利用空气幕冷却提升涡轮叶栅端壁气膜冷却有效度的实验研究  

An Experimental Investigation of Improving Film Cooling Effectiveness for Turbine Vane Endwalls Using Curtain Cooling

作　　者：杨星 赵强[3] 吴航 丰镇平[1,2] YANG Xing;ZHAO Qiang;WU Hang;FENG Zhenping(School of Energy and Power Engineering,Xi’an Jiaotong University,Xi’an 710049,China;Shaanxi Engineering Laboratory of Turbomachinery and Power Equipment,Xi’an 710049,China;Xi’an Aerospace Propulsion Institute,Xi’an 710100,China)

机构地区：[1]西安交通大学能源与动力学院,西安710049 [2]陕西省叶轮机械及动力装备工程实验室,西安710049 [3]西安航天动力研究所,西安710100

出　　处：《西安交通大学学报》2023年第8期1-10,共10页Journal of Xi'an Jiaotong University

基　　金：国家科技重大专项资助项目(2017-Ⅲ-0003-0027)。

摘　　要：针对涡轮叶栅压力面侧由于横向压力梯度及复杂二次流导致的冷却难题,通过在端壁通道进口靠近压力面侧布置两排离散气膜孔,形成具有独特冷却特征的空气幕冷却,采用压力敏感漆测量技术(PSP)详细研究了空气幕冷却在端壁表面的冷却分布规律及其对通道中离散气膜孔冷气射流的影响。在此基础上,将空气幕冷却应用于端壁气膜冷却形成了改进设计方案。实验结果表明,空气幕冷却具有与通道中离散气膜冷却完全不同的冷却特征,其几乎不受叶栅通道中横向压力梯度的影响,可以有效冷却端壁通道的压力面侧甚至喉部及下游区域;随着冷气量的增大,空气幕冷却的冷却效果不断增强,并会提高下游端壁通道中离散气膜孔的冷却性能;与端壁原型气膜冷却方案相比,在冷气量相同的情况下,改进方案将端壁表面的面积平均气膜冷却有效度提高了33%。叶栅出口的流场结构表明,改进方案还可以削弱叶栅的气动损失。Due to circumferential pressure gradients and complicated secondary flows,pressureside regions of turbine endwalls are one of the most difficulttocool regions.This study aims to solve this cooling problem by placing two rows of discrete film holes in the upstream region of the endwall passage inlet near the pressure side to obtain curtain cooling with unique cooling characteristics.A pressure sensitive paint(PSP)technique was implemented to examine film cooling patterns of curtain cooling over the endwall surfaces and its effects on discrete film injection within the passage in detail.Furthermore,an optimized endwall film cooling scheme was obtained by applying curtain cooling to the endwall.The experimental results reveal that curtain cooling has quite different cooling patterns compared with discrete film cooling within the passage.The curtain cooling was slightly influenced by the pressure gradients circumferentially across the endwall passage,resulting in efficient cooling for the pressureside region and even for the endwall passage throat and its downstream regions.Increasing coolant rates enhances film cooling effectiveness for curtain cooling and improves film cooling performance of the discrete film holes within the endwall passage.The optimized cooling scheme improves film cooling effectiveness over the endwall by 33%relative to a baseline cooling configuration.Additionally,flow structures at the vane cascade exit demonstrate that the optimized cooling configuration reduces the cascade aerodynamic losses,somewhat.

关 键 词：航空发动机 涡轮端壁 气膜冷却 空气幕冷却 气动损失 

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]