Liquid rocket propellants

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

Contents

[edit] History

[edit] Early development

The first rockets used liquid oxygen and gasoline as propellants. Both are readily available, cheap, high energy, 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 heroic insulation measures. Gasoline has since been replaced by RP-1, a highly refined grade of kerosene.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.

[edit] 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 not stable, 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.)

[edit] 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.

[edit] 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.


[edit] 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, Energia, 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. Hydrazine decomposes energetically to nitrogen and hydrogen, making it a fairly good monopropellant all by itself.

[edit] 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:

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
Optimum expansion from
68.05 atm to 1 atm		 Optimum expansion from
68.05 atm to 0 atm (vacuum) (Areanozzle = 40:1)
Oxidizer Fuel Ve r Tc d C* Ve r Tc d C*
LOX H2 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 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 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 2789 2.97 3407 1.42 1739 3274 3.01 3413 1.42 1739
N2H4 2885 2.81 3650 1.49 1824 3356 2.89 3666 1.50 1822
N2O4 MMH 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 2862 1.36 2992 1.21 1781 3369 1.42 2993 1.22 1770
N2H4/UDMH 50/50 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 2638 3.26 2848 1.30 1627 3123 3.41 2839 1.31 1617
MMH 2690 2.59 2849 1.27 1665 3178 2.71 2841 1.28 1655
UDMH 2668 3.13 2874 1.26 1648 3157 3.31 2864 1.27 1634
IRFNA IV HDA UDMH/DETA 60/40 2689 3.06 2903 1.32 1656 3187 3.25 2951 1.33 1641
MMH 2742 2.43 2953 1.29 1696 3242 2.58 2947 1.31 1680
UDMH 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 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

[edit] See also

  • Cpropep-Web an online computer program to calculate propellant performance in rocket engines
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