Experimental and computational investigations of shock wave – boundary layer interaction on a transonic airfoil with and without bump

In this study, experimental and numerical studies have been performed to investigate shock wave-boundary layer interaction (SWBLI) on a wing at transonic speed. The wing has rectangular planform and NACA 64-A-307 airfoil in cross-section. Experiments are conducted at angles of attack-2°, 0°, 8° and freestream Mach numbers 0.8 & 0.85 for both 2D-3D wing configurations. In the experimental part, shock wave patterns and separated flow induced by SWBLI are observed by the help of high-speed schlieren method. Numerical computations are employed using open source computational fluid dynamics (CFD) software OpenFOAM^® . In addition to experimental studies, wings with bump surfaces are also examined in computational analyses. In pre-processing stage, the computational domain and flow conditions are defined. Afterwards, analyses are run with turbulence model and computational results such as density gradients (numerical schlieren) are compared with the experimental results. In the light of computational and experimental outcomes; each parameter affecting flow field and wing performance are discussed.