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. Notably, these engines were one of the main factors of NASA's success in reaching the Moon by the Saturn V rocket.
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.
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. Both components are easily and cheaply available, and when burned have one of the highest enthalpy releases by combustion, 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 at 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.
LOX+LH2 rocket engines by country
Currently, six countries have successfully developed and deployed cryogenic rocket engines:
Comparison of Cryogenic rocket engine for first stage
|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 stage
|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
|Expander bleed cycle
|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)||98 kN (10.0 tf)||102.9 kN (10.5 tf)||r121.5 kN (12.4 tf)||137.2 kN (14 tf)|
|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|
|Dry weight kg||135||165||550||282||435||558||236||550||242||255.8||259.4||255||248||285|
- Bilstein, Roger E. (1996). 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.
- Biblarz, Oscar; Sutton, George H. (2009). Rocket Propulsion Elements. New York: Wiley. p. 597. ISBN 0-470-08024-8.
- 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.
- 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."
- without nozzle 48.52kN (4.9 tf)
- without nozzle 66.64kN (6.8 tf)
- without nozzle 286.8
- without nozzle 291.6