跨声速轴流压气机径向涡现象与失稳机理  

Radial vortex phenomenon and instability mechanism of transonic axial-flow compressor

作　　者：胡加国[1] 王如根[1] 李少伟[1] 甘甜[1] 

机构地区：[1]空军工程大学航空航天工程学院,西安710038

出　　处：《航空动力学报》2014年第9期2239-2246,共8页Journal of Aerospace Power

摘　　要：对NASA Rotor 37进行数值模拟并与实验结果对比,计算了堵塞点到失稳点的全部工况,详细探究了跨声速轴流压气机附面层分离规律与失稳机理.研究发现:激波后的吸力面附面层中存在一条径向涡,它增强了附面层分离,使部分靠近吸力面的主流向叶尖堆积.随着工况向失稳点推进,压气机转子叶尖出现两块堵塞区,由叶尖泄漏涡与激波作用引起的堵塞区位于压力面前端,由叶尖泄漏涡与径向附面层分离涡耦合作用引起的堵塞区位于吸力面50%弦长后,两块堵塞区的叠加作用最终引起压气机失稳.Numerical investigation was made on NASA Rotor 37 and the result was com- pared with an experiment. Working conditions from blocking point to stall point were calculated to study the rules of boundary layer separation and the instability mechanism of tran- sonic axial-flow compressor. The research discovers a radial vortex behind passage shock wave in suction side boundary layer, which facilitates boundary layer separation and converges part of airflow near suction side to blade tip. With working conditions go to stall point, blade tip passage is jammed by two pieces of blocking zone; one located in the front of blade pressure side is caused by the interference of clearance leakage vortex and passage shock, while the other located on suction side after 50% chord is caused by the interference of radial vortex and leakage flow. The two blocking zones＇ expansion arouses instability of the compressor.

关 键 词：跨声速轴流压气机 流动分离 径向涡 叶尖间隙泄漏 失稳机理 

分 类 号：V231.9[航空宇航科学与技术—航空宇航推进理论与工程]