叶尖小翼结合凹槽对跨声速压气机转子间隙流场的影响  

Influence of Blade Tip Winglet Combined With Tip Groove on the Leakage Flow of Transonic Compressor Rotor

作　　者：韩少冰 赵傲 赵毓舟 钟兢军 HAN Shaobing;ZHAO Ao;ZHAO Yuzhou;ZHONG Jingjun(Naval Architecture and Ocean Engineering College,Dalian Maritime University,Dalian 116026,China;Merchant Marine College,Shanghai Maritime University,Shanghai 201306,China)

机构地区：[1]大连海事大学船舶与海洋工程学院,大连116026 [2]上海海事大学商船学院,上海201306

出　　处：《工程热物理学报》2022年第7期1859-1869,共11页Journal of Engineering Thermophysics

基　　金：国家自然科学基金资助项目(No.51406021,No.51436002);辽宁省自然科学基金项目(No.2019-MS-030);大连市高层次人才创新支持计划(No.2018RQ03)。

摘　　要：为了揭示叶尖小翼结合叶顶凹槽对跨声速压气机转子气动性能的影响机理,采用数值模拟方法研究了NASA Rotor35附加不同凹槽型叶尖小翼的气动特性,并在分析Rotor35失稳机制的基础上探究了叶尖小翼结合叶顶凹槽的扩稳机理。研究结果表明:不同深度的凹槽型小翼均使得压气机转子总压比和等熵效率有所降低。随着凹槽深度的增加,压气机转子失速裕度先增大后减小,存在最佳深度的凹槽型小翼使压气机转子失速裕度增加5.4%。凹槽型小翼方案中压气机转子叶尖负荷降低,叶尖泄漏涡强度减弱,转子叶尖通道激波后移,叶尖泄漏涡与激波干涉后发生螺旋型破裂,降低了波/涡干涉所导致的通道堵塞程度及抑制了叶片吸力面附面层分离是其扩稳的机制所在。In order to reveal the influence mechanism of the grooved blade tip winglet on the aerodynamic characteristics of the transonic compressor rotor,a numerical simulation of aerodynamic performance of NASA Rotor35 with different tip winglets was carried out.Based on the instability mechanism of the transonic compressor rotor,the mechanism of stability enhancement by the grooved blade tip winglet was explored.The research results show that the total pressure ratio and isentropic efficiency of the compressor rotor decrease with the different depth of grooved winglets.With the increase of the groove depth,the stall margin of the compressor rotor first increases and then decreases,and the winglet with optimum depth of the groove increases the stall margin by 5.4%.In the scheme of grooved winglet,the tip load of compressor rotor is reduced,the strength of tip leakage vortex is weakened,the shock wave of rotor passage moves backward,and the spiral type vortex breakdown occurs after the interference between tip leakage vortex and shock wave,which reduces the channel blockage caused by wave/vortex interference and inhibits the boundary layer separation of blade suction surface.

关 键 词：叶尖小翼 叶顶凹槽 跨声速压气机转子 间隙流场 数值模拟 

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]