Combustion Characteristics of a Hypermixer Scramjet Engine

A scramjet engine with a wall-mounted hypermixer injector, which generates streamwise vortices for enhancing supersonic mixing and combustion, is examined at a Mach 8 simulated flight condition in the High Enthalpy Shock Tunnel (HIEST). The engine and the fuel injector are full scale models of the HyShot-IV flight experiment planned for 2005 by JAXA and University of Queensland (UQ). Main purpose of the present study is to clarify the combustion and operation characteristics of the hypermixer scramjet owing to the ability of the streamwise vortices for mixing enhancement and boundary layer control. For comparison, two injectors with normal and parallel injection without streamwise vortex generation are also examined. The results show the superior performance of the hypermixer injector in scramjet mode obtaining higher pressure rise in a shorter distance compared to the other two injectors. In the case of the hypermixer injector, a 1D analysis of an inviscid nozzle flow shows the increment in the specific impulse due to combustion to be 2,649 and 2,224 sec for the equivalence ratio, Φ=0.3 and 0.6, respectively. At Φ=1.0 and 1.5, sudden rapid combustion of the premixed fuel at the end of the combustor generates a strong pressure wave, which propagates upstream up to the injector location and decades there. As a result, a new quasi-steady combusting flow is established throughout the combustor downstream of the injector. The pressure wave is identified as a kind of detonation wave, which is suggested to propagate upstream mainly through the streamwise vortices. As a driving force of the upstream propagation of the detonation wave, mixing and combustion enhanced through the interaction between the detonation wave and the streamwise vortices are considered.