Liquid rocket propellant

Main article: Liquid-propellant rocket

The highest specific impulse chemical rockets use liquid propellants (liquid-propellant rockets). They can consist of a single chemical (a monopropellant) or a mix of two chemicals, called bipropellants. Bipropellants can further be divided into two categories; hypergolic propellants, which ignite when the fuel and oxidizer make contact, and non-hypergolic propellants which require an ignition source. ^[1]

About 170 different propellants made of liquid fuel have been tested, excluding minor changes to a specific propellant such as propellant additives, corrosion inhibitors, or stabilizers. In the U.S. alone at least 25 different propellant combinations have been flown. ^[2]

Many factors go into choosing a propellant for a liquid-propellant rocket engine. The primary factors include ease of operation, cost, hazards/environment and performance.^{[ citation needed ]}

History

Development in early 20th century

Robert H. Goddard on March 16, 1926, holding the launching frame of the first liquid-fueled rocket

Konstantin Tsiolkovsky proposed the use of liquid propellants in 1903, in his article Exploration of Outer Space by Means of Rocket Devices. ^{[3] [4]}

On March 16, 1926, Robert H. Goddard used liquid oxygen (LOX) and gasoline as propellants for his first partially successful liquid-propellant rocket launch. Both propellants are readily available, cheap and highly energetic. Oxygen is a moderate cryogen as air will not liquefy against a liquid oxygen tank, so it is possible to store LOX briefly in a rocket without excessive insulation. ^{[ clarification needed ]}

In Germany, engineers and scientists began building and testing liquid propulsion rockets in the late 1920s. ^[5] According to Max Valier, two liquid-propellant Opel RAK rockets were launched in Rüsselsheim on April 10 and April 12, 1929. ^[6]

World War II era

Germany had very active rocket development before and during World War II, both for the strategic V-2 rocket and other missiles. The V-2 used an alcohol/LOX liquid-propellant engine, with hydrogen peroxide to drive the fuel pumps. ^{[7] : 9} The alcohol was mixed with water for engine cooling. Both Germany and the United States developed reusable liquid-propellant rocket engines that used a storeable liquid oxidizer with much greater density than LOX and a liquid fuel that ignited spontaneously on contact with the high density oxidizer.^{[ citation needed ]}

The major manufacturer of German rocket engines for military use, the HWK firm, ^[8] manufactured the RLM-numbered 109-500-designation series of rocket engine systems, and either used hydrogen peroxide as a monopropellant for Starthilfe rocket-propulsive assisted takeoff needs; ^[9] or as a form of thrust for MCLOS-guided air-sea glide bombs; ^[10] and used in a bipropellant combination of the same oxidizer with a fuel mixture of hydrazine hydrate and methyl alcohol for rocket engine systems intended for manned combat aircraft propulsion purposes. ^[11]

The U.S. engine designs were fueled with the bipropellant combination of nitric acid as the oxidizer; and aniline as the fuel. Both engines were used to power aircraft, the Me 163 Komet interceptor in the case of the Walter 509-series German engine designs, and RATO units from both nations (as with the Starthilfe system for the Luftwaffe) to assist take-off of aircraft, which comprised the primary purpose for the case of the U.S. liquid-fueled rocket engine technology - much of it coming from the mind of U.S. Navy officer Robert Truax. ^[12]

1950s and 1960s

During the 1950s and 1960s there was a great burst of activity by propellant chemists to find high-energy liquid and solid propellants better suited to the military. Large strategic missiles need to sit in land-based or submarine-based silos for many years, able to launch at a moment's notice. Propellants requiring continuous refrigeration, which cause their rockets to grow ever-thicker blankets of ice, were not practical. As the military was willing to handle and use hazardous materials, a great number of dangerous chemicals were brewed up in large batches, most of which wound up being deemed unsuitable for operational systems.^{[ citation needed ]} In the case of nitric acid, the acid itself (HNO
_³) was unstable, and corroded most metals, making it difficult to store. The addition of a modest amount of nitrogen tetroxide, N
_²O
_⁴, turned the mixture red and kept it from changing composition, but left the problem that nitric acid corrodes containers it is placed in, releasing gases that can build up pressure in the process. The breakthrough was the addition of a little hydrogen fluoride (HF), which forms a self-sealing metal fluoride on the interior of tank walls that Inhibited Red Fuming Nitric Acid. This made "IRFNA" storeable.

