改善压气机端区流动的新方法——前缘边条叶片技术  

New approach to improve the endwall flow of compressor—leading edge strake blade technique

作　　者：伊卫林[1] 唐方明[2] 陈志民[1] 季路成[3] 

机构地区：[1]北京理工大学机械与车辆学院,北京100081 [2]中国航空工业集团公司中国航空动力机械研究所,湖南株洲412002 [3]北京理工大学宇航学院,北京100081

出　　处：《航空动力学报》2015年第7期1691-1698,共8页Journal of Aerospace Power

基　　金：国家自然科学基金(51006100;51176012)

摘　　要：针对压气机叶片进气存在端区附面层扭曲而造成局部大攻角问题,借鉴飞机边条翼理论,阐释了LESB(前缘边条叶片)概念,开发了对主叶片施加修型形成边条叶片的造型方法,从而形成LESB技术.为验证其技术效果,选取折转角为60°的NACA65扩压叶栅进行了LESB修型,在利用叶栅试验数据确认CFD模拟精度及掌握使用经验后,对主流区0°攻角、5°攻角带端区附面层扭曲来流条件下NACA65原型叶片及LESB流场进行了数值模拟,对其中流场结构、性能参数及作用机理进行了分析.结果表明:LESB技术能有效组织端区流场,改善压气机性能,15%叶高的LESB修型在0°攻角、5°攻角下改善区域分别可至30%和40%叶高.Targeting the high flow incidence angle induced by inlet endwall boundary layer skew in compressor,the LESB（leading edge strake blade）concept was elucidated with reference to the strake wing theory from aircraft.The modeling method was developed by adding leading edge strake to main blade,and the LESB technique was formed.In order to validate the effect of technique,the LESB modeling was applied to the NACA65 diffuser cascade with turning angle 60°after the numerical method was verified by cascade test data.The performance improvement was compared between original blade and LESB at 0°and 5°incidence angles with endwall boundary layer skew through the flow field structure,performance parameters and action mechanism.The results show that LESB technique can organize effectively the flow field near endwall and improve the compressor performance.The LESBmodeling with 15%span can improve the performances of 30%and 40%span at 0°and 5°incidence angles separately.

关 键 词：压气机 叶轮机 前缘边条叶片 端区 分离与损失 

分 类 号：V231.1[航空宇航科学与技术—航空宇航推进理论与工程]