Liquid rocket propellants

The highest specific impulse chemical rockets use liquid propellants. This type of propellent has a long history going back to the first rockets and is still in use in for example the Space Shuttle and Ariane 5.

History

Early development

On March 16, 1926, Robert H. Goddard used liquid oxygen and gasoline as propellants for his first successful liquid rocket launch. Both are readily available, cheap, highly energetic, and dense. Oxygen is a moderate cryogen — air will not liquify 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 Gasoline has since been replaced by RP-1, a highly refined grade of kerosene.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 most orbital launchers, as well as the long-range offensive missiles of China and North Korea.

1950s

During the 1950s there was a great burst of activity by propellant chemists to find high-energy liquid propellants better suited to the military. Military rockets need to sit in 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, virtually all of which were dead ends.

For instance, in the case of nitric acid, the acid itself (HNO₃) is unstable, and gives off NO₂ fumes (hence the name white fuming nitric acid). Unlike nitrous oxide (N₂O), these nitrogen dioxide fumes are extremely toxic. The addition of large amounts of dinitrogen tetroxide (N₂O₄) makes the mixture red, but keeps it from changing composition, leaving the problem that nitric acid will eat any container it is placed in, releasing gases that can build up pressure in the process. The breakthrough was the addition of a little hydrofluoric acid (HF), which forms a self-healing metal fluoride on the interior of tank walls and makes "Inhibited" Red Fuming Nitric Acid storable. Although the development of military propellants was treated with the greatest secrecy, the trick to inhibiting nitric acid was published shortly after its discovery in 1954 and Russian rockets with the same fuel appeared shortly afterwards, the first being the SS-1B ("Scud"). Eventually the chemists gave up stabilizing HNO₃ with N₂O₄, and just used straight N₂O₄, which is a slightly better oxidizer anyway. (In the propellant table below, note that N₂O₄ is always in equilibrium with NO₂, and so mixtures are sometimes quoted with the latter.)

Hydrogen

Many early rocket theorists believed that hydrogen would be a marvellous propellant, since it gives the highest specific impulse. As hydrogen in any state is very bulky, for flightweight vehicles it is typically stored as a deeply cryogenic liquid. This storage technique was mastered in the 1960s as part of the Saturn and Centaur upper-stage programs. Even as a liquid, hydrogen has low density, requiring large, heavy tanks and pumps, and the extreme cold requires heavy and potentially dangerous tank insulation. This extra weight reduces the mass fraction of the vehicle and offsets the specific impulse advantage. Most rockets that use hydrogen fuel use it in upper stages only, where a low thrust-to-empty-mass ratio can be tolerated and where a hydrogen stage's low total mass reduces the size of the lower stages. Those rockets that use hydrogen fuel in their lower stages, like the Space Shuttle, Delta IV, and Ariane 5, often use powerful and dense solid rocket motors at liftoff to improve their acceleration off the pad and thus reduce gravity losses early in flight.

Lithium/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, and hydrogen, while not hypergolic, is an explosive hazard. 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. The rocket exhaust is also ionized, which would interfere with radio communication with the rocket. Finally, both lithium and fluorine are expensive and rare, enough to actually matter. This combination has therefore never flown.

Monopropellants

*Hydrogen peroxide decomposes to steam and oxygen
*Hydrazine decomposes energetically to nitrogen and hydrogen, making it a fairly good monopropellant all by itself.
*Nitrous oxide decomposes to nitrogen and oxygen
*Steam when externally heated gives a reasonably modest Isp of up to 190 seconds, depending on material corrosion and thermal limits

Current use

Here are some common liquid fuel combinations in use today:

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

* LOX and liquid hydrogen, used in the Space Shuttle, Ariane 5, Delta IV and the Centaur upper stage.

* Nitrogen tetroxide (N₂O₄) and hydrazine (N₂H₄). Used in military, orbital and deep space rockets, because both liquids are storable for long periods at reasonable temperatures and pressures.

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

"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

Bipropellants

			Optimum expansion from 68.05 atm to 1 atm					Optimum expansion from 68.05 atm to 0 atm (vacuum) (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		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		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/H20 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				Optimum expansion from 68.05 atm to 0 atm (vacuum) (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
Propellant	comment	"Ve"	"Tc"	"d"	"C*"	"Ve"	"Tc"	"d"	"C*"

ee also

* [http://rocketworkbench.sourceforge.net/equil.phtml Cpropep-Web] an online computer program to calculate propellant performance in rocket engines