The Study Of Blade Design Of Aspirated Transonic Compressor

compressor is one of the most key technologies in the moden aeroengine. Pressure ratio of single stage is increasing incessantly, so the whole compressor size and weight was reduce. It is a very important approach to increase the tip velocity of rotor for the high pressure ratio. With the increasing of rotor tip velocity, shock will appear because of the supersonic flow at the tip of rotor. By using CFD technology, this paper studies the design of the transonic aspirated compressor rotor with large turning angles. These mainly including: the effects of the rotor maximum thickness location and the cascades solidity on performance of transonic compressor cascade , the effects of suction flow rate and axial position on the performance of transonic compressor cascades .Boundary layer suction can improve the flows of Cascades, transonic cascade with boundary layer suction can educe higher adverse pressure gradient and increase blade load and improve cascades capacity. The location of the maximum thickness in aspirated transonic compressor is particularly important, with reasonable position of maximum thickness that will reduce the losses of shock and also provide reasonable distribution of load in diffuser. Suction flow rate and axial position on the performance are also very important, reasonable suction location and suction flow rate can remove the high entropy gas, so that the boundary layer separation can be curtailed.The results showed that at the same conditions, boundary removal could increase the blade loading, decrease the separated flow region on the blade suction surface, enhance the turning capacity of the blade and finally decrease the total-pressure loss at the cascade exit. It optimizes the boundary layer behavior. Anyway the potential improvements in pressure ratio and in efficiency that can be attained by boundary layer removal.