涡轮叶栅超声速流场流动特征与气膜冷却特性

应用shear strain transport(SST) k-ω 两方程湍流模型,对超声速涡轮叶栅通道内气膜冷却特性进行数值研究,得到不同气膜孔倾角和吹风比下叶栅通道内流场流动特征以及气膜冷却效率的变化规律.在激波入射点附近的气膜射流能够向分离区边界层中补充动量,克服逆压力梯度,有效改善由于激波引起的局部过热.亚声速流动状态下的气膜入射角度对冷却效率的影响能够在较大吹风比下得以体现,而超声速主流状态下,气膜冷却效率与入射角度基本无关,说明亚声速的气膜冷却射流对超声速主流的穿透力要弱于对亚声速主流的穿透力;超声速主流条件下,在激波入射位置的气膜冷却效率要高于激波入射位置下游的气膜冷却效率,这与气膜孔出流在当地的湍流度有关. 

Flow field and film cooling characteristics in supersonic turbine cascade

The shear strain transport(SST) k-ω two-equation turbulence model was employed to numerically simulate the film cooling characteristics in the supersonic turbine cascade channel.The flow characteristics in the cascade channel and the change law of the film cooling efficiency were obtained at different gas film hole angles and blow ratios.The film jet near shock incident point can complement momentum to the boundary layer of separation zone,overcome the adverse pressure gradient and effectively improve the local overheating caused by shock wave.In subsonic flow state,the film incidence angle's impact on the efficiency of cooling can be embodied at larger blow ratio;on the contrary,in supersonic flow state,the film cooling efficiency is almost independent of the incidence angle,showing that the penetrating power of subsonic film to the supersonic mainstream is weaker than that to the subsonic mainstream;in the supersonic mainstream state,the film cooling efficiency of the shock incident location is higher than that of the downstream,which is mainly related to local turbulence intensity of the gas film.