Heat transfer experiment on film cooling of turbine blade leading edge

被引：0

作者：

Tan, Xiao-Ming ^{[1
]}

Zhu, Xing-Dan ^{[1
]}

Guo, Wen ^{[2
]}

Zhang, Jing-Zhou ^{[1
]}

Wang, Yong-Ming ^{[2
]}

Pang, Bing-Hua ^{[2
]}

Su, Yun-Liang ^{[2
]}

Liu, Song ^{[2
]}

机构：

[1] Jiangsu Province Key Laboratory of Aerospace Power System, College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing

[2] China Gas Turbine Establishment, Aviation Industry Corporation of China, Chengdu

来源：

Hangkong Dongli Xuebao/Journal of Aerospace Power | 2014年 / 29卷 / 11期

关键词：

Film angle; Film cooling; Infrared radiation camera; Pressure loss coefficient; Turbine blade;

D O I：

10.13224/j.cnki.jasp.2014.11.018

中图分类号：

学科分类号：

摘要：

Detailed experimental study on film cooling effect of one enlarged model of turbine blade leading edge cooling structure was carried out. The surface temperature distribution of blade was captured by the infrared radiation camera. The influence of adiabatic cooling efficiency and pressure loss were analyzed by different film angles of leading edge, blow ratios, main flow Reynolds numbers. In the experiment, the range of three-row film angle on leading edge was 35 degree to 90 degree; the range of main flow Reynolds number was 76 112-142 624, and the range of blow ratio was 0.44-2.64. The results show that: the film cooling on the stagnation region of leading edge is getting better with the decrease of film angle; the pressure loss coefficient is lowest with film angle of 45 degree and highest with film angle of 75 degree; with increase of main flow Reynolds number, the adiabatic cooling efficiency decreases, and the pressure loss coefficient increases; the adiabatic cooling efficiency reaches to the maximum when blow ratio increases to 1.32 and then decreases when blow ratio keeps increasing. ©, 2014, BUAA Press. All right reserved.

引用

页码：2672 / 2678

页数：6

共 16 条

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