跨声叶栅角区激波/附面层干扰的开槽控制措施  

Control of blade end slot on shock wave/boundary layer interaction of transonic cascade corner

作　　者：马健词 周玲 季路成 MA Jianci;ZHOU Ling;JI Lucheng(School of Aerospace Engineering,Beijing Institute of Technology,Beijing 100081,China;Institute for Aero Engine,Tsinghua University,Beijing 100084,China)

机构地区：[1]北京理工大学宇航学院,北京100081 [2]清华大学航空发动机研究院,北京100084

出　　处：《航空动力学报》2021年第7期1377-1387,共11页Journal of Aerospace Power

基　　金：国家自然科学基金(51676015,51976010,52006011);国家科技重大专项(2017-Ⅱ-0006-0020,2017-Ⅱ-0001-0013);北京理工大学青年教师学术启动计划。

摘　　要：提出了使用叶根槽作为一种被动控制手段来控制跨声叶栅的角区分离问题。在压力面与吸力面的压差作用下,叶根槽可产生自发射流,为叶栅吸力面侧角区注入高能流体,从而控制跨声叶栅的角区分离问题。通过数值模拟的方法分析了在不同攻角下叶根槽对压气机叶栅性能的影响及作用机理。结果表明:在小攻角下,叶根槽射流可破坏角区环形涡,从而有效减小跨声叶栅角区分离,提高叶栅的流通能力,改善叶栅性能;在大攻角下,叶根槽射流已不能破坏角区环形涡,但仍能为角区低能流体充能,减弱角区分离,从而拓宽叶栅工作范围。在0°攻角下总压损失系数可降低11.6%,同时叶栅攻角裕度由2°拓宽为3°。A method of using the blade end slot for passive control of the corner separation in a transonic compressor cascade was proposed.Because of the pressure difference between the pressure and suction sides,the blade end slot enables one to induce jet flow into the corner region,thereby restraining the corner separation in transonic cascade.The effect and mechanism of the blade end slot on the performance of the compressor cascade at different incidence angles were investigated numerically.The results demonstrate that,at smaller incidence angle,high-speed jet flow through the blade end slot can destroy the ring vortex in corner region,thereby reducing the corner separation,improving the performance of the cascade;at larger incidence angle,the jet flow can no longer destroy the ring vortex,but still re-energized the low-momentum fluid in the corner region,suppressed the corner separation,and the operation range was extended.The blade end slot reduced the total pressure loss coefficient by 11.6%at 0°incidence angle,and the operation range were increased from 2°to 3°.

关 键 词：叶根槽 跨声叶栅 角区分离 激波/附面层干扰 射流 

分 类 号：V231.3[航空宇航科学与技术—航空宇航推进理论与工程]