School of Engineering, The University of Queensland
Abstract
The successful operation of scramjet combustors requires compression of hypersonic viscous ducted flows and avoidance of separation effects which may preclude steady flow. Separation effects in scramjet inlets and combustors can be caused by shock wave/boundary layer interactions. The hypersonic turbulent flow experiments needed are inherently difficult to design because of the high sensitivity of the macroscopic flow parameters which cause the turbulent flow processes. Hence computational fluid dynamics (CFD) is a useful tool for the design and characterisation of models in hypersonic flows before model construction. One of the greatest challenges however is to ensure that the flow is being modeled accurately. In this paper, a commercial code has been used to model an experiment performed in a small reflected shock tunnel using a Mach 8.65 condition. The research being carried out in this facility is concerned with separation due to incident shock wave/turbulent boundary interactions in hypersonic ducted flows. The model is designed to produce two conical shocks which interact with a turbulent boundary layer and it is instrumented with pressure transducers and thin film heat transfer gauges. The measurements have allowed graphical representation of unseparated static wall pressure and heat flux prior to and after each wall interaction. The results of the simulations are in excellent agreement with the experimental data. The code has been applied to identify parameter boundaries in the design of a model of similar scale that will produce separated flow
