The possibilities of manipulating shock/boundary-layer interactions are demonstrated with regard to an application for hypersonic inlets. Experiments on generic models have been carried out in the hypersonic wind tunnel of the German Aerospace Research Center at Mach 6 and laminar ow conditions. Experimental results were validated by numerical ow simulations using a two-dimensional nite element scheme. In the case of shockinduced boundary-layer separation, it could be shown that the implementation of bleed leads to a reduction of the separation bubble thickness by almost 50%. Further experimental investigations dealt with the achievable reduction of the heat loads on the wall surface depending on the amount and the position of the boundary-layer bleed. These examinations were extended to three-dimensional corner ows and favorable design parameters for a boundary-layer bleed setup were found. Finally, the results obtained using the generic models were transferred to a hypersonic inlet model. There the application of a correctly designed and positioned bleed system showed a signi cant increase of the attainable total pressure recovery.
NomenclatureB A = bleed slot width c f = skin-friction coef cient H = effective inlet height h = height of the internal cross section L A = length of the bleed slot l = model length M = Mach number m = mass ow rate Re = Reynolds number St = Stanton number v = velocity x, y = coordinates b = turning angle of the ow (angle of the shock generator wedge) c A = bleed slot angle D x AS = distance of the bleed slot from the shock impingement position d = boundary-layer thickness d = displacement thickness Subscripts A = bleed G = undisturbed boundary layer at the shock impingement location = freestream conditions
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