Tetranitromet hane as Oxidizer in Rocket Propellants JOHANN G. TSCHINKEL
U. S.
Department o f the Army, Redsfone Arsenal, Huntsville, Ala.
Performance calculations for tetranitromethane, a potential oxidizer in liquid rocket propellants, have been carried out comparing tetranitromethane and common oxidizers, all in combination with hydrocarbon fuel. For the tetranitromethane system the ideal cutoff velocity as range parameter is 14% higher than for the other systems; however, the high freezing temperature o f tetranitromethane is a drawback in its use. In order to lower this freezing temperature, freezing point depressants may b e added. These depressants also increase the explosive sensitivity o f tetranitromethane, which i s low.
E T R A N I T R O M E T H A S E aroused interest as a rocket oxidizer during the dwelopment of thc Y-2 rocket in Germany, when a boost of range was sought by use of propellants of higher density without major design changes ( I S ) . A pilot plant for production was set up by I. G. Farbenindustrie ( 7 ) ,and about 10 tons of material (mixture with nitlo$en tetroxide) were accumulated a t the German Rocket Center, Peenemunde, with the intention of testing it in rocket motors. Capture of Peenemunde by the Russians prevented these tests. Work in the United States v a s continued mainly by Hefco Labs during 1948-50 (10-12). Emphasis TYas on study of handling properties and freezing point depression. A pilot scale manufacture was set up by Sitroform, Inc., Newark, N. J., but in 1953 was destroyed by an explosion, Tvhich, t o the author's knoxledge, discouraged any further attempts of manufacture in larger scale. However it can be obtained, in small quantity, from various sources a t the high price of about 520 per pound. Cost in mass production has been estimated a t a fraction of a dollar per pound ( 7 ) . Best process is nitration of acetylene with nitric acid.
Pkysico-Chemical Properties A meritorius property of tetranitromethane is its high density. Table I gives a comparison of this density and that of some common oxidizers.
Tu ble I.
Density and Oxygen Content o f Rocket Oxidizers
Oxidizer Liquid oxygen (at -183' C.) Tetranitromethane Liquid nitrogen tetroxide Nitric acid P u r e hydrogen peroxide 90 wt. % hydrogen peroxide
Iiorniula 0 2
C(N0z)i KsO1
"01
IIzOz HzOz 0 . 2 1 HzO
+
Wt. %
Density
0 2
Kg /Liter a t 20° C.
100 65.3 69.5 63.5 47.0 42.2
1,14 1.64 1.45 1.50 1.460 1.405
Active 02, ICg.,'Liter 1.14
1.07 1.01 0.97 0.69 0.59
ume of carbon of 11 ml. (4, p. 524) and a specific tetranitromethane of SI
=
1.6741 - 0,001734 X t (" C.j (gram/cu. am.)
732
of (1)
the molar volume a t boiling temperature minus carbon volume is calculated to be 124 cu. cm. The net density of the nitrogroups follows 184/124 = 1.484 a t 126" C.; using the temperature coefficient of Equation I,vie obtain a net oxygen-density a t 20" C. = 1.671 x 0.696 = 1.16 kg. oxygen per liter net oxidizer. This means the nitrogroups in t,etranitromethane provide a somewhat higher packing density of oxygen than liquid oxygen itself. Other physical properties (5) important for a rocket oxidizer are: Freezing temperature Boiling temperature Viscosity at, 20" C .
13.8" (1. 126" C . 1 . 7 6 cp.
Although the boiling temperature and viscosity are a t a level convenient for handling of a propellant, the high freezing temperature places a severe restriction on use of tetranitromethane at winter temperatures. Therefore, much effort of research (6) was expended to find depressants of freezing point that ~5-ould not impair appreciably the performance and other desirable properties. Energy of Decomposition. It is desirable that a rocket oxidizer decompose8 exothermally. Burning carbon to carbon dioxide, the equation of decomposition yields about 455 cal. per gram at 25" C. reference temperature (I! 6). 4 considerably higher value of 610 cal. per gram was quoted by two laboratories (3,8) as their determination. As the new value has not been published, the following calculations for enthalpy of decomposition of tetranitromethane use thc classical value. I-Iowever, use of the new value would improve the thrust performance by several per cent. cal. ( +-+ 2N2 + 302+- 89,500 94,450 cal. + 2x2 f 402 - 4,800 cal.
C(N02)* = COz
c0z
=
c,
C(NO2)& = C, Tetranitromethane contains nearly as much oxygen Tveight per unit volume as liquid oxygen itself. I n order to appraise the oxygen-density justly, i t must be considered that the carbon atom is combustible and that it should be treated as fuel. A net oxygen-density was derived as follows: "sing an atomic vol-
ma28
=
455 cnl. per gram)
0 2
This heat evolution is due entirely to the combustion of carbon, and the separation into elements is even endothermal. The resulting adiabatic explosion temperature is quoted as 2675" C. (5, 1 4 ) .
