Shock-wave/boundary layer interaction plays a major role in any circumstances where the flow becomes supersonic, either locally or in totality. This phenomenon is not clearly understood when the transitional regime (from laminar to turbulent) of the boundary layer appears during the interaction process, which is the case for compressor or turbine cascades configurations and for laminar transport/business aircraft wing.
An experimental investigation in the S8Ch research wind tunnel of the ONERA Meudon Centre at a moderate supersonic (Mach number equal to 1.6) regime is carried out to quantify the effect of the shock-wave intensity on the boundary layer transition. The detection of the transition region is obtained by means of Schlieren visualizations, IR (Infra-red) thermography and TSP (Temperature Sensitive Paint) measurements. LES computation is performed on this configuration by using the block-structured solver elsA of ONERA. Comparisons of results are performed on two configurations, one at a moderate shock intensity and the other for a strong shock intensity leading to a massive boundary layer separation.