Propellant combinations based on IRFNA or pure N
_²O
_⁴ as oxidizer and kerosene or hypergolic (self igniting) aniline, hydrazine or unsymmetrical dimethylhydrazine (UDMH) as fuel were then adopted in the United States and the Soviet Union for use in strategic and tactical missiles. The self-igniting storeable liquid bi-propellants have somewhat lower specific impulse than LOX/kerosene but have higher density so a greater mass of propellant can be placed in the same sized tanks. Gasoline was replaced by different hydrocarbon fuels, ^[7] for example RP-1  – a highly refined grade of kerosene. This combination is quite practical for rockets that need not be stored.

Kerosene

The V-2 rockets developed by Nazi Germany used LOX and ethyl alcohol. One of the main advantages of alcohol was its water content, which provided cooling in larger rocket engines. Petroleum-based fuels offered more power than alcohol, but standard gasoline and kerosene left too much soot and combustion by-products that could clog engine plumbing. In addition, they lacked the cooling properties of ethyl alcohol.

During the early 1950s, the chemical industry in the US was assigned the task of formulating an improved petroleum-based rocket propellant which would not leave residue behind and also ensure that the engines would remain cool. The result was RP-1, the specifications of which were finalized by 1954. A highly refined form of jet fuel, RP-1 burned much more cleanly than conventional petroleum fuels and also posed less of a danger to ground personnel from explosive vapours. It became the propellant for most of the early American rockets and ballistic missiles such as the Atlas, Titan I, and Thor. The Soviets quickly adopted RP-1 for their R-7 missile, but the majority of Soviet launch vehicles ultimately used storable hypergolic propellants. As of 2017 ^[update] , it is used in the first stages of many orbital launchers.

Hydrogen

Many early rocket theorists believed that hydrogen would be a marvelous propellant, since it gives the highest specific impulse. It is also considered the cleanest when oxidized with oxygen because the only by-product is water. Steam reforming of natural gas is the most common method of producing commercial bulk hydrogen at about 95% of the world production ^{[13] [14]} of 500 billion m³ in 1998. ^[15] At high temperatures (700–1100 °C) and in the presence of a metal-based catalyst (nickel), steam reacts with methane to yield carbon monoxide and hydrogen.

Hydrogen is very bulky compared to other fuels; it is typically stored as a cryogenic liquid, a technique mastered in the early 1950s as part of the hydrogen bomb development program at Los Alamos. Liquid hydrogen can be stored and transported without boil-off, by using helium as a cooling refrigerant, since helium has an even lower boiling point than hydrogen. Hydrogen is lost via venting to the atmosphere only after it is loaded onto a launch vehicle, where there is no refrigeration. ^[16]

In the late 1950s and early 1960s it was adopted for hydrogen-fuelled stages such as Centaur and Saturn upper stages.^{[ citation needed ]} Hydrogen has low density even as a liquid, requiring large tanks and pumps; maintaining the necessary extreme cold requires tank insulation. This extra weight reduces the mass fraction of the stage or requires extraordinary measures such as pressure stabilization of the tanks to reduce weight. (Pressure stabilized tanks support most of the loads with internal pressure rather than with solid structures, employing primarily the tensile strength of the tank material.^{[ citation needed ]})

The Soviet rocket programme, in part due to a lack of technical capability, did not use liquid hydrogen as a propellant until the Energia core stage in the 1980s.^{[ citation needed ]}

