Liquid rocket propellant

Main article: Liquid-propellant rocket

The highest specific impulse chemical rockets (liquid-propellant rockets) use liquid fuel propellants. Approximately 170 different liquid propellants have undergone lab testing. This estimate excludes 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. However, there has not been a completely new propellant used in flight for nearly 30 years. 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. Bipropellants can be either hypergolic propellant or nonhypergolic. A hypergolic combination of oxidizer and fuel will start to burn upon contact. A nonhypergolic needs an ignition source.

History

Early development

Robert H. Goddard on March 16, 1926, holding the launching frame of his most notable invention — the first liquid-fueled rocket.

On March 16, 1926, Robert H. Goddard used liquid oxygen (LOX) and gasoline as propellants for his first partially successful liquid rocket launch. Both are readily available, cheap and highly energetic. Oxygen is a moderate cryogen — air will not liquefy against a liquid oxygen tank, so it is possible to store LOX briefly in a rocket without excessive insulation. Gasoline has since been replaced by different hydrocarbon fuels, for example RP-1 - a highly refined grade of kerosene. This combination is quite practical for rockets that need not be stored, and to this day, it is used in the first stages of many orbital launchers.

Wartime

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. 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 would ignite spontaneously on contact with the high density oxidizer. The German engine was powered by hydrogen peroxide and a fuel mixture of hydrazine hydrate and methyl alcohol. The U.S. engine was powered by nitric acid oxidizer and aniline. Both engines were used to power aircraft, the Me-163B Komet interceptor in the case of the German engine and RATO units to assist take-off of aircraft in the case of the U.S. engine.

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, and which cause their rockets to grow ever-thicker blankets of ice, are not practical. As the military is 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. In the case of nitric acid, the acid itself (HNO₃) is unstable, and corrodes most metals, making it difficult to store. The addition of a modest amount of nitrogen tetroxide, N₂O₄, turns the mixture red and keeps it from changing composition, but leaves 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 N₂O₄ 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.

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 used with a liquid oxygen oxidizer because the only by-product is water. As hydrogen in any state is very bulky, for lightweight vehicles it is typically stored as a deeply cryogenic liquid. This storage technique was mastered in the early 1950s as part of the hydrogen bomb development program at Los Alamos. It was then adopted for hydrogen fueled stages such as Centaur and Saturn upper stages in the late 50s and early 1960s. Even as a liquid, hydrogen has low density, requiring large tanks and pumps, and the 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. Most rockets that use hydrogen fuel use it in upper stages only.

Gaseous hydrogen is commercially produced by the fuel-rich burning of natural gas. Carbon forms a stronger bond with oxygen so the gaseous hydrogen is left behind. Liquid hydrogen is stored and transported without boil-off because helium, which has a lower boiling point than hydrogen, is the cooling refrigerant. Only when hydrogen is loaded on a launch vehicle (where there is no refrigeration) does it vent to the atmosphere.

Comparison to kerosene

Launch pad fires due to spilled kerosene are more damaging than hydrogen fires, primarily for two reasons. First, kerosene burns about 20% hotter (absolute temperature) than hydrogen. The second and more significant reason is 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 H₂O that is formed has a molecular weight of only 18 u compared to 29.9 u for air, so it rises quickly as well. Kerosene on the other hand falls to the ground and burns for hours when spilled in large quantities, unavoidably causing extensive heat damage that requires time consuming repairs and rebuilding. This is a lesson most frequently experienced by test stand crews involved with firings of large, unproven rocket engines. Hydrogen-fueled engines also have some special design requirements such as running propellant lines horizontally so traps do not form in the lines and cause ruptures due to boiling in confined spaces. These considerations, however, apply to all cryogens such as liquid oxygen and liquid natural gas as well. Use of liquid hydrogen fuel has an excellent safety record and superb performance that is well above that of all other practical chemical rocket propellants. (See bipropellant rocket engine performance table below.)

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 a 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, fluorine ignites on contact with most fuels, 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 that much more difficult. Finally, both lithium and fluorine are expensive compared to most rocket propellants. This combination has therefore never flown.

Methane

In November 2012, SpaceX CEO Elon Musk announced a new direction for the propulsion side of SpaceX: developing methane/LOX rocket engines. SpaceX had previously used only LOX/RP-1 for all of their primary propulsion engines. As of March 2014, SpaceX is actively developing the Raptor methalox bipropellant rocket engine with over 1,000,000 lbf (4,400 kN) of thrust. The engine is slated to be used on a future super-heavy rocket, the MCT launch vehicle.

