跨声速叶栅尾缘激波对层板冷却的影响研究  

Effects of trailing-edge shock wave on laminated cooling in transonic cascade

作　　者：张卓 王春华 刘一帆 张靖周 张树林 王东 ZHANG Zhuo;WANG Chunhua;LIU Yifan;ZHANG Jingzhou;ZHANG Shulin;WANG Dong(Jiangsu Province Key Laboratory of Aerospace Power System,College of Energy and Power Engineering,Nanjing University of Aeronautics and Astronautics,Nanjing 210016,China;Collaborative Innovation Center for Advanced Aero-Engine,Beijing 100191,China;Shenyang Aviation Fuel Technology Co.,Ltd.,Shenyang 110000,China)

机构地区：[1]南京航空航天大学能源与动力学院江苏省航空动力系统重点实验室,江苏南京210016 [2]先进航空发动机协同创新中心,北京100191 [3]沈阳航燃科技有限公司,辽宁沈阳110000

出　　处：《推进技术》2024年第6期177-189,共13页Journal of Propulsion Technology

基　　金：国家科技重大专项(J2019-III-0019-0063)。

摘　　要：采用数值模拟与实验验证相结合的方法针对跨声速叶栅尾缘激波影响下叶片吸力面侧层板冷却结构内外耦合流动传热特性展开研究,获得了来流热力参数和主要结构参数对层板冷却效率及压力损失等的影响规律。研究结果表明:斜激波入射形成的逆压梯度易使得吸力面边界层流动分离;冷却气射流的引入对流动分离有抑制效果,但局部的分离也使层板热侧表面形成局部热斑,冷却效率发生突降(本研究范围内降幅达27.83%);吹风比增大可明显提高冷却效率,但相对压力损失也随之增大;压比由1.89增大到3.67时,叶栅通道内激波作用位置后移,平均冷效提升40.98%,对冷却性能起改善作用。随着气膜孔直径的增大,冷却效果逐渐提升,相对压力损失逐渐增大;冲击孔直径的增大削弱了综合冷却效果,相对压力损失有所下降;在本研究范围内,改变扰流柱直径与高度则对流动换热规律则影响较小。Numerical simulation and experimental verification are combined to investigate the internal and external coupling flow and heat transfer characteristics of the blade suction-side cooling structure under the influence of the trailing edge shock wave of the transonic cascade.The effects of inflow thermal parameters and major structural parameters on the cooling efficiency and pressure loss of the laminate were obtained.The results show that the reverse pressure gradient formed by the oblique shock wave incidence can easily cause the separation of the boundary layer flow on the suction surface.The introduction of cooling air jet has a suppressing effect on flow separation,but the local separation also causes local hot spots on the hot side surface of the laminate,and the cooling efficiency drops sharply(by 27.83%in the present study).Increasing the blowing ratio can significantly improve the cooling efficiency,but the relative pressure loss also increases accordingly.When the pressure ratio increases from 1.89 to 3.67,the position of the shock wave in the cascade channel moves backward,and the average cooling efficiency increases by 40.98%,which improves the cooling performance.As the diameter of the film cooling hole increases,the cooling effect gradually improves,while the relative pressure loss gradually increases.Increasing the diameter of the impingement hole weakens the overall cooling effect,and the relative pressure loss decreases slightly.Within the scope of this study,changing the diameter and height of the turbulence promoter has a relatively small effect on the flow and heat transfer characteristics.

关 键 词：跨声速叶栅 激波 层板冷却 冷却效率 相对压力损失 

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]