基于大涡模拟及实验的压气机叶栅角区分离研究  

Study on Compressor Cascade Corner Separation based on Large Eddy Simulation and Experiment

作　　者：李赫飞 郑群[1] 姜斌[1] 闫巍 LI HE-fei;ZHENG Qun;JIANG Bin;YAN Wei(College of Power and Energy Engineering,Harbin Engineering University,Harbin,China,Post Code:150001)

机构地区：[1]哈尔滨工程大学动力与能源工程学院,黑龙江哈尔滨150001

出　　处：《热能动力工程》2021年第9期117-125,131,共10页Journal of Engineering for Thermal Energy and Power

基　　金：国家科技重大专项(2017-II-0006-0019,2017-I-0009-0010)。

摘　　要：为了研究叶栅内部的流动特性,以及不同攻角下的角区分离模式,对压气机叶栅流场进行了分析。针对两种攻角条件下的平面叶栅模型,采用瞬态雷诺时均(URANS)以及大涡模拟(LES)湍流模型进行了数值模拟研究,并结合叶栅风洞实验验证了数值模拟结果的准确度。对比研究了零攻角以及10°攻角下的叶栅出口流场,叶栅、端壁表面极限流线,以及角区分离结构。研究结果表明:LES能够较好地对角区、尾迹损失进行预测,但URANS在大攻角下的模拟则与实验偏差较大;零攻角下吸力面出现层流分离泡、转捩以及再附现象,而大攻角下吸力面前缘未出现层流分离,而是直接发生转捩以及再附现象;与零攻角相比,10°攻角下的角区分离在展向范围未发生明显变化,在横向范围有小幅度增加,但吸力面附面层分离导致尾迹范围扩大了接近130%,同时总压损失系数提高了接近135%,即大攻角下的主要损失是由吸力面附面层分离以及尾迹损失带来的,而非角区分离。In order to study the flow characteristics inside the cascades and the angle separation modes under different angle of attack, the flow field of compressor cascades was analyzed.URANS and LES turbulence models were used to carry out numerical simulation for planar cascade models under two kinds of incidences, and the accuracy of numerical simulation results was verified by cascade wind tunnel experiment.The flow field at cascade outlet, the limit streamlines on cascade and endwall surfaces and the angular separation structure at 0° and 10° angle of attack are compared and studied.The results show that LES can predict the angle region and wake loss well, but the simulation of URANS has a large deviation from the experiment at high angle attack.Laminar flow separation bubbles, transition and reattachment appear on the suction surface at 0° angle of attack, while transition and reattachment directly occur on the suction leading-edge without laminar flow separation at high angle of attack.Compared with 0° angle of attack, the spanwise range of angle separation at 10° angle of attack does not change obviously, the crosswise range increases slightly.However, the weak range expanded by nearly 130% due to suction boundary layer separation,and the total pressure loss coefficient was increased by nearly 135% at the same time, that is, the main loss at high angle of attack was caused by the separation of the suction boundary layer and the wake loss rather than angle separation.

关 键 词：压气机叶栅 角区分离 风洞实验 大涡模拟 总压损失 

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]