INDUSTRIAL AND ENGINEERING CHEMISTRY
Vol. 48, No. 4
ROCKET PROPELLANTS
Table II.
Thrust Performance of Various Oxidizers with Same Fuel
Fuel: hydrocarbon of (CH1.8)s average composition. Combustion chamber presaure: 16 atm. Expansion pressure ratio: 16 to 1. All Components liquid at 25' C. (except 02) Ah Combustion (Ah)kin. Isp wt. % (H~O-GW) i6:i (*h)kin. i6:i rap. Oxidizer On F/O (F/O)equiv. Cal./G. Cal./G. (Ah)oomb. Sec. (Relative) C(N0n)r 65.3 0.16 1.10 1602 578 0.361 224 0.91 HaOn 47 0.14 1.06 1502 596 0.397 228 0.93 (N0z)Iiquid 69.5 0.25 1.28 1536 589 0.383 226 0.92 HNOi 63.5 0.20 1.125 1345 553 0.411 219 0.89 Hz0n f 0 . 2 1 HzO .-(go wt. % HaOn) 42 0.126 1.155 1158 557 0.481 220 0.90 (Un)liquid at -183O C. 100 0.32 1.14 2096 687 0.328 245 1.00
.
Propellant Energy Content The energy content of a propellant, expressed as enthalpy of complete combustion, is not an adequate parameter to compare rocket motor performance: Complete combustion cannot be realized because of dissociation, and the flame gas properties influence what enthalpy fraction is convertible into kinetic energy of the stream. Therefore, as commonly done, calculations of exhaust velocities for the various oxidizers combining with the same fuel have been carried out (Table 11) supplementing earlier calculations (W, 13). Hydrocarbon fuel oil of the average composition, (CHI.S)z, was chosen as fuel, combustion pressure was assumed a t 16 atm., and the gases were expanded isentropically to 1 atm. The result is compared as specific impulse, thrust developed per unit weight flow of propellant. The mixing ratio of fuel and oxidizer was selected a t such a value as to give a maximum thrust per unit propellant volume. As expected, all oxidizers show a lower thrust performance than liquid oxygen, since they are diluted oxygen, so to speak. Also, all oxidizers fall about 10% short of liquid oxygen in specific impulse, although hydrogen peroxide only has about 40 weight per cent available oxygen; however, this compound makes up for this deficiency by its exothermal decomposition which contributes about 370 cal. per gram of hydrogen peroxide to the heat of combustion.
the design of structure, pumps, etc., needs t o be adjusted t o the use of a heavier propellant. This variation was neglected t o simplify the following estimation. Using this value of 2.15 liters per kg. and inserting exhaust velocities and propellant densities calculated before, a comparison of ideal cutoff velocities for the selected propellants is obtained. Table I11 shows the low density propellant (fuel oil-oxygen) to fall off to the last place in the range parameter, UId.'. Under the assumed ideal conditions, range is proportional t o U,d.*. The parameter shows tetranitromethane a t the top. Range calculations were carried out (13) on a V-2 type rocket where necessary adjustments in size of some components, such as pumps, were made, which amounts t o adjusting the parameter Vp/M,. Calculations refer to vertical flight in vacuum. Range came out t o be proportional t o about 1.5 power of U , d . . The range improvement of tetranitromethane versus hydrocarbonoxygen comes out better than 20%.
Table 111. Oxidizer C(NOt)r HzOt (Nodliquid
"Or
Propellant Density and Flight Performance For a crude comparison of the influence of propellant density, the thrust per unit volume flow, so-called density impulse, is often used. A more refined, but still simple, parameter is readily obtained from the basic rocket flight equation.
This ideal cutoff velocity is the velocity that a rocket would obtain in gravity-free and empty space if its initial mass, M,, is decreased t o its cutoff mass, M,, by exhausting it with a jet velocity, u,. The propellant mass, the propellant specific mass, and propellant tank volume are introduced into the mass ratio.
The parameter, Vp/M,, signifies the tank volume realized per unit structural mass and is thus a measure for the perfection of lightweight design. Selecting a rocket of V-2 size and type for comparison, the ratio of V p / M , is found t o be near 2.15 liters per kg. This parameter is not strictly independent of propellant density since April 1956
Ideal Cutoff Velocity for Selected Propellants
+
HeOt 0.21 HnO (0e)lisuid
UB
2202 2235 2222 2153 2160 2400
SP
ad.