Upper stage use

The liquid-rocket engine bipropellant liquid oxygen and hydrogen offers the highest specific impulse for conventional rockets. This extra performance largely offsets the disadvantage of low density, which requires larger fuel tanks. However, a small increase in specific impulse in an upper stage application can give a significant increase in payload-to-orbit mass. ^[17]

Comparison to kerosene

Launch pad fires due to spilled kerosene are more damaging than hydrogen fires, for two main reasons:

• Kerosene burns about 20% hotter in absolute temperature than hydrogen.
• Hydrogen's buoyancy. Since hydrogen is a deep cryogen it boils quickly and rises, due to its very low density as a gas. Even when hydrogen burns, the gaseous H
  _²O that is formed has a molecular weight of only 18  AMU compared to 29.9  AMU for air, so it also rises quickly. Spilled kerosene fuel, on the other hand, falls to the ground and if ignited can burn for hours when spilled in large quantities.

Kerosene fires unavoidably cause extensive heat damage that requires time-consuming repairs and rebuilding. This is most frequently experienced by test stand crews involved with firings of large, unproven rocket engines.

Hydrogen-fuelled engines require special design, such as running propellant lines horizontally, so that no "traps" form in the lines, which would cause pipe ruptures due to boiling in confined spaces. (The same caution applies to other cryogens such as liquid oxygen and liquid natural gas (LNG).) Liquid hydrogen fuel has an excellent safety record and performance that is well above all other practical chemical rocket propellants.

Lithium and fluorine

The highest-specific-impulse chemistry ever test-fired in a rocket engine was lithium and fluorine, with hydrogen added to improve the exhaust thermodynamics (all propellants had to be kept in their own tanks, making this a tripropellant). The combination delivered 542 s specific impulse in vacuum, equivalent to an exhaust velocity of 5320 m/s. The impracticality of this chemistry highlights why exotic propellants are not actually used: to make all three components liquids, the hydrogen must be kept below −252 °C (just 21 K), and the lithium must be kept above 180 °C (453 K). Lithium and fluorine are both extremely corrosive. Lithium ignites on contact with air, and fluorine ignites most fuels on contact, including hydrogen. Fluorine and the hydrogen fluoride (HF) in the exhaust are very toxic, which makes working around the launch pad difficult, damages the environment, and makes getting a launch license more difficult. Both lithium and fluorine are expensive compared to most rocket propellants. This combination has therefore never flown. ^[18]

During the 1950s, the Department of Defense proposed lithium/fluorine as ballistic-missile propellants. A 1954 accident at a chemical works that released a cloud of fluorine into the atmosphere convinced them to use LOX/RP-1 instead.^{[ citation needed ]}

Methane

Using liquid methane and liquid oxygen as propellants is sometimes called methalox propulsion. ^[19] Liquid methane has a lower specific impulse than liquid hydrogen, but is easier to store due to its higher boiling point and density, as well as its lack of hydrogen embrittlement. It also leaves less residue in the engines compared to kerosene, which is beneficial for reusability. ^{[20] [21]} In addition, it is expected that its production on Mars will be possible via the Sabatier reaction. In NASA's Mars Design Reference Mission 5.0 documents (between 2009 and 2012), liquid methane/LOX (methalox) was the chosen propellant mixture for the lander module.

Due to the advantages methane fuel offers, some private space launch providers aimed to develop methane-based launch systems during the 2010s and 2020s. The competition between countries was dubbed the Methalox Race to Orbit, with the LandSpace's Zhuque-2 methalox rocket becoming the first to reach orbit. ^{[22] [23] [24]}

As of January 2024 ^[update] , two methane-fueled rockets have reached orbit. Several others are in development and two orbital launch attempts failed:

