An experimental investigation of three-dimensional shock control bumps applied to transonic airfoils

Previous research on the behavior of shock control bumps (SCBs) on transonic airfoils has been largely limited to numerical studies, with experimental investigations primarily limited to basic ow elds in small wind tunnels. This paper examines the possibility of simulating the conditions on a wing in a blow-down supersonic wind tunnel to allow a relatively inexpensive and simple experimental study of the fundamental physics of SCBs. The main requirements are a post-shock adverse pressure gradient and a representative incoming turbulent boundary layer. Tests were carried out at a Mach number of 1.3 using a variety of measurement techniques and the results compared with computations. The ow conditions in the proposed wind tunnel set-up were highly comparable with the computational results for a representative ight condition on a typical transonic airfoil. A contour SCB was tested in the new wind tunnel set-up, and its ow features are discussed. It was found that the SCB brought about an improved total pressure recovery in the boundary layer by the end of the di user (corresponding to the airfoil trailing edge) and this was attributed to a vortical wake generated by baroclinic e ects. This provides direct evidence in support of the suggestion that SCBs could also be used as a form of boundary layer control.