Uid. 1 . 4 5 5 1 . 4 1 8 3122 1 . 3 4 0 1 . 3 7 9 3082 1 . 2 9 5 1 . 3 3 0 2955 1 . 3 3 7 1 . 3 5 3 2913 1 . 3 0 5 1 . 3 3 7 2888 0 . 9 9 2 1 . 1 4 2 2741 ue
Uid.
(Rela-
tive) 1.14 1.12 1.08 1.06 1.05 1.00
Uid.l (Relative) 1.30 1.25 1.17 1.12 1.10 1.00
Range (Relative) 1.22 1.16 1.12 1.OQ 1.07 1.00
Figure 1 shows the exhaust velocity and the ideal cutoff velocity as a function of the mixture ratio for the system tetranitromethane-(CHI&. The mixture ratio is in terms of fuel equivalence kaction. The maximum of exhaust velocity and of the range parameter, U i d . , is on the fuel rich side a t q = 0.6 or (F/O)equiv,= 1.50; stoichiometric weight ratio is F/O = 0.1455.
Freezing Point Depression Although i t should be very possible t o keep a propellant in a large rocket from freezing by some form of heating, it would be preferable to lower the freezing temperature by a suitable additive without altering the favorable properties too much. Such an additive should have a low molecular weight, high content of available oxygen, high density, high boiling temperature, low corrosivity, and, finally, a low content in combustible atoms. The molar depression constant has been estimated theoretically t o be about 32" to 35" C. per gram mole depressant per kilogram tetranitromethane (6). However, in experiments of Hefco Labs ( l a ) , a value of the molar depression of only 11.5' C. per mole per kilogram was found, corresponding to a heat of fusion of tetranitromethane of 2800 cal. per gram mole. Inserting this value into Equation 4 from reference ( 4 ) :
I N D U S T R I A L A N D E N G I N E E R I N G CHEMISTRY
133
temperature to less than 0" C. The additive lowered the ideal cutoff velocity by only 2 to 3%. Pe IDEAL VOLUME
CUTOFF PER
VELOCITY KG
DRY
Uid
WEIGHT
I ATM
F O R ROCKET OF 2.15 L I T TAN
.of nitromethane can be composed of methyl glycol and tetranitromethane t o a con-
Table IV.
7.
Freezing Temperature of Tetranitromethane Mixtures
Wt. of Freezing Additive in Temp., Mixture C. CHaS02,mol. 2 . 0 1 0 . 5 (12) wt.=61 5.0 6.5 15.0 - 1 . 3 25.0 - 8 . 5 50.0 -26 PIT201, mol. 10.0 0 (16) w t . = 60 2 0 . 0 -14 3 5 . 0 -30
xiole Fraction TNbI 0.938 0 . 855 0.638 0 483 0.237 0 734 0.551 0,362
I N D U S T R I A L A N D E N G I N E E R I N G CHEMISTRY
Freeein:: Temp. Depression Calcd. Eq. 5 Exptl. 9 26 42
3.3 7.3 15 22
84
40
4
18
35
60
14 28
44
ti^ of Exptl. Calod Depression
0.83 0.81
0.57 0.52 0.47 0.77 0.80 0.73
Vol. 48, No. 4
ROCKET PROPELLANTS
Table V.
Performance Comparison on Oxidizer Tetranitromethane Mixed with Freezing Point Depressants
Chamber pressure: 16 a t m . ; expansion pressure ratio: 16 t o 1 ; fuel: Oxidizer TNM Wt. Fraction mole T!, TNM Additive fraction ( F / O ) e q u i v . Sox. C. F/O (F/O)equiv. 1.00 1.00 0.25 1.64 +14 0.16 1.10 0.85 0.15Mechylnitrate 0.69 0.38 1.57 - 0 0.163 1.42 0.70 0.30Methylnitrate 0.48 0.62 1.49 - 8 0.151 1.64 0.85 0.15Nitromethane 0.64 0.44 1.55 2 0,148 1.40 0.80 0.20Nz04 0.65 0.20 1.61 -15 0.220 1.50
-
hydrocarbon H CHI.^)^: Propellant SP, PB, ! ; :S
1.455 1.410 1.355 1,400 1.400
2200 2195 2220 2180 2190
ad. (Uid.)(Rrlntivo~. 3120 3050 3030 3030 3040
1.00 0.98 0.97 0.97 0.975
(A/L).~~L,. = enthalpy difference of combustion, cal./g., 2.5’
Table VI.
Detonation in Trautzel Block
Blasting C a p Size No. (Germany TNM Army S t d . ) Pure 1 12
2 3 4 5
TZ;RI/NsOp
70/30 (Wt.)
