Experimental and numerical studies of shock wave -- boundary-layer interactions over a cone-cylinder-flare in hypersonic flows

The present study provides experimental and numerical results for turbulent shockwave boundary layer interactions (SBLIs) over a blunt cone-cylinder-flare geometry, on which there is a lack of available data. Experiments were conducted in the Cranfield hypersonic gun tunnel, at Mach 8.2. Two configurations with varying nose length were studied using experimental testing and computational fluid dynamics (CFD). Three dimensionality has been generated by placing the model at incidence. The objective was to extend the range of data available for hypersonic flows and contribute to the design of re-entry vehicles. Experimental results cover forces and moments, that the numerical study aimed to replicate with Reynolds Averaged Navier-Stokes (RANS) modeling, and extended by analyzing the separated flow regions over the model. Additionally, the effect of the nose length on the aerodynamic forces and boundary layer transition were presented.