ECCOMAS Congress 2016 5 - 10 June 2016, Crete Island, Greece STS 1 - The CAero2 Platform: Dissemination of Computational Case Studies in Aeronautics Transition effect on a shock-wave / boundary layer interaction Reynald Bur and Eric Garnier ONERA, 8 rue des Vertugadins, 92190 Meudon, France Abstract 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. Definition of the test case The experimental investigation is executed in the S8Ch wind tunnel of the ONERA Meudon Centre. This facility is a continuous wind tunnel supplied with desiccated atmospheric air. The stagnation conditions are near ambient pressure and temperature: p st =0.96×10 5 ± 300 Pa and T st =300 ± 10 K. It is constituted by a rectilinear channel having a test section with a height of 120mm and a span of 120mm (in full nozzle configuration). The test section side walls and upper wall are equipped with high quality glass windows to allow using optical techniques, mainly to detect the location of the boundary layer transition. Figure 1 shows the test set-up with the Mach 1.6 full nozzle configuration, the unit Reynolds number being around 14 × 10 6 m -1 for this Mach number. The shock generator wedge (with its displacement device) has been installed on the upper wall of the test section. The flat plate under study is mounted above the lower wall and the re-generated boundary layer will interact with the shock-wave. The location of the flat plate has to be movable in both the longitudinal and vertical directions to allow the study of the boundary layer behaviour with respect to the shock impingement. The leading edge shape of the flat plate is designed to take into account