Combustion of hydrogen and methane to simulate expansion of storablepropellants

An experimental investigation of exhaust-nozzle temperatures for the storable aystem, 50 percent UDMH-50 percent hydrazine/nitrogen tetroxide, was conducted using hydrogen and methane fuel burned in oxygen-enriched air to provide the s'ame atomic constituents as the storable propellants. Oxidant-fuel mass ratios of 1.6 to 2.5 at 3.7 atmospheres combustionchamber pressure were simulated by control of the fuel. oxygeh, and ai< flows; inlet enthalpies were duplicated by preheating the oxidant in a pebble-bed storage heater. Static temperatures, measured by a spectral line reversal pyrometer at 5 stations in the expanding portion of a 5.5-area-ratio, Mach 3 nozzle, were found ta be nearly insensitive to oxidat-fuel ratio and close to temperatures calculated for a frozen expansion. This study demonstrated the feasibility of the simulation technique for exhaust-nozzle kinetic studies. Results agreed well with a simplified kinetic analysis based on a I'sudden freezing" of the nozzle recombination reactions at an area ratio of 1.04 upstream of the throat. c. Introduction ' I The storable propellant system consisting of a 50 percent by weight mixture of hydrazine and unsymmetrical dimethylhydrazine (UDMH) fuel and nitrogen tetroxide oxidizer is used in many advanced rocket engines, and experimental performance of several space-vehicle engines, using this propellant system, has been reported. ( l ) The actual engine performance is degraded by combustion, aerodynamic, and kinetic losses. The latter result from the reduction in nozzle energy release due to the incomplete recombination of the atoms and free radicals formed during the high-temperature combustion. These chemical recbmbination reactions have finite reaction times, which are often slow with respect to the residence time of the exhaust products in the nozzle. The prediction of kinetic losses for the storable propellant system has been the subject of several 3, 4); it% desirable, however, to supplement these predictions, which are based on idealized laboratory reaction-rate data, with experimental measurements under conditions more closely approaching those encountered in rocket engine environments. ,. chamber products and enthalpy of the desired storablepropellant system using gaseous propellants, eliminating combustor inefficiencies and instabilities. A recent report also suggests this method of simulation. (5) Experimental static temperatures are reported for the simulated storable propellants at a combustionchamber pressure of 3.7' atmospheres, covering a range of oxidant-fuel weight ratios of 1.6 to 2.5 (equivalence ratios of 1.4 to 0.9). The experimental data are compared to analytical calculations based on theoretical equilibrium and kinetic expansions, and the application of the data to performance predictions of rocket engines is discussed.