000 kg. Thrust 4,549.05 kN. Vacuum specific impulse 459 seconds. Delta Winged, Cross Range 2,742 km More...
  • Spacemaster-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,224,490/203,342 kg. Thrust 28,082.54 kN. Vacuum specific impulse 442 seconds. Unique Catamaran configuration. More...
  • Starlifter Lox/LH2 propellant rocket stage. Loaded/empty mass 42,630/19,955 kg. Thrust 5,217.00 kN. Vacuum specific impulse 455 seconds. More...
  • Titan 5 Lox/LH2 propellant rocket stage. Loaded/empty mass 500,000/45,000 kg. Thrust 4,457.10 kN. Vacuum specific impulse 425 seconds. Part of launch vehicle proposed by Martin as alternative to NLS. All figures estimated based on 1,000,000 lb thrust single engine. More...
  • Titan C-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 57,400/6,000 kg. Thrust 666.00 kN. Vacuum specific impulse 403 seconds. Engine developed 1958-1960, but launch vehicle cancelled 1961. More...
  • UR-700M-3 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,000,000/200,000 kg. Thrust 11,700.00 kN. Vacuum specific impulse 455 seconds. Total mass, length, estimated based on empty mass, total vehicle mass. Engine specific impulse estimated based on performance requirements. More...
  • Venturestar Lox/LH2 propellant rocket stage. Loaded/empty mass 991,000/89,300 kg. Thrust 15,413.89 kN. Vacuum specific impulse 455 seconds. More...
  • VTOHL 45t Lox/LH2 propellant rocket stage. Loaded/empty mass 1,158,192/96,650 kg. Thrust 18,643.47 kN. Vacuum specific impulse 467 seconds. More...
  • VTOHL 9t Lox/LH2 propellant rocket stage. Loaded/empty mass 687,503/71,901 kg. Thrust 11,186.09 kN. Vacuum specific impulse 467 seconds. More...
  • VTOVL 150t Lox/LH2 propellant rocket stage. Loaded/empty mass 4,093,761/419,154 kg. Thrust 63,387.80 kN. Vacuum specific impulse 485 seconds. More...
  • Vulkan 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 800,000/80,000 kg. Thrust 7,450.00 kN. Vacuum specific impulse 452 seconds. 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. More...
  • Vulkan Blok V Lox/LH2 propellant rocket stage. Loaded/empty mass 142,000/15,000 kg. Thrust 411.00 kN. Vacuum specific impulse 460 seconds. Upper stage design by KB Saturn for manned lunar expedition, large geosynchronous platform launch. More...
  • X-33 Lox/LH2 propellant rocket stage. Loaded/empty mass 123,800/28,600 kg. Thrust 2,284.00 kN. Vacuum specific impulse 439 seconds. More...

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