• Zhuque-2 successfully reached orbit on its second flight on 12 July 2023, becoming the first methane-fueled rocket to do so. ^[25] It had failed to reach orbit on its maiden flight on 14 December 2022. The rocket, developed by LandSpace, uses the TQ-12 engine.
• Vulcan Centaur successfully reached orbit on its first try, called Cert-1, on 8 January 2024. ^[26] The rocket, developed by United Launch Alliance, uses the Blue Origin's BE-4 engine, though the second stage uses the hydrolox RL10.
• Terran 1 had a failed orbital launch attempt on its maiden flight on 22 March 2023. The rocket, developed by Relativity Space, uses the Aeon 1 engine.
• Starship achieved a transatmospheric orbit on its third flight on 14 March 2024, ^[27] after two failed attempts. The rocket, developed by SpaceX, uses the Raptor engine.

SpaceX developed the Raptor engine for its Starship super-heavy-lift launch vehicle. ^[28] It has been used in test flights since 2019. SpaceX had previously used only RP-1/LOX in their engines.

Blue Origin developed the BE-4 LOX/LNG engine for their New Glenn and the United Launch Alliance Vulcan Centaur. The BE-4 provides 2,400 kN (550,000 lbf) of thrust. Two flight engines had been delivered to ULA by mid 2023.

In July 2014, Firefly Space Systems announced plans to use methane fuel for their small satellite launch vehicle, Firefly Alpha with an aerospike engine design. ^[29]

ESA is developing a 980kN methalox Prometheus rocket engine which was test fired in 2023. ^[30]

Monopropellants

High-test peroxide

High test peroxide is concentrated Hydrogen peroxide, with around 2% to 30% water. It decomposes to steam and oxygen when passed over a catalyst. This was historically used for reaction control systems, due to being easily storable. It is often used to drive Turbopumps, being used on the V2 rocket, and modern Soyuz.

Hydrazine

decomposes energetically to nitrogen, hydrogen, and ammonia (2N₂H₄ → N₂+H₂+2NH₃) and is the most widely used in space vehicles. (Non-oxidized ammonia decomposition is endothermic and would decrease performance).

Nitrous oxide

decomposes to nitrogen and oxygen.

Steam

when externally heated gives a reasonably modest I_sp of up to 190 seconds, depending on material corrosion and thermal limits.

Present use

As of June 2024 ^[update] , liquid fuel combinations in common use:

Kerosene (RP-1) / liquid oxygen (LOX)

Used for the lower stages of the Soyuz-2 boosters, the first stage of Atlas V, and both stages of Electron, Falcon 9, Falcon Heavy, and Firefly Alpha.

Liquid hydrogen (LH) / LOX

Used in the stages of the Space Launch System, New Shepard, H-IIB, GSLV and Centaur.

Liquid methane (LNG) / LOX

Used in both stages of Zhuque-2, Starship (doing nearly orbital test flights), and the first stage of the Vulcan Centaur.

Unsymmetrical dimethylhydrazine (UDMH) or monomethylhydrazine (MMH) / dinitrogen tetroxide (NTO or N
_²O
_⁴)

Used in three first stages of the Russian Proton booster, Indian Vikas engine for PSLV and GSLV rockets, most Chinese boosters, a number of military, orbital and deep space rockets, as this fuel combination is hypergolic and storable for long periods at reasonable temperatures and pressures.

Hydrazine (N
_²H
_⁴)

Used in deep space missions because it is storable and hypergolic, and can be used as a monopropellant with a catalyst.

Aerozine-50 (50/50 hydrazine and UDMH)

Used in deep space missions because it is storable and hypergolic, and can be used as a monopropellant with a catalyst.

