涡轮叶片气膜冷却的研究

RESEARCH ON FILM-COOLING OF TURBINE BLADE

本文针对涡轮叶片离散气膜孔结构及叶片表面具有不同曲率的特点 ,采用传质比拟、数值计算及可视化技术 ,通过比较平板、弯曲壁面及涡轮叶片的气膜冷却流动及传热特性 ,系统地研究了离散孔三维效应、壁面曲率及压力梯度诸因素对叶片气膜冷却的影响。在此基础上 ,提出了估算叶片头部、中弧区及尾缘气膜冷却有效温比的准则公式。

The effects of discrete film-cooling holes,wall curvature and the pressure gradient on the turbine blade film cooling are systematically studied by means of heat-mass analogy,numerical methods and flow visualization techniques.The flow and heat transfer performances of the cooling film over the flat plate,curved wall and blade surfaces are investigated.It is proven that the characteristics of discrete-hole film cooling are quite different from those of continuous-slot film cooling in the vicinity of the injection holes because the coolant jets from holes tend to separate from the cooled surfaces. An optimal blowing ratio is expected for ideal distribution of film cooling effectiveness.Wall curvature exerted an additional force on the cooling film.When momentum blowing ratio I(=ρ 2U 2/ρ ∞U 2 ∞)>1,this force tends to seperate the film from the convex surface and to keep it close up to the concave surface,but when I<1,it acts contrary.For a given film cooling scheme,the effectiveness of turbine blade deponds on the blowing ratio,wall curvature and the pressure gradient.Some of the correlation formulas,exprimental data and computational results are presented to evaluate the film cooling effectivenesses and heat transfer coefficients at the leading edge,mid-chord and tailing edge of the turbine blade.