In high-speed aircrafts, a boundary layer may interact with an externally generated or self-induced shock wave depending upon geometry. The impact of such an interaction will be harmful and must be factored into design tradeoffs. Non-stationary shock strengths and impingement locations on hypersonic vehicles exist due to changes in operating conditions and subsystem dynamics. Consequences of these non-stationary shock conditions include degraded engine performance and structural integrity due to increased pressure and thermal loads at the surface. Most of the work in the literature dealing with shock/boundary-layer interaction considers a calorically perfect gas. However, in hypersonic flight conditions, high-temperature effects must be considered to compute the flow field.

The current work aims to study the non-equilibrium effects of an oscillating shock wave interacting with a laminar boundary layer in hypersonic flows. For this purpose, a finite volume based flow solver is developed using Steger-warming flux splitting scheme with 5th order WENO reconstruction for calculating convective fluxes and Green's theorem for gradient evaluation of primitive variables associated with viscous terms. Park's two-temperature model for five species is used to develop the non-equilibrium part of the solver. Fourth order Runge-Kutta scheme is used for time marching.