Table

To approximate I_sp at other chamber pressures^{[ clarification needed ]}

Absolute pressure kPa ; atm (psi)	Multiply by
6,895 kPa; 68.05 atm (1,000 psi)	1.00
6,205 kPa; 61.24 atm (900 psi)	0.99
5,516 kPa; 54.44 atm (800 psi)	0.98
4,826 kPa; 47.63 atm (700 psi)	0.97
4,137 kPa; 40.83 atm (600 psi)	0.95
3,447 kPa; 34.02 atm (500 psi)	0.93
2,758 kPa; 27.22 atm (400 psi)	0.91
2,068 kPa; 20.41 atm (300 psi)	0.88

The table uses data from the JANNAF thermochemical tables (Joint Army-Navy-NASA-Air Force (JANNAF) Interagency Propulsion Committee) throughout, with best-possible specific impulse calculated by Rocketdyne under the assumptions of adiabatic combustion, isentropic expansion, one-dimensional expansion and shifting equilibrium. ^[31] Some units have been converted to metric, but pressures have not.

Definitions

V_e

Average exhaust velocity, m/s. The same measure as specific impulse in different units, numerically equal to specific impulse in N·s/kg.

r

Mixture ratio: mass oxidizer / mass fuel

T_c

Chamber temperature, °C

d

Bulk density of fuel and oxidizer, g/cm³

C*

Characteristic velocity, m/s. Equal to chamber pressure multiplied by throat area, divided by mass flow rate. Used to check experimental rocket's combustion efficiency.

