Application of compound compressible flow to nonuniformities in hypersonic propulsion systems

Abstract

An analytical model is presented to study the propagation of the forebody boundary layer inside the engine channel of a hypersonic engine, based on a compound compressible one-dimensional streamtube method. It is shown that the compound compressible flow model can be greatly simplified in the hypersonic regime. This model is appropriate for studying the effects of nonuniform flow inside a supersonic combustion ramjet. It is found that even a small boundary layer can produce noticeable changes in freestream properties inside the inlet. In the combustor models, it is seen that freestream Mach number is generally increased, and so static pressure and temperature are decreased. It is suggested that these effects will reduce the combustion reaction rate relative to that expected in a uniform flow and will therefore reduce the total heat release and increase losses inside the engine.