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

Abstract: 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.

Key words: Nozzle guide vane；Suction side；Heat transfer；Film cooling；Oblique shock wave；Boundary layer separation；Boundary layer transition