This study can be useful in turbine blade cooling design.
为涡轮叶片的冷却结构设计提供了依据。
The experimental results are valuable reference for aero-engine turbine blade cooling designers.
本文的实验结果对于航空发动机涡轮叶片冷却结构设计具有参考价值。
This essay analyzes the cooling principle and the mode of operation of the M701F gas turbine blade cooling technology to verify its efficiency in enhancing the safety of the gas turbine.
通过分析三菱M701F燃气轮机叶片冷却技术的冷却原理和运行方式，为燃气轮机的安全高效运行提供参考。
Film cooling of the surface of a gas turbine blade was studied in a large-scale low-speed opening wind tunnel according to actual requirement of the design of aero-engine.
根据航空发动机设计的实际需要，在大尺寸低速叶栅传热风洞中对涡轮叶片表面的气膜冷却进行了综合性实验研究。
Hot wire measurements and flow visualization are presented for studying the turbulent flow field over a flat gas turbine film cooling blade with lateral expanded holes.
报道了用热线风速仪和流动可视化技术对带有横向扩展型孔的燃气轮机气膜冷却叶片紊流流场进行的研究。
Based on the gas turbine blade's practical cooling condition, a blade cooling computation model is introduced and perfected, and its validity is verified using reported data.
根据燃气轮机透平叶片冷却的实际情况，引入并完善叶片冷却的计算模型，并采用文献报道的数据验证计算模型的正确性。
Experimental investigation of the turbine blade local heat transfer characteristics at the inner side of the mid-chord region near film holes was conducted, with and without impingement cooling.
对叶片弦中区内部有、无冲击射流的气膜出流冷却方式中，冷气侧气膜孔局部换热特性进行了实验研究。
The cooling technique of the blade inner passage plays a significant role in the design of blades for gas turbine.
叶片内部通道冷却技术在航空发动机叶片的设计中占有重要的位置。
Several problems involved in the prediction of heat transfer coefficient on the turbine blade profiles without film cooling are studied in this paper for the purpose of improving the accuracy.
本文用数值方法计算了无气膜冷却涡轮叶片上包括前驻点的整个型面上的换热系数。
In view of the whole process of composite cooling on a turbine blade, this paper is composed of three parts.
按照涡轮叶片复合冷却的全过程，本文由三部分组成。
This essay mainly specifies the impacts of the blade cooling technique on the turbine initial temperature.
着重叙述的是叶片的冷却技术对提高透平初温的影响。
The results show that the cooling effectiveness on concave side is better than suction side with laying film cooling holes in concave side at leading edges of turbine blade.
结果表明，叶片前缘压力面侧布置的气膜孔对叶片压力面有很好的冷却效果；
The numerical study was carried out on the trailing edge cooling of turbine blade using the circular pin fin that was arranged in cross type.
对一个有交叉排列圆形扰流柱的涡轮叶片尾缘冷却方案进行了数值研究。
Film cooling effectiveness of hole rows on leading edge of turbine blade has been studied experimentally. The model is a blunt body with a half cylinder leading edge with two flat side walls.
采用放大的半圆柱状表面模拟涡轮叶片前缘的形状，对叶片前缘单排及两排圆柱形孔的气膜冷却效率进行了测量。
Compared to the channel with uniform section area, those with various section areas are much similar to the internal cooling structure of the turbine blade.
与等截面回转通道相比较，变截面通道更接近叶片内冷通道原型。
Compared to the channel with uniform section area, those with various section areas are much similar to the internal cooling structure of the turbine blade.
与等截面回转通道相比较，变截面通道更接近叶片内冷通道原型。
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