In order to effectively reduce both losses induced by the first passage shock and resulted from boundary layer separation inside a supersonic compressor cascade at variable inlet Mach numbers

a type of method employing a new kind of 2-D bump structure with straight transition region on the suction side of the supersonic compressor cascade was presented to control the shock-boundary layer interaction. Numerical simulations and then detailed analyses to demonstrate the control mechanism of the 2-D bump on shocks and shock-boundary layer nteractions were conducted as well to investigate the influences of geometric dimensions and positions of the 2-D bump on control results. The investigations show that the compression wave system induced from concave windward surface of the 2-D bump effectively weakens the intensity of the first passage shock and that the expansion wave system produced by the transition region of the bump accelerates the flow in boundary layer

both of which suppress the local flow separation significantly and make the separation reattach in advance. While the first passage shock inside a supersonic cascade impinged on the straight section of the bump

optimal control effect was achieved at dimensionless height of the bump which retained 0.35 times of boundary layer thickness

and also at dimensionless length of windward and leeward side which remained 4 and 5 times of transition region respectively. Moreover

compared with the cascade without 2-D bump

the level for total pressure loss reduced by the bump reached 4.6% with the free stream Mach number 1.75

while the aerodynamic loss of the supersonic compressor cascade with bump was still better than that without bump when the inlet Mach number varied from 1.65 to 1.80.