涡轮叶片前缘气膜冷却机理及流场分析

Film cooling mechanism and flow field analysis of leading edge of turbine blade

为探究航空发动机涡轮叶片前缘气膜冷却机理及气膜孔射流对叶片表面气体流场的影响,为涡轮叶片气膜冷却设计和维修提供参考。以普惠PW4084一级HPT叶片为参照构造叶身及内部流道结构、并为叶片设置合适的主流流道;建立叶片前缘气膜冷却模型,在不同吹风比下的气膜冷却进行了数值模拟。结果表明:叶片表面气膜冷却效率具有较强三维性,吸力面前缘处在小吹风比时,射流不能延伸到下一排孔边界无法对孔排间区域进行充分冷却保护,压力面前缘射流贴附壁面后,受到凹曲率表面的导向和约束作用向下游更好地延伸;叶片展向平均气膜冷却效率随吹风比增大而增加,增加速率随吹风比增大而降低;射流使压力面前缘孔排间气流流向沿叶片径向朝叶顶方向偏转、使吸力面叶顶凸曲率处发生气流分离而引起回流,叶顶易产生涡流消耗气体动能。

In order to explore the film cooling mechanism of the leading edge of aero-engine turbine blades and the effect of film hole jets on the gas flow field on the blade surface,and to provide references for the design and maintenance of the film cooling of turbine blades,we set a suitable main flow channel for the blade,taking the Pratt&Whitney PW4084 first-level HPT blade as a reference to construct the blade body and internal flow channel structure.A film cooling model for the leading edge of the blade was established,and the film cooling of this model under different blowing ratios was numerically simulated.The results show that the cooling efficiency of the air film on the blade surface has a strong three-dimensionality.When the suction front edge is at a small blowing ratio,the jet cannot extend to the boundary of the next row of holes,and the area between the rows of holes cannot be fully cooled and protected.After the pressure front edge jet adheres to the wall,it extends to the downstream better under the guidance and restraint of the concave curvature surface.The spanwise average film cooling efficiency increases with the increase of blowing ratio,and the increase rate decreases with the increase of blowing ratio.The jet deflects the air flow direction between the rows of edge holes in front of the pressure along the radial direction of the blade towards the blade top,and causes the air flow separation at the convex curvature of the blade top on the suction surface to cause backflow.The blade top is easy to produce vortex and consume gas kinetic energy.