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Cryogenic rocket engine

From Wikipedia, the free encyclopedia
Vulcain engine of Ariane 5 rocket

A cryogenic rocket engine is a rocket engine that uses a cryogenic fuel and oxidizer; that is, both its fuel and oxidizer are gases which have been liquefied and are stored at very low temperatures.[1] These highly efficient engines were first flown on the US Atlas-Centaur and were one of the main factors of NASA's success in reaching the Moon by the Saturn V rocket.[1]

Rocket engines burning cryogenic propellants remain in use today on high performance upper stages and boosters. Upper stages are numerous. Boosters include ESA's Ariane 5, JAXA's H-II, ISRO's GSLV, LVM3, United States Delta IV and Space Launch System. The United States, Russia, Japan, India, France and China are the only countries that have operational cryogenic rocket engines.

Cryogenic propellants

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

Rocket engines need high mass flow rates of both oxidizer and fuel to generate useful thrust. Oxygen, the simplest and most common oxidizer, is in the gas phase at standard temperature and pressure, as is hydrogen, the simplest fuel. While it is possible to store propellants as pressurized gases, this would require large, heavy tanks that would make achieving orbital spaceflight difficult if not impossible. On the other hand, if the propellants are cooled sufficiently, they exist in the liquid phase at higher density and lower pressure, simplifying tankage. These cryogenic temperatures vary depending on the propellant, with liquid oxygen existing below −183 °C (−297.4 °F; 90.1 K) and liquid hydrogen below −253 °C (−423.4 °F; 20.1 K). Since one or more of the propellants is in the liquid phase, all cryogenic rocket engines are by definition liquid-propellant 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 in combustion,[4] producing a specific impulse of up to 450 s at an effective exhaust velocity of 4.4 kilometres per second (2.7 mi/s; Mach 13).

Components and combustion cycles

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The major components of a cryogenic rocket engine are the combustion chamber, pyrotechnic initiator, fuel injector, fuel and oxidizer turbopumps, cryo valves, regulators, the fuel tanks, and rocket engine nozzle. In terms of feeding propellants to the combustion chamber, cryogenic rocket engines are almost exclusively pump-fed. Pump-fed engines work in a gas-generator cycle, a staged-combustion cycle, or an expander cycle. Gas-generator engines tend to be used on booster engines due to their lower efficiency, staged-combustion engines can fill both roles at the cost of greater complexity, and expander engines are exclusively used on upper stages due to their low thrust.[citation needed]

LOX+LH2 rocket engines by country

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Chinese YF-77 engine used by Long March 5
Chinese YF-77 engine used by Long March 5

Currently, six countries have successfully developed and deployed cryogenic rocket engines:

Country Engine Cycle Use Status
 United States RL-10 Expander Upper stage Active
J-2 Gas-generator lower stage Retired
SSME (aka RS-25) Staged combustion Booster Active
RS-68 Gas-generator Booster Retired
BE-3 Combustion tap-off New Shepard Active
BE-7 Dual Expander Blue Moon (spacecraft) Active
J-2X Gas-generator Upper stage Developmental
 Russia RD-0120 Staged combustion Booster Retired
KVD-1 Staged combustion Upper stage Retired
RD-0146 Expander Upper stage Developmental
 France Vulcain Gas-generator Booster Active
HM7B Gas-generator Upper stage Active
Vinci Expander Upper stage Developmental
 India CE-7.5 Staged combustion Upper stage Active
CE-20 Gas-generator Upper stage Active
 China YF-73 Gas-generator Upper stage Retired
YF-75 Gas-generator Upper stage Active
YF-75D Expander cycle Upper stage Active
YF-77 Gas-generator Booster Active
 Japan LE-7 / 7A[5] Staged combustion Booster Active
LE-5 / 5A / 5B[6] Gas-generator(LE-5)
Expander bleed(5A/5B) 		Upper stage Active
LE-9[7] Expander bleed Booster Active

Comparison of first stage cryogenic rocket engines

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model SSME/RS-25 LE-7A RD-0120 Vulcain 2 RS-68 YF-77
Country of origin  United States  Japan  Soviet Union  France  United States  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 2.6 m
Diameter 1.63 m 1.82 m 2.42 m 1.76 m 2.43 m 1.5 m
Dry weight 3,177 kg 1,832 kg 3,449 kg 1,686 kg 6,696 kg 1,054 kg
Propellant LOX/LH2 LOX/LH2 LOX/LH2 LOX/LH2 LOX/LH2 LOX/LH2
Chamber pressure 18.9 MPa 12.0MPa 21.8 MPa 11.7 MPa 9.7 MPa 10.1 MPa
Isp (vac.) 453 sec 440 sec 454 sec 433 sec 409 sec 428 sec
Thrust (vac.) 2.278MN 1.098MN 1.961MN 1.120MN 3.37MN 0.7MN
Thrust (SL) 1.817MN 0.87MN 1.517MN 0.800MN 2.949MN 0.518MN
Used in Space Shuttle
Space Launch System		 H-IIA
H-IIB		 Energia Ariane 5 Delta IV Long March 5

Comparison of upper stage cryogenic rocket engines

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Specifications
　 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  China  China  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 186.36 kN 44.15 kN 83.585 kN 88.36 kN 98.1 kN (22,054 lbf) 68.6 kN (7.0 tf)[8] 98 kN (10.0 tf)[9] 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.) 433 444.2 465 462 454 442 420 438 442.6 463 425[10] 425[11] 450 452 447
Chamber pressure :MPa 2.35 3.5 6.1 5.6 5.8 6.0 2.59 3.68 4.1 5.9 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 135 165 550 282 435 558 236 245 265 242 255.8 259.4 255 248 285

References

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  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 978-0-470-08024-5.
  3. ^ The liquefaction temperature of oxygen is 89 kelvins, and at this temperature it has a density of 1.14 kg/L. For hydrogen it is 20 K, 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. ^ https://www.rocket.jaxa.jp/rocket/engine/le7/
  6. ^ https://www.rocket.jaxa.jp/rocket/engine/le5b/
  7. ^ https://www.rocket.jaxa.jp/rocket/engine/le9/
  8. ^ without nozzle 48.52kN (4.9 tf)
  9. ^ without nozzle 66.64kN (6.8 tf)
  10. ^ without nozzle 286.8
  11. ^ without nozzle 291.6
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