Effects of Number of Bleed Holes on Shock-Wave/Boundary-Layer Interactions in a Transonic Compressor Stator

An extensive numerical investigation is conducted to characterize the flow separation control in a transonic compressor cascade with a porous bleed. The bleed holes are arranged on the suction surface in a single row, two staggered rows and three staggered rows. For each bleed scheme, five bleed pressure ratios are examined at an inlet Mach number of 1.0. The results indicate that the aerodynamic performance of the cascade is significantly improved by the porous bleed. For the single-row scheme, the maximum reduction in total pressure losses is 57%. For the two-staggered-row and three-staggered-row schemes, there is an optimal bleed pressure ratio of 1.0, and the maximum reductions in total pressure loss are 68% and 75%, respectively. The low loss in the cascade is due to the well-controlled boundary layer. The new local supersonic region created by the bleed hole is the key reason for the improved boundary layer. The vortex induced by side bleeding provides another mechanism for delaying flow separation. Increasing the bleed holes could create multiple local supersonic regions, which reduce the range of the adverse pressure gradient that the boundary layer needs to withstand. This is the reason why cascades with more bleed holes perform better.