Supersonic combustion of storable liquid fuels in mach 3.0 to 5.0 air streams

Abstract

Supersonic combustion of reactive aluminum alkyl fuels has been experimentally demonstrated in two-dimensional ducted combustors and adjacent to a flat plate. Fuel was injected from the combustor walls through multiple orifices and ignited spontaneously. Stable supersonic heat release was maintained as evidenced by schlieren and direct motion pictures of the flow field and deduced from static and pitot pressure measurements in the combustion zone. zone. The results from the ducted combustor tests were correlated with simple theoretical models thus enabling a reasonable determination of combustion efficiency to be made. A theoretical model of constant-pressure heat release on a flat plate in supersonic flow is postulated. Normal force, coefficients and specific impulse values are tabulated for a variety of flight Mach Numbers and altitudes. Additional refinements in this theoretical model were required to adequately describe the experimental results. In a test simulating Mach 5 flight at 66,000-feet altitude, a side force specific impulse of 1350 sec was measured at equivalence ratio of one. Combustion was only partially completed 12 in. downstream of fuel injection. Based on the theoretical model an additional 12 in. of combustor length and 36 in. of expansion length would be required to obtain, the estimated theoretical impulse of 5760 sec. The interaction of a vaporizing liquid droplet with a supersonic stream is considered. Additional refinements were made in the existing theories on droplet trajectory to include the influences of a separated zone and the normal component of velocity of the external stream. Calculations of the trajectory and evaporation of the estimated mean droplet size based on the modified technic were in general agreement with the observed fame zone and deduced combustion efficiency.