Shock Wave Turbulent Boundary Layer Interaction in a 2-D Compression Corner ASMELASH HAFTU AMAHA Department of Mechanical Engineering (Gas Turbine Technology), Defence Institute of Advanced Technology Girinagar, Pune-411025, Maharashtra, India AMARJIT SINGH Department of Aerospace Engineering, Defence Institute of Advanced Technology Girinagar, Pune-411025, Maharashtra, India ROSCHELLE R. MARTIS Department of Aerospace Engineering, Defence Institute of Advanced Technology Girinagar, Pune-411025, Maharashtra, India Abstract: A computational study has been carried out to analyze the supersonic shock wave turbulent boundary layer interaction in a 2-D compression corner for a free stream Mach number of 2.94. The study has been done for a unit Reynolds number of 36.4x 10 6 per meter and 20 0 corner angle. The model has been analyzed using 2-D numerical simulations based on a commercially available Computational Fluid Dynamics (CFD) Code that employs k-ω Shear Stress Transport (SST) turbulence model. The substantiation of the CFD code and the turbulence model used is obtained by comparing with the experimental results available in literature. Comparison of the surface pressure distribution with experiment exhibited good engineering agreement. Numerical results indicate that the extent of the separated zone has increased and thus show increased separation and reattachment points when compared to experiment. Keywords: Compression corner, Shock wave, Shock wave/Boundary layer Interaction, Supersonic flow 1. Introduction The phenomena of shock-wave / turbulent boundary layer interactions (SWTBLIs) is frequently encountered on the surfaces of aeronautical/aerospace devices like air intake compression ramps of an air breathing propulsion system, jet nozzles, control surfaces and various parts of high speed vehicle. The knowledge of the boundary layer which develops on the walls of these components is essential to optimise the use of these vehicles or equipments. The separation of this boundary layer or its disturbance by a shock wave are two phenomena, which can involve increase of losses of total pressure, high peak heat transfer rates, hence drag and can sometimes even be catastrophic if the shock is strong enough to cause separation [1, 2, 3]. SWTBLIs are a fact of life in the practical world of supersonic flows and that is why valuable attention is given here. A significant amount of work has been under taken to study turbulent shock-separated compression corner flows. Some of the authors who have extensively studied the flow field in a compression corner include: (Delery (1985) [1]), (Daniel Arnal and J.M. Delery (2004) [2]), (A.B. Oliver et al (2007) [3]), and (Settles et al (1994) [4]). This research work focuses on a significantly separated 20 0 2-D compression corner flow-field of Mach number 2.94 to find out the extent of the separation bubble which causes increase in losses and thereby adversely affects the performance. It is hoped that a more physical representation of the SWTBLI in the corner region will improve the capability to predict the surface pressure and wall shear stress for these interactions. 2. Theory The basic interactions between a shock wave and a boundary layer are 2-dimensional and 3-dimensional. 2-D interactions include: the ramp flow, the impinging reflecting shock, and the pressure discontinuity resulting from adaptation to a higher downstream pressure level. 3-D interactions include, swept shock boundary layer Asmelash Haftu Amaha et al. / International Journal of Engineering Science and Technology (IJEST) ISSN : 0975-5462 Vol. 3 No. 3 March 2011 2256