TNT
0
0 13 48 37
44 65 86 71
0 0
218 268 332
Nitroglycerin
TNM/Benzene, 87/13.(Wt.) (Stoioh.)
171 172 379 407 397
413 404 404 645
AH! (Ah)k,, I E L
Mc MO Mp
+
P
Table VII. Fuel Component Nitromethane,
Sm
=
1.136
Methanol, 520 = 0.792
Methyl glycol, CHaO.CHI,CH20H. 820= 0.965
Liquid Monopropellants
Wt. %
100 81.3 55.5 100 48 46 37 26 100 43 41.5 33 23
TNM, Wt. % ’ (100)
(F/O)equiv.
to C O t o GO, (0.125) (0.25) 1.25 1.75 18:7 1.00 1.42 44.5 0.69 1.00 .. 3.00 52 1.30 1175 54 1.25 1.70 63 1.00 1.35 74 0.70 1.00 .. 3.5 5.0 57 1.35 1.75 48.5 1.25 1.65 67 1.00 1.35 77 0.71 1.00
Density a t 20° C.
(1.64) 1.136 1.205 1.316 0.792 1.085 1.100
1.175 1.265 0.965 1.260 1.270 1.332 1.413
S SP t
AT Tj US
Uld.
V p X
c.
reference temperature, COzand HzO gas as products = molar enthalpy of fusion, cal. /mole (see Equation 4) = enthalpy difference converted into kinetic energy of the exhaust gas stream, cal./g. = specific impulse, thrust produced per unit weight flow rate, sec. = mass of rocket a t motor cutoff = take-off mass of rocket = mass of propellant - (J!’/U)equiv. = equivalence fuel fraction, g = 1 (F/O)equiv.’ q = 0.5 signifies stoichiometric (COZ, HzO) = specific mass (weight), density, g./cu. cm. (kg./l.) = propellant specific weight (kg./l.) = temperature, a C. = freezing temperature depression, O C. = temperature of fusion, absolute, K. = exhaust velocity, theoretical, m./sec. = ideal cutoff velocity, m./sec. (see Equation 2 ) = volume of propellant, effective tank volume = mole fraction (of tetranitromethane)
.
literature Cited (1) Berger, C., C o m p t . rend. 151, 815 (1910); Bull. soc. c h i m . 9,
31 (1911).
(2) Braun, W. von, Hager, K. F., Tschinkel, J. G., Redstone Arsenal, GMDD, Tech. Rept. No. 5, p. 6 (February 1946).
siderably higher density than nitromethane itself (1.260 compared to 1.136), but no data on safety of such mixtures are available at this time. Nitromethane-tetranitromethane mixtures are much less sensitive to initiation than stoichiometrically equivalent hydrocarbon-tetranitromethane mixtures (11); the carbon dioxidestoichiometric nitromethane- tetranitromethane mixture is more sensitive than the carbon monoxide-stoichiometric mixture. Nitromethane-tetranitroniethanemixtures are subjects of several patents (9).
Nomenclature
F/O = fuel to oxidizer weight ratio (F/O),,,,,. = fuel to oxidizer equivalence ratio =
/O (F/O)atoioh.’
g =
where stoichiometric is based on complete combustion to Con and HzO gravitational acceleration
April 1956
*
(3) California Institute of Technology, Jet Propulsion Lab, Director’s Office, letter dated Dec. 8, 1955. (4) Glasstone, S., “Physical Chemistry,” 2nd ed., p. 644, Van Nostrand, Kew York, 1946. (5) Gmelin Institute, P e e n e m u n d e Archiv. 124, KO.2, (6) Ibid., 124, No. 4, (7) Hager, IC. F., ISD.EN*. CHEW41, 2168 (1949). (8) Hannum, J. A,, Hefco Labs Inc., Detroit, Mich., oral com-
munication. (9) Hannum, J. A., U. S. Patent 2,537,526 (Jan. 9, 1951); Ibid.,
2,538,516 (Jan. 16, 1951). (10) Hefco Labs Inc., Detroit, Mich., Tech. Rept. on Rocket Fuels, Office of Naval Research Proiect No. 220019, May-December 1948. (11) I b i d . , December 1948-May 1949. (12) Ibid., June 1949-February 1950. (13) Heller, G., Thiel, W., Peenemunde Archiv. 81, KO.12 (October 1941). (14) Roth, J. F., 2.g e s . Schiess-u. S p r e n g s t o f w . 36, 53 (1941). (15) Schabert, R., Tschinkel, J. G., Peenemunde Archiv. 110, S o . 15 (March 1, 1944). RECEIVEDfor review September 24, 1955.
INDUSTRIAL AND ENGINEERING CHEMISTRY
ACCEPTED J a n u a r y 28, 1956.
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