摘要
为了了解涡轮导叶吸力面在跨声速条件下的换热特性,采用数值模拟的方法,分析了出口马赫数对平面叶栅内流场与换热特性的影响,以及气膜出流对吸力面气膜冷却特性的影响。结果表明,跨声速条件下,斜激波导致的逆压梯度导致了吸力面层流边界层分离和转捩;亚声速条件下,吹风比从0.5增大至1.5时,转捩位置前移了约0.1倍弦长;跨声速条件的转捩位置随吹风比增大未发生变化,但是边界层分离现象被抑制,分离泡的尺寸明显变小。在吸力面小吹风比更容易获得更高的冷却效率;边界层的分离导致冷效率分布不同于亚声速条件,在分离区冷却效率迅速降低,在吹风比0.75时降低约50%。
In order to understand the heat transfer characteristic of nozzle guide vane＇s suction side at transonic condition,the effects of exit Mach number on flow field and heat transfer characteristic in plane cascade and the effects of cooling film outflow on film cooling characteristic on suction side have been investigated using CFD method. The results indicate that,the anti-pressure gradient induced by oblique shock wave leads to the laminar boundary layer separation and transition on the suction side at transonic condition. The position of boundary layer transition moves forward about 10% of chord length with the blowing ratios increasing from 0.5 to 1.5 at subsonic condition. The transition position is not changed with the blowing ratio increasing at transonic condition,but the separation is restrained and the size of separation bubble becomes smaller obviously. The higher cooling effectiveness is obtained at smaller blowing ratios on suction side. The cooling effectiveness distribution is different from that of subsonic condition because of the boundary layer separation. The cooling effectiveness decreases rapidly at the separated region. It decreases by nearly 50% when the blowing ratio is 0.75.
出处
《推进技术》
EI
CAS
CSCD
北大核心
2015年第7期1046-1053,共8页
Journal of Propulsion Technology
基金
国家重点基础研究发展规划资助项目(2013CB035702)
关键词
涡轮导叶
吸力面
换热
气膜冷却
斜激波
边界层分离
边界层转捩
Nozzle guide vane
Suction side
Heat transfer
Film cooling
Oblique shock wave
Boundary layer separation
Boundary layer transition