Firefly Space Systems announced in July 2014 their plans to use methane fuel for their small satellite launch vehicle, Firefly Alpha, utilizing an aerospike engine design.

Blue Origin and United Launch Alliance announced in September 2014 the joint development of the BE-4 lox/methane engine. The BE-4 will provide 550,000 lbf of thrust.

Monopropellants

• Hydrogen peroxide decomposes to steam and oxygen
• Hydrazine decomposes energetically to nitrogen, hydrogen and ammonia (2N₂H₄-->N₂+H₂+2NH₃) and is the most widely used in space vehicles. (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

Current use

Typical performances of common propellants

Propellant mix	Vacuum Isp (seconds)	Effective exhaust velocity (m/s)
liquid oxygen/ liquid hydrogen	455	4462
liquid oxygen/ kerosene (RP-1)	358	3510
nitrogen tetroxide/ hydrazine	344	3369
n.b. All performances at a nozzle expansion ratio of 40

Here are some common liquid fuel combinations in use today:

• LOX and kerosene (RP-1). Used for the lower stages of the Soyuz boosters, and the first stage of the U.S. Saturn V, Atlas, and Falcon 9 boosters. Very similar to Robert Goddard's first rocket.

• LOX and liquid hydrogen, used in the stages of the Space Shuttle, Space Launch System, Ariane 5, Delta IV, H-IIB, GSLV and Centaur.

• Nitrogen tetroxide (N₂O₄) and UDMH or MMH. 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 (N₂H₄) and Aerozine-50 are also used in deep space missions because they are storable and hypergolic, and can be used as a monopropellant with a catalyst.

Upper stage use

The liquid rocket engine propellant combination of liquid oxygen and hydrogen offers the highest specific impulse of currently used conventional rockets. This extra performance largely offsets the disadvantage of low density. Low density of a propellant leads to larger fuel tanks. However, a small increase in specific impulse in an upper stage application can have a significant increase in payload to orbit capability.

Propellant table

To approximate Isp at other chamber pressures
Absolute pressure (atm) {psi}	Multiply by
6,895 kPa (68.05) {1000}	1.00
6,205 kPa (61.24) {900}	0.99
5,516 kPa (54.44) {800}	0.98
4,826 kPa (47.63) {700}	0.97
4,137 kPa (40.83) {600}	0.95
3,447 kPa (34.02) {500}	0.93
2,758 kPa (27.22) {400}	0.91
2,068 kPa (20.41) {300}	0.88

JANAF thermochemical data used throughout. Calculations performed by Rocketdyne, results appear in "Modern Engineering for Design of Liquid-Propellant Rocket Engines", Huzel and Huang. Some of the units have been converted to metric, but pressures have not. These are best-possible specific impulse calculations.

Assumptions:

• adiabatic combustion
• isentropic expansion
• one-dimensional expansion
• shifting equilibrium

Definitions

Ve	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
Tc	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

			Optimum expansion from 68.05 atm to 1 atm					Expansion in vacuum (0 atm) from 68.05 atm to almost 0 atm (Areanozzle = 40:1)
Oxidizer	Fuel	comment	Ve	r	Tc	d	C*	Ve	r	Tc	d	C*
LOX	H2	common	3816	4.13	2740	0.29	2416	4462	4.83	2978	0.32	2386
	H2-Be 49/51		4498	0.87	2558	0.23	2833	5295	0.91	2589	0.24	2850
	CH4 (methane)		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	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/N2H5NO3 55/26/19	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
N2H4	B2H6		3342	1.16	2231	0.63	2080	3953	1.16	2231	0.63	2080
	B5H9		3204	1.27	2441	0.80	1960	3819	1.27	2441	0.80	1960
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 C₁₀H₁₈)
• MMH: CH₃NHNH₂

r	Mixture ratio: mass oxidizer / mass fuel
Ve	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.
Tc	Chamber temperature, °C
d	Bulk density of fuel and oxidizer, g/cm³

Monopropellants

		Optimum expansion from 68.05 atm to 1 atm				Expansion in vacuum (0 atm) from 68.05 atm to almost 0 atm (Areanozzle = 40:1)
Propellant	comment	Ve	Tc	d	C*	Ve	Tc	d	C*
Hydrazine	common
100% Hydrogen peroxide	common	1610	1270	1.4	1040	1860	1270	1.4	1040
LMP-103s	recently test-flown						1608
Nitromethane
Propellant	comment	Ve	Tc	d	C*	Ve	Tc	d	C*

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.
• for a history of liquid rocket propellants in the US by a pioneering rocket fuel developer.