| 涡轮叶片前缘气膜冷却换热实验 |
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| 引用本文： | 谭晓茗,朱兴丹,郭文,张靖周,王永明,潘炳华,苏云亮,刘松.涡轮叶片前缘气膜冷却换热实验[J].航空动力学报,2014,29(11):2672-2678. |
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| 作者姓名： | 谭晓茗 朱兴丹 郭文 张靖周 王永明 潘炳华 苏云亮 刘松 |
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| 作者单位： | 南京航空航天大学 能源与动力学院 江苏省航空动力系统重点实验室, 南京 210016;南京航空航天大学 能源与动力学院 江苏省航空动力系统重点实验室, 南京 210016;中国航空工业集团公司 中国燃气涡轮研究院, 成都 610500;南京航空航天大学 能源与动力学院 江苏省航空动力系统重点实验室, 南京 210016;中国航空工业集团公司 中国燃气涡轮研究院, 成都 610500;中国航空工业集团公司 中国燃气涡轮研究院, 成都 610500;中国航空工业集团公司 中国燃气涡轮研究院, 成都 610500;中国航空工业集团公司 中国燃气涡轮研究院, 成都 610500 |
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| 摘 要： | 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.
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| 关 键 词： | 涡轮叶片 气膜冷却 气膜孔倾角 红外热像仪 压力损失系数 |
| 收稿时间： | 2013/10/21 0:00:00 |
Heat transfer experiment on film cooling of turbine blade leading edge |
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| TAN Xiao-ming,ZHU Xing-dan,GUO Wen,ZHANG Jing-zhou,WANG Yong-ming,PANG Bing-hu,SU Yun-liang and LIU Song.Heat transfer experiment on film cooling of turbine blade leading edge[J].Journal of Aerospace Power,2014,29(11):2672-2678. |
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| Authors: | TAN Xiao-ming ZHU Xing-dan GUO Wen ZHANG Jing-zhou WANG Yong-ming PANG Bing-hu SU Yun-liang and LIU Song |
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| Institution: | Jiangsu Province Key Laboratory of Aerospace Power System, College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China;Jiangsu Province Key Laboratory of Aerospace Power System, College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China;China Gas Turbine Establishment, Aviation Industry Corporation of China, Chengdu 610500, China;Jiangsu Province Key Laboratory of Aerospace Power System, College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China;China Gas Turbine Establishment, Aviation Industry Corporation of China, Chengdu 610500, China;China Gas Turbine Establishment, Aviation Industry Corporation of China, Chengdu 610500, China;China Gas Turbine Establishment, Aviation Industry Corporation of China, Chengdu 610500, China;China Gas Turbine Establishment, Aviation Industry Corporation of China, Chengdu 610500, China |
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| Abstract: | 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 76112-142624, 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. |
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| Keywords: | turbine blade film cooling film angle infrared radiation camera pressure loss coefficient |
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