涡轮叶片非对称扇形气膜孔冷却特性数值研究  

Numerical Investigation on Film Cooling Performance of Turbine Blade with Asymmetrical Fan-shaped Holes

作　　者：徐虹艳[1] 张靖周[1] 姚玉[1] 

机构地区：[1]南京航空航天大学能源与动力学院,南京210016

出　　处：《机械工程学报》2011年第18期152-157,共6页Journal of Mechanical Engineering

摘　　要：针对涡轮导向叶片吸力面和压力面上特定位置上的单排气膜孔,在吹风比为0.44～2.67范围内,数值研究非对称扇形气膜孔的冷却特性。基准对称扇形孔侧向扩展角为20°,后向扩展角为10°。研究结果表明,在扇形总扩展角相等的条件下,非对称型扇形气膜孔的气膜出流穿透能力与对称型扇形气膜孔基本相当,但气膜出流侧向覆盖范围较对称型扇形气膜孔有一定程度的改善,在高吹风比下扇形气膜孔侧向扩展角的影响较为显著。相对而言,非对称扇形气膜孔改善气膜冷却的效果在涡轮叶片压力面侧能得到更好的体现。A numerical study is conducted to investigate the effect of the lateral diffusion angle of fan-shaped holes on the blade film cooling effectiveness at the suction and pressure surfaces.The baseline fan geometry has a 20° lateral diffusion and 10° laidback expansion.Film cooling effectiveness for asymmetrical fan-shaped holes is studied at different blowing ratios（from 0.44 to 2.67）.Under the same total lateral diffusion angle of fan-shaped holes,the coolant jet penetration from the asymmetrical fan-shaped holes is equal to that from baseline fan geometry,but the film coverage in lateral direction is improved a little,which is obvious at the higher blowing ratio.And the asymmetrical fan-shaped holes are suitable to be applied to the pressure surface of turbine blade.

关 键 词：航空宇航推进系统 扇形孔 非对称扇形孔 气膜冷却 涡轮叶片 

分 类 号：V231[航空宇航科学与技术—航空宇航推进理论与工程]