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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.
By using the shear strain transport (SST) k-ω two-equation turbulence model, the numerical simulation of the gas-film cooling in the supersonic turbine cascade was conducted. The flow characteristics of the gas flow in the cascade at different gas-film hole angles and blowing ratios were obtained, Film cooling efficiency changes in the vicinity of the shock point of the gas film jet to the separation zone boundary layer to add momentum to overcome the inverse pressure gradient to effectively improve the local overheating caused by the shock wave subsonic flow in the film The impact of incident angle on cooling efficiency can be reflected in a larger blow ratio, while the film cooling efficiency in the supersonic mainstream state is basically independent of the incident angle, which indicates that the subsonic film cooling jet has a weak penetrating force in the supersonic mainstream. In the mainstream of supersonic flow, the cooling efficiency of the gas film at the incident location of the shock wave is higher than that of the gas film downstream of the shock incidence location, which is consistent with the local turbulence Degree related.