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Vulcain engine of Ariane 5 rocket
RL-10 is an early example of cryogenic rocket engine.

A cryogenic rocket engine is a rocket engine that uses a cryogenic fuel or oxidizer, that is, its fuel or oxidizer (or both) are gases liquefied and stored at very low temperatures.[1] Notably, these engines were one of the main factors of NASA's success in reaching the Moon by the Saturn V rocket.[1]

During World War II, when powerful rocket engines were first considered by the German, American and Soviet engineers independently, all discovered that rocket engines need high mass flow rate of both oxidizer and fuel to generate a sufficient thrust. At that time oxygen and low molecular weight hydrocarbons were used as oxidizer and fuel pair. At room temperature and pressure, both are in gaseous state. Hypothetically, if propellants had been stored as pressurized gases, the size and mass of fuel tanks themselves would severely decrease rocket efficiency. Therefore, to get the required mass flow rate, the only option was to cool the propellants down to cryogenic temperatures (below −183 °C [90 K], −253 °C [20 K]), converting them to liquid form. Hence, all cryogenic rocket engines are also, by definition, either liquid-propellant rocket engines or hybrid rocket engines.[2]

Various cryogenic fuel-oxidizer combinations have been tried, but the combination of liquid hydrogen (LH2) fuel and the liquid oxygen (LOX) oxidizer is one of the most widely used.[1][3] Both components are easily and cheaply available, and when burned have one of the highest enthalpy releases by combustion,[4] producing specific impulse up to 450 s (effective exhaust velocity 4.4 km/s).



The major components of a cryogenic rocket engine are the combustion chamber (thrust chamber), pyrotechnic initiator, fuel injector, fuel cryopumps, oxidizer cryopumps, gas turbine, cryo valves, regulators, the fuel tanks, and rocket engine nozzle. In terms of feeding propellants to the combustion chamber, cryogenic rocket engines (or, generally, all liquid-propellant engines) are either pressure-fed or pump-fed, and pump-fed engines work in either a gas-generator cycle, a staged-combustion cycle, or an expander cycle.

The cryopumps are always turbopumps powered by a flow of fuel through gas turbines. Looking through this aspect, engines can be differentiated into a main flow or a bypass flow configuration. In the main flow design, all the pumped fuel is fed through the gas turbines, and in the end injected to the combustion chamber. In the bypass configuration, the fuel flow is split; the main part goes directly to the combustion chamber to generate thrust, while only a small amount of the fuel goes to the turbine.[citation needed]

LOX+LH2 rocket engines by countryEdit

Comparison of Cryogenic rocket engine for first stageEdit

model SSME LE-7A RD-0120 Vulcain2 RS-68 YF-77
Country of origin   United States   Japan   Soviet Union   France   United States   People's Republic of China
Cycle Staged combustion Staged combustion Staged combustion Gas-generator Gas-generator Gas-generator
Length 4.24 m 3.7 m 4.55 m 3.00 m 5.20 m 4.20 m
Diameter 1.63 m 1.82 m 2.42 m 1.76 m 2.43 m -
Dry weight 3,177 kg 1,832 kg 3,449 kg 1,686 kg 6,696 kg 2,700 kg
Chamber pressure 18.9 MPa 12.0MPa 21.8 MPa 11.7 MPa 9.7 MPa 10.2 MPa
Isp (vac.) 453 sec 440 sec 454 sec 433 sec 409 sec 438 sec
Thrust (vac.) 2.278MN 1.098MN 1.961MN 1.120MN 3.37MN 673 kN
Thrust (SL) 1.817MN 0.87MN 1.517MN 0.800MN 2.949MN 550 kN
Used in Space Shuttle H-IIA
Energia Ariane 5 Delta IV Long March 5

Comparison of cryogenic rocket engine for upper stageEdit

  RL-10 HM7B Vinci KVD-1 CE-7.5 CE-20 YF-73 YF-75 YF-75D RD-0146 ES-702 ES-1001 LE-5 LE-5A LE-5B
Country of origin   United States   France   France   Soviet Union   India   India   People's Republic of China   People's Republic of China   People's Republic of China   Russia   Japan   Japan   Japan   Japan   Japan
Cycle Expander Gas-generator Expander Staged combustion Staged combustion Gas-generator Gas-generator Gas-generator Expander Expander Gas-generator Gas-generator Gas-generator Expander bleed cycle
(Nozzle Expander)
Expander bleed cycle
(Chamber Expander)
Thrust (vac.) 66.7 kN (15,000 lbf) 62.7 kN 180 kN 69.6 kN 73 kN 200 kN 44.15 kN 78.45 kN 88.26 kN 98.1 kN (22,054 lbf) 68.6 kN (7.0 tf)[5] 98 kN (10.0 tf)[6] 102.9 kN (10.5 tf) r121.5 kN (12.4 tf) 137.2 kN (14 tf)
Mixture ratio 5.5:1 or 5.88:1 5.0 5.8 5.05 5.0 5.2 6.0 5.2 6.0 5.5 5 5
Nozzle ratio 40 83.1 100 40 80 80 40 40 140 130 110
Isp (vac.) 433 444.2 465 462 454 443 420 438 442 463 425[7] 425[8] 450 452 447
Chamber pressure :MPa 2.35 3.5 6.1 5.6 5.8 6.0 2.59 3.68 7.74 2.45 3.51 3.65 3.98 3.58
LH2 TP rpm 90,000 42,000 65,000 125,000 41,000 46,310 50,000 51,000 52,000
LOX TP rpm 18,000 16,680 21,080 16,000 17,000 18,000
Length m 1.73 1.8 2.2~4.2 2.14 2.14 1.44 2.8 2.2 2.68 2.69 2.79
Dry weight kg 135 165 550 282 435 558 236 550 242 255.8 259.4 255 248 285


  1. ^ a b c Bilstein, Roger E. (1995). Stages to Saturn: A Technological History of the Apollo/Saturn Launch Vehicles (NASA SP-4206) (The NASA History Series). NASA History Office. pp. 89–91. ISBN 0-7881-8186-6.
  2. ^ Biblarz, Oscar; Sutton, George H. (2009). Rocket Propulsion Elements. New York: Wiley. p. 597. ISBN 0-470-08024-8.
  3. ^ The liquefaction temperature of oxygen is 89 kelvins and at this temperature it has a density of 1.14 kg/l, and for hydrogen it is 20 kelvins, just above absolute zero, and has a density of 0.07 kg/l.
  4. ^ Biswas, S. (2000). Cosmic perspectives in space physics. Bruxelles: Kluwer. p. 23. ISBN 0-7923-5813-9. "... [LH2+LOX] has almost the highest specific impulse."
  5. ^ without nozzle 48.52kN (4.9 tf)
  6. ^ without nozzle 66.64kN (6.8 tf)
  7. ^ without nozzle 286.8
  8. ^ without nozzle 291.6

External linksEdit