Study On Design Method For Pre-compression Blade Profile Of Supersonic Compressor

Compressor is the core component of aero engine,the increase of per-stage pressure ratio can effectively improve the ratio of thrust to weight.The improvement of per-stage pressure ratio can achieve by improving the circumferential speed and the twist speed,high circumferential speed lead to the emergence of supersonic compressor.The static pressure rise can achieve by shock wave,while shock wave also cause energy loss and become one of the core problems of supersonic compressor,and blade profile has a great impact on the shock wave.This paper propose two design method for the supersonic blade profile and carries out an intensive study on the effect of pre-compression design and geometry parameters on shock wave structure and aerodynamic performance.Firstly,DLR-PAV-1.5 supersonic cascade being the numerical modal,the verification of numerical method and grid-independence have been carried out and numerical results shows a great consistence with experimental results,so the numerical method is correct.Based on Levine's unique incidence effect,establishing geometric relationships between airflow parameters and suction surface profiles and proposing the way of suction surface profiles past thickness,and study on two types of slight turning supersonic blades: arch and pre-compression blade profile and it shows that the inlet Mach number and inlet flow angle equal to design value.Then based on the design method,studying the effect of the change of profile maximum thickness,deflection and leading edge radius on the shock wave structure and aerodynamic performance.It shows that in the high inlet Mach number,supersonic airfoil should not be designed as an arch,and pre-compression design can reduce the pre-shock Mach number;the maximum thickness position in the middle and rear part of the airfoil contributes to shock distribution;pressure ratio increases ae the deflection decreases.Moreover,the smaller the leading edge radius is better to the airfoil aerodynamic performance.Proposing a blade design method that directly forms suction surface and pressure surface for the high turning supersonic blade and study on the suction surface pre-compression design and modify pressure surface.It shows that suitable pre-compression design of inlet section of suction surface can reduce the pre-shock Mach number and the separation of suction surface and generates a pressure buffer at the 30% chord length on the suction surface,maximum reduction of loss by 2.32%,but too much pre-compression design will let the leading edge shock changes from one oblique shock to one positive one oblique shocks.After considering a suitable pre-compression design,the length of the inlet section also has a great influence on the shock structure and the length can't be too short.When the length of the inlet section increased from 30% chord length to 45%,shock out of the leading edge and a multi-weak shock system instead of the one strong passage shock,the pre-shock Mach number reduce form 1.3 to 1.2,and multi-shock system result in a smooth deceleration with lower shock losses,maximum reduction of loss by 3.34%.Pressure surface modification will not change shock structure and it will reduce the overall aerodynamic performance of the blade.