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针对某型涡轮叶片放大模型的前缘冷却结构气膜冷却效果开展了细致的实验研究,利用红外热像仪测量了叶片表面的温度场分布,分析了前缘的气膜孔倾角、吹风比、主流雷诺数等参数对绝热冷却效率和压力损失的影响.实验中前缘的3排气膜孔倾角变化范围是35°~90°,主流雷诺数变化范围是76 112~142 624,吹风比变化范围是0.44~2.64.结果表明:气膜孔倾角越小,前缘驻点附近的气膜覆盖效果越好;气膜孔倾角为45°的叶片压力损失系数最小,气膜孔倾角为75°的叶片压力损失系数最大;主流雷诺数增大,绝热冷却效率下降,压力损失系数增加;吹风比增大到1.32时,绝热冷却效率达到最大,吹风比再增大绝热冷却效率反而下降.
Aiming at the cooling effect of the cooling film on the leading edge of a turbine blade model, a detailed experimental study has been carried out. The temperature field distribution on the blade surface has been measured by infrared thermography. The inclination angle of the film hole, The main Reynolds number and other parameters on the adiabatic cooling efficiency and pressure loss.In the experiment the leading edge of the three rows of gas vent hole angle variation range is 35 ° ~ 90 °, the mainstream Reynolds number range is 76 112 ~ 142 624, the hair blowing ratio changes The range of 0.44 ~ 2.64.The results show that: the smaller the angle of the gas film hole, the better the film coverage effect near the stagnation point of the front edge; the smallest of the pressure loss coefficient of the film with the film hole angle of 45 ° and the angle of the film hole angle of 75 ° The maximum pressure loss coefficient is the largest; the mainstream Reynolds number increases, the adiabatic cooling efficiency decreases, the pressure loss coefficient increases; when the blowing ratio increases to 1.32, the adiabatic cooling efficiency reaches the maximum, the blowing ratio increases and the adiabatic cooling efficiency decreases.