采用计算流体动力学软件ANSYS CFX11.0、以NASA跨声速透平第一级动叶为研究对象,对带气膜冷却孔的叶栅近前缘端壁区域的流动和换热特性进行了研究,计算获得了3种气膜孔分布条件下,吹风比分别为0.3、0.5、0.7以及孔径分别为1 mm、1.5 mm时叶栅端壁处的流场结构和斯坦顿数分布.计算结果表明:气膜孔的数目及分布对端壁换热性能和换热均匀性有显著影响,减小孔间距与孔径的比值可以降低前缘端壁的换热系数、提高端壁换热的均匀性;吹风比对冷却流的作用范围和贴壁性有很大影响,所研究的3种吹风比中,吹风比为0.5时壁面换热系数最小,吹风比为0.7时换热系数最大;当吹风比保持0.5不变且气膜孔的直径由1 mm增大到1.5mm时,冷却流在端壁上影响的距离增加,相邻冷却流之间区域的换热强度降低.该结果可为透平动叶端壁换热特性的改善和气膜冷却特性的提高提供参考.%Taking the first stage rotor blade of NASA transonic turbine as the research object and using computational fluid dynamics software ANSYS-CFX11.0,the flow and heat transfer performance of the endwall near the blade leading edge was investigated.The flow field structure and Stanton number distribution near the blade endwall were numerically analyzed at three blowing ratios and two different cooling-hole diameters with three kinds of cooling-hole arrangements.The numerical results obtained indicate that the heat transfer performance and heat transfer uniformity on the blade endwall are greatly affected by the number and layout of film-cooling holes.The heat transfer coefficient on the blade endwall near leading edge is reduced by decreasing the hole-pitch to hole-diameter ratio.However,the heat transfer uniformity on the endwall is increased correspondingly.It is also seen that the influencing area and adherence property of the coolant to the endwall are significantly influenced by the blowing ratio.Among the studied three blowing ratios,the lowest heat transfer rate on the blade endwall near leading edge is obtained at the blowing ratio 0.5,while the largest heat transfer rate is derived at a blowing ratio of 0.7.At a fixed blowing ratio 0.5,the coverage of the ejected coolant downstream the cooling hole is enlarged as the cooling-hole diameter increases from 1mm to 1.5 mm.However,the heat transfer rate on the endwall between adjacent cooling jets is decreased subsequently.The present research results can provide a reference to improve the heat transfer performance of the turbine rotor blade endwall as well as the film cooling effect.
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