基于空气幕冷却的涡轮端壁改进冷却结构的高温数值验证  

Computational verification of an optimized cooling configuration for turbine endwalls with curtain cooling under high temperatures

作　　者：蔡海扬 吴航 杨星[1,2] 丰镇平 CAI Haiyang;WU Hang;YANG Xing;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)

机构地区：[1]西安交通大学能源与动力工程学院,陕西西安710049 [2]陕西省叶轮机械及动力装备工程实验室,陕西西安710049

出　　处：《推进技术》2025年第1期172-183,共12页Journal of Propulsion Technology

基　　金：国家科技重大专项(2017-III-0003-0027)。

摘　　要：叶栅端壁是航空发动机高压涡轮中流动结构与传热冷却特征最为复杂的区域之一。为提高采用轴向气膜冷却孔布局的原型涡轮端壁的综合冷却性能,本文依据各区域的流场结构和热负荷特性有针对性地进行局部高效气膜冷却强化设计,包括应用端壁通道进口空气幕冷却、扇形气膜孔以及等马赫数线优化布局等方法,使得冷气可以克服横流和二次流的影响,冷气附壁性和覆盖面积大幅提升。本文采用数值模拟方法在某型航空发动机高压涡轮真实进口条件下,验证了端壁改进冷却设计的耦合换热特性。与端壁原型冷却设计相比,在相同冷气消耗量下,端壁改进冷却设计在涡轮真实平均进口温度2150 K下实现了更低且更均匀的金属温度分布,有效消除了原型冷却设计中存在的局部高温热斑,端壁面积平均综合冷却有效度提升了14.2%。叶栅气动分析表明,端壁改进冷却设计还可以降低叶栅出口总压损失,改善叶栅气动性能。冷气流线分布规律表明,空气幕冷却在端壁应用中表现出良好的冷却性能;当空气幕冷气吹风比达到1.85时,可以克服端壁附近涡系的影响而达到尾缘区域形成冷却,并削弱二次流对压力面侧气膜冷却的不利影响;但当空气幕冷气吹风比达到2.12时则会由于过大的射流动量而脱离端壁表面,从而损失部分冷却效果,因此需要合理分配冷气用量。The cascade endwall is one of the regions with the most complex flow structure and heat transfer characteristics in the high-pressure turbine of an aero-engine.In order to improve overall cooling performance of a baseline turbine endwall with axially-arranged film cooling holes,this paper redesigned the film cooling scheme for critical heat transfer regions based on the flow structures and heat loads over the turbine endwall,by applying curtain cooling,fan-shaped film holes,and an iso-Mach number line row pattern for film holes,which enables the film coolant to overcome the influence of crossflows and secondary flows.As a result,the coolant adhesion and coverage area were significantly improved.In this study,conjugate heat transfer characteristics of the optimized cooling scheme for the endwall were verified using numerical simulations for an actual inlet condition for the high-pressure turbine of the aero-engine.The computational results revealed that,compared with the baseline cooling scheme,the optimized cooling scheme achieved a lower and more uniform metal temperature distribution for the same amount of coolant consumption under the inlet temperature of 2150 K,effectively eliminating localized high-temperature hot spots observed in the baseline cooling scheme and increasing area-averaged overall cooling effectiveness on the endwall by 14.2%on average.The reduction of total pressure losses at the vane cascade exit by the optimized cooling scheme demonstrated that it improved the aerodynamic performance of the cascade.The distributions of coolant streamlines showed that curtain cooling provided better cooling performance for the endwall.When the blowing ratio of the curtain cooling reached at 1.85,it overcame the influence of the endwall-nearby vortices,so as to reach and cool the trailing-edge region,and to weaken the adverse influence of the secondary flows on film cooling in the pressure-side region.However,with the excessive blowing ratio of 2.12,the curtain coolant detached from the endwall surface and lo

关 键 词：涡轮端壁 耦合换热 气膜冷却 空气幕冷却 综合冷却有效度 

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]