A Computational Study of Supersonic Combustion Relevant to Air–Breathing Engines

The development of high–speed flight and space access vehicles requires the solution of many technical challenges associated with the comparatively small net thrust at supersonic or hypersonic flight speeds. One of the more essential issues is the design of an air–breathing propulsion system capable of operating over the wide range of Mach (Ma) numbers, desired to facilitate the advancement of high–speed flight and space access vehicles. At flight speeds above Ma≈3 turbofan engines fall short since the compressed air through the engine reaches such temperatures that the compressor stage fan blades begin to fail. Instead ramjet engines, in which the profile of the air intake guarantees that the supersonic approach flow is decelerated to a subsonic flow through the combustor, where fuel is injected prior to mixing, self-ignition and combustion, may be used. However, beyond Ma≈5 extreme temperatures and pressure losses occur when decelerating the supersonic airflow to subsonic conditions, making the ramjet unpractical at higher flight speeds. At flight speeds beyond Ma≈5, supersonic combustion ramjets, or scramjets, in which the flow trough the inlet and combustor remain supersonic may be used. Achieving high combustion efficiency under such conditions, with residence time on the order of 1 ms, places extreme demands on the inlet, combustor, fuel–injector as well as on the nozzle design, [1]. The mixing of fuel and air, the self–ignition and the flame stabilization are thus critical processes.