Effect of a blade end slot on supersonic compressor cascade hub-corner separation

被引：13
|
作者：
Ma, Jianci [1 ]
Yang, Guang [2 ]
Zhou, Ling [3 ]
Ji, Lucheng [1 ]
Zhang, Chun [3 ]
机构：
[1] Tsinghua Univ, Beijing 100084, Peoples R China
[2] Poly Technol Inc, Beijing 100010, Peoples R China
[3] Beijing Inst Technol, Beijing 100081, Peoples R China
来源：
AEROSPACE SCIENCE AND TECHNOLOGY | 2021年 / 118卷
基金：
中国国家自然科学基金;
关键词：
Blade end slot; Supersonic compressor cascade; Corner separation; Shock wave/boundary layer interaction; Jet flow; FLOW; STALL; PERFORMANCE;
D O I：
10.1016/j.ast.2021.107032
中图分类号：
V [航空、航天];
学科分类号：
08 ; 0825 ;
摘要：
The strong adverse pressure gradient and shock wave/boundary layer interaction that occurs in supersonic compressor cascade corner regions usually leads to large three-dimensional corner separations with extensive total pressure losses. This substantially restricts compressor load performance. In this paper, the effect of a blade end slot on corner separation caused by shock wave/boundary layer interaction was investigated. Because of the pressure difference between the pressure and suction sides, the slot enables one to induce jet flow into the endwall/blade suction-side corner region. Four different blade end slot locations are designed and their influences are investigated numerically. The mechanism of blade end slot influence on corner separation varies by position. The optimal location is slightly downstream of the initial separation point, which improves corner separation control and can reduce the cascade total pressure loss by 9.7% using the currently prescribed slot size. (C) 2021 Elsevier Masson SAS. All rights reserved.
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页数：11
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