摘要
以二维平板模拟高速导弹头罩上的探测窗口,用电加温产生的热气流模拟导弹飞行时的气动加热,对采用气膜冷却控制导弹探测窗口温度的效果及规律进行了研究。着重探讨了气膜冷却临界长度LC与气膜冷却有效度η随射流缝高s、喷射率λ及主射流夹角α等参数的变化规律,并总结出一个相关参数及一些经验公式。研究结果与国外有关文献基本一致。
An experiment using a two dimensional plate wall model of an interceptor forebody configuation is studied to determine the effectiveness of film cooling on temperature control of an infrared window during the high speed flight. The effect of slot height, mass flow ratio, angle of freestream and injection coolant, on the cooling length and cooling effectness, is emphatically studied. In addition, a correlation parameter is obtained. The test results are a general in agreement with the reference date of tunnel at Mach 8.
出处
《南京航空航天大学学报》
EI
CAS
CSCD
北大核心
1998年第5期491-495,共5页
Journal of Nanjing University of Aeronautics & Astronautics
关键词
薄膜冷却
超音速射流
导弹
射流
film cooling
supersonic velocity
missile
injection flow