Experimental Study on Impingement and Film Overall Cooling Characteristics of Turbine Shroud  

作　　者：LI Ziqiang WANG Longfei MAO Junkui BI Shuai WANG Feilong 

机构地区：[1]College of Energy and Power Engineering,Nanjing University of Aeronautics and Astronautics,Nanjing 210016,China

出　　处：《Journal of Thermal Science》2022年第1期239-250,共12页热科学学报（英文版）

基　　金：The authors are grateful for the supported by China Postdoctoral Science Foundation(NO.2020TQ0147);Natural Science Foundation of Jiangsu Province(NO.BK20200454);the Fundamental Research Funds for the Central Universities(NO.NJ2020014).

摘　　要：Based on the variable characteristics of the actual operating conditions of the turbine shroud and the purpose of improving the cooling effect of the turbine shroud,this paper builds a test system of the impingement-film cooling shroud with two gas inlet angles(90°,167°).The effects of film cooling hole arrangement,gas inlet angle,blowing ratio(0.7,1.0,1.5,2.0,2.5,3.0)and temperature ratio(1.2,1.3,1.4,1.5,1.6)on the cooling characteristics of the impingement-film cooling shroud were experimentally studied by infrared temperature measurement technology,especially the effects of gas inlet angle and temperature ratio.The results showed that the film covering effect of the film cooling hole vertical or the same direction of the high-temperature gas incoming flow is better than the film covering effect of the reverse direction with the incoming flow,and the optimal arrangement of film cooling holes can improve the cooling effectiveness of the shroud.Compared with 90°intake gas,the film coverage area on the shroud surface of the 167°intake gas is expanded,and the surface average overall cooling effectiveness is increased by 1.03%to 12.6%.The overall cooling effectiveness of turbine shroud increases with the increase of blowing ratio,which increases the flow rate and pressure of cooling gas,and the corresponding increase rate is between 1.04%and 9.96%.However,the increase in the temperature ratio increases the mainstream heating capacity,resulting in a decrease in the cooling effectiveness of the shroud,with a maximum reduction rate of 11.04%.

关 键 词：turbine shroud overall cooling effectiveness impingement-film cooling blowing ratio temperature ratio 

分 类 号：V263.3[航空宇航科学与技术—航空宇航制造工程]