Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow,which always accompanies aerodynamic perfor-mance penalties.A loss reduction method for smearing the passage shock foot via Shock Control Bump (SCB) located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation.The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model.Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes (RANS) method to reveal flow physics of SCB.Comparisons and analyses on simulation results have also been carried out,showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer.It is found that adiabatic effi-ciency and total pressure ratio of improved blade exceeds those of baseline at 95％-100％ design rotor speed,and then slightly worsens with decrease of rotatory speed till both equal below 60％rated speed.The investigated conclusion implies a potential promise for future practical applica-tions of SCB in both transonic and supersonic compressors.