Supersonic and Hypersonic Compression Corner flows

Significant research effort is currently being focused all around the world to understand the interaction of turbulent boundary layers with shock waves. These flow phenomena play crucial role in inlets and isolators of high-speed propulsion systems like scramjet engines. The sensitivity of the inlet design to the pressure and heat transfer rate in these interactions make it very important to predict them accurately.

Conventional Reynolds averaged Navier-Stokes (RANS) methods often give erroneous results. Newer models are being developed to improve predictions. A parallel effort is currently under-way to use theoretical analysis and direct numerical simulation data to come up with advanced turbulence models for flows involving shock-turbulence interaction. Details of this work can be found here. These new models have worked exceptionally well in fundamental flows involving amplification  of homogeneous turbulence through normal shock waves.

Compression corners and shocks impinging on boundary layers are common in scramjet inlets and isolators. Due to the simplicity of the configuration, compression ramps have been extensively studied both experimentally and computationally. Conventional turbulence models, like k-epsilon and k-omega, usually predict too high an amplification of turbulence through the shock and this results in a smaller separation region than what is measured in experiments. The shock-unsteadiness modfication developed previously (Physics of Fluids, August 2003) gives a more realistic estimate of the separation region (AIAA Journal, March 2005). A typical pressure prediction for a 24 degree compression ramp in a Mach 2.84 flow in shown below.

At hypersonic Mach numbers typical of scramjet engines, the shock pattern caused by boundary layer separation can be quite complex, thereby making it harder to predict. Inaccurate turbulence models can result in flow field that is entirely different from reality. An example of such an interaction on a cone-flare geometry at Mach 9.1 is shown below.