Bipropellants

Oxidizer	Fuel	Comment	Optimal expansion from 68.05 atm to[ citation needed ]
			1 atm					0 atm, vacuum (nozzle area ratio 40:1)
			Ve	r	Tc	d	C*	Ve	r	Tc	d	C*
LOX	H 2	Hydrolox. Common.	3816	4.13	2740	0.29	2416	4462	4.83	2978	0.32	2386
	H 2:Be 49:51		4498	0.87	2558	0.23	2833	5295	0.91	2589	0.24	2850
	CH 4 (methane)	Methalox . Many engines under development in the 2010s.	3034	3.21	3260	0.82	1857	3615	3.45	3290	0.83	1838
	C2H6		3006	2.89	3320	0.90	1840	3584	3.10	3351	0.91	1825
	C2H4		3053	2.38	3486	0.88	1875	3635	2.59	3521	0.89	1855
	RP-1 (kerosene)	Kerolox. Common.	2941	2.58	3403	1.03	1799	3510	2.77	3428	1.03	1783
	N2H4		3065	0.92	3132	1.07	1892	3460	0.98	3146	1.07	1878
	B5H9		3124	2.12	3834	0.92	1895	3758	2.16	3863	0.92	1894
	B2H6		3351	1.96	3489	0.74	2041	4016	2.06	3563	0.75	2039
	CH4:H2 92.6:7.4		3126	3.36	3245	0.71	1920	3719	3.63	3287	0.72	1897
GOX	GH2	Gaseous form	3997	3.29	2576	—	2550	4485	3.92	2862	—	2519
F2	H2		4036	7.94	3689	0.46	2556	4697	9.74	3985	0.52	2530
	H2:Li 65.2:34.0		4256	0.96	1830	0.19	2680
	H2:Li 60.7:39.3							5050	1.08	1974	0.21	2656
	CH4		3414	4.53	3918	1.03	2068	4075	4.74	3933	1.04	2064
	C2H6		3335	3.68	3914	1.09	2019	3987	3.78	3923	1.10	2014
	MMH		3413	2.39	4074	1.24	2063	4071	2.47	4091	1.24	1987
	N2H4		3580	2.32	4461	1.31	2219	4215	2.37	4468	1.31	2122
	NH3		3531	3.32	4337	1.12	2194	4143	3.35	4341	1.12	2193
	B5H9		3502	5.14	5050	1.23	2147	4191	5.58	5083	1.25	2140
OF2	H2		4014	5.92	3311	0.39	2542	4679	7.37	3587	0.44	2499
	CH4		3485	4.94	4157	1.06	2160	4131	5.58	4207	1.09	2139
	C2H6		3511	3.87	4539	1.13	2176	4137	3.86	4538	1.13	2176
	RP-1		3424	3.87	4436	1.28	2132	4021	3.85	4432	1.28	2130
	MMH		3427	2.28	4075	1.24	2119	4067	2.58	4133	1.26	2106
	N2H4		3381	1.51	3769	1.26	2087	4008	1.65	3814	1.27	2081
	MMH:N2H4:H2O 50.5:29.8:19.7		3286	1.75	3726	1.24	2025	3908	1.92	3769	1.25	2018
	B2H6		3653	3.95	4479	1.01	2244	4367	3.98	4486	1.02	2167
	B5H9		3539	4.16	4825	1.20	2163	4239	4.30	4844	1.21	2161
F2:O2 30:70	H2		3871	4.80	2954	0.32	2453	4520	5.70	3195	0.36	2417
	RP-1		3103	3.01	3665	1.09	1908	3697	3.30	3692	1.10	1889
F2:O2 70:30	RP-1		3377	3.84	4361	1.20	2106	3955	3.84	4361	1.20	2104
F2:O2 87.8:12.2	MMH		3525	2.82	4454	1.24	2191	4148	2.83	4453	1.23	2186
Oxidizer	Fuel	Comment	Ve	r	Tc	d	C*	Ve	r	Tc	d	C*
N2F4	CH4		3127	6.44	3705	1.15	1917	3692	6.51	3707	1.15	1915
	C2H4		3035	3.67	3741	1.13	1844	3612	3.71	3743	1.14	1843
	MMH		3163	3.35	3819	1.32	1928	3730	3.39	3823	1.32	1926
	N2H4		3283	3.22	4214	1.38	2059	3827	3.25	4216	1.38	2058
	NH3		3204	4.58	4062	1.22	2020	3723	4.58	4062	1.22	2021
	B5H9		3259	7.76	4791	1.34	1997	3898	8.31	4803	1.35	1992
ClF5	MMH		2962	2.82	3577	1.40	1837	3488	2.83	3579	1.40	1837
	N2H4		3069	2.66	3894	1.47	1935	3580	2.71	3905	1.47	1934
	MMH:N2H4 86:14		2971	2.78	3575	1.41	1844	3498	2.81	3579	1.41	1844
	MMH:N2H4:N2H5NO3 55:26:19		2989	2.46	3717	1.46	1864	3500	2.49	3722	1.46	1863
ClF3	MMH:N2H4:N2H5NO355:26:19	Hypergolic	2789	2.97	3407	1.42	1739	3274	3.01	3413	1.42	1739
	N2H4	Hypergolic	2885	2.81	3650	1.49	1824	3356	2.89	3666	1.50	1822
N2O4	MMH	Hypergolic, common	2827	2.17	3122	1.19	1745	3347	2.37	3125	1.20	1724
	MMH:Be 76.6:29.4		3106	0.99	3193	1.17	1858	3720	1.10	3451	1.24	1849
	MMH:Al 63:27		2891	0.85	3294	1.27	1785
	MMH:Al 58:42							3460	0.87	3450	1.31	1771
	N2H4	Hypergolic, common	2862	1.36	2992	1.21	1781	3369	1.42	2993	1.22	1770
	N2H4:UDMH 50:50	Hypergolic, common	2831	1.98	3095	1.12	1747	3349	2.15	3096	1.20	1731
	N2H4:Be 80:20		3209	0.51	3038	1.20	1918
	N2H4:Be 76.6:23.4							3849	0.60	3230	1.22	1913
	B5H9		2927	3.18	3678	1.11	1782	3513	3.26	3706	1.11	1781
NO:N2O4 25:75	MMH		2839	2.28	3153	1.17	1753	3360	2.50	3158	1.18	1732
	N2H4:Be 76.6:23.4		2872	1.43	3023	1.19	1787	3381	1.51	3026	1.20	1775
IRFNA IIIa	UDMH:DETA 60:40	Hypergolic	2638	3.26	2848	1.30	1627	3123	3.41	2839	1.31	1617
	MMH	Hypergolic	2690	2.59	2849	1.27	1665	3178	2.71	2841	1.28	1655
	UDMH	Hypergolic	2668	3.13	2874	1.26	1648	3157	3.31	2864	1.27	1634
IRFNA IV HDA	UDMH:DETA 60:40	Hypergolic	2689	3.06	2903	1.32	1656	3187	3.25	2951	1.33	1641
	MMH	Hypergolic	2742	2.43	2953	1.29	1696	3242	2.58	2947	1.31	1680
	UDMH	Hypergolic	2719	2.95	2983	1.28	1676	3220	3.12	2977	1.29	1662
H2O2	MMH		2790	3.46	2720	1.24	1726	3301	3.69	2707	1.24	1714
	N2H4		2810	2.05	2651	1.24	1751	3308	2.12	2645	1.25	1744
	N2H4:Be 74.5:25.5		3289	0.48	2915	1.21	1943	3954	0.57	3098	1.24	1940
	B5H9		3016	2.20	2667	1.02	1828	3642	2.09	2597	1.01	1817
Oxidizer	Fuel	Comment	Ve	r	Tc	d	C*	Ve	r	Tc	d	C*

