Cryogenic rocket engine: Difference between revisions

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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).

Construction

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.^{[citation needed]}

LOX+LH2 rocket engines by country

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

Country	Engine	Cycle
United States	SSME	Staged combustion
	J-2	Gas-generator
	RL-10	Expander
	RS-68	Gas-generator
	BE-3	Combustion tap-off
Russia	RD-0120	Staged combustion
	RD-0146	Expander
	KVD-1	Staged combustion
France	Vulcain	Gas-generator
	HM7B	Gas-generator
	Vinci	Expander
India	CE-7.5	Staged combustion
India	CE-20	Gas-generator
People's Republic of China	YF-50t	Staged combustion
	YF-73	Gas-generator
	YF-75	Gas-generator
	YF-75D	Expander cycle
	YF-77	Gas-generator
Japan	LE-7 / 7A	Staged combustion
Japan	LE-5 / 5A / 5B	Gas-generator(LE-5) Expander(5A/5B)

Comparison of Cryogenic rocket engine for first stage

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
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	10.2 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 H-IIB	Energia	Ariane 5	Delta IV	Long March 5

Comparison of Cryogenic rocket engine for upper stage

**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	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		5.0	5.8				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
I_sp (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

References

^ ^a ^b ^c 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

External links

[saturn-1] 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.

[isbn0-470-08024-8-2] 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.

[isbn0-7923-5813-9-4] 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

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 ==External links==
-*[http://www.astronautix.com/engines/rl10b2.htm  USA's Cryogenic Rocket engine RL10B-2]
+*[https://web.archive.org/web/20120204144940/http://www.astronautix.com/engines/rl10b2.htm  USA's Cryogenic Rocket engine RL10B-2]
 *[http://www.lpre.de/energomash/RD-170/index.htm Russian Cryogenic Rocket Engines]