Definitions of some of the mixtures:

IRFNA IIIa

83.4% HNO₃, 14% NO₂, 2% H₂O, 0.6% HF

IRFNA IV HDA

54.3% HNO₃, 44% NO₂, 1% H₂O, 0.7% HF

RP-1

See MIL-P-25576C, basically kerosene (approximately C
_¹⁰H
_¹⁸)

MMH monomethylhydrazine

CH
_³NHNH
_²

Has not all data for CO/O₂, purposed for NASA for Martian-based rockets, only a specific impulse about 250 s.

r

Mixture ratio: mass oxidizer / mass fuel

V_e

Average exhaust velocity, m/s. The same measure as specific impulse in different units, numerically equal to specific impulse in N·s/kg.

C*

Characteristic velocity, m/s. Equal to chamber pressure multiplied by throat area, divided by mass flow rate. Used to check experimental rocket's combustion efficiency.

T_c

Chamber temperature, °C

d

Bulk density of fuel and oxidizer, g/cm³

Monopropellants

Propellant	Comment	Optimal expansion from 68.05 atm to 1 atm[ citation needed ]				Expansion from 68.05 atm to vacuum (0 atm) (Areanozzle = 40:1)[ citation needed ]
		Ve	Tc	d	C*	Ve	Tc	d	C*
Ammonium dinitramide (LMP-103S) [32] [33]	PRISMA mission (2010–2015) 5 S/Cs launched 2016 [34]		1608	1.24			1608	1.24
Hydrazine [33]	Common		883	1.01			883	1.01
Hydrogen peroxide	Common	1610	1270	1.45	1040	1860	1270	1.45	1040
Hydroxylammonium nitrate (AF-M315E) [33]			1893	1.46			1893	1.46
Nitromethane
Propellant	Comment	Ve	Tc	d	C*	Ve	Tc	d	C*

References

1. Larson, W.J.; Wertz, J.R. (1992). Space Mission Analysis and Design. Boston: Kluver Academic Publishers.
2. Sutton, G. P. (2003). "History of liquid propellant rocket engines in the united states". Journal of Propulsion and Power. 19 (6): 978–1007. doi:10.2514/2.6942.
3. Tsiolkovsky, Konstantin E. (1903), "The Exploration of Cosmic Space by Means of Reaction Devices (Исследование мировых пространств реактивными приборами)", The Science Review (in Russian) (5), archived from the original on 19 October 2008, retrieved 22 September 2008
4. Zumerchik, John, ed. (2001). Macmillan encyclopedia of energy . New York: Macmillan Reference USA. ISBN   0028650212. OCLC   44774933.
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External links

• Cpropep-Web an online computer program to calculate propellant performance in rocket engines
• Design Tool for Liquid Rocket Engine Thermodynamic Analysis is a computer program to predict the performance of the liquid-propellant rocket engines.