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。
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.
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。
The experimental results are valuable reference for aero-engine turbine blade cooling designers.
本实验选取航空发动机涡轮叶片中段内部冷却通道为研究对象，对不同肋间距的横肋变截面U型通道的换热特性进行了研究。
The experiment has studied the heat transfer characteristic inside U shaped variable cross-section channels with different rib pitches, The channel is the middle part of a gas turbine blade.
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。
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.
涡轮叶片之间冷却空气缝隙的堵塞主要是因为存在高比表面积的超细粒子。
Clogging up of the cooling air slits in the turbine blades is primarily caused by ultrafine particles with a high specific surface area.
本文对涡轮叶片尾缘中具有束腰结构扰流柱的冷却通道的传热和流动阻力特性进行了实验研究，重点研究了雷诺数、扰流柱的束腰比以及不同组合的影响。
The heat transfer and characteristics of the fluid flow in the duct with waist-shaped pin - fin arrays for the trailing edge of the turbine blade were experimentally investigated.
为涡轮叶片的冷却结构设计提供了依据。
实验中对涡轮叶片采用四排气膜孔的冷却方案，气膜孔为圆柱形。
The surface of a gas turbine blade was cooled by four rows of cylindrical holes in the experiment.
按照涡轮叶片复合冷却的全过程，本文由三部分组成。
In view of the whole process of composite cooling on a turbine blade, this paper is composed of three parts.
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。
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.
以带气膜冷却孔的航空燃气涡轮发动机涡轮叶片为研究背景，引入了等效概念对密布小孔结构进行了详细的应力-应变分析。
An equivalent no-perforated plate concept is introduced in the stress-strain analysis for densely-distributed film cooling holes on turbine blades of gas turbine engines.
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。
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.
着重介绍了气膜冷却、涡轮叶片内流冷却技术和气膜孔流量系数的研究进展。
Particular emphasis is given to the development of film cooling, internal cooling technologies and film hole discharge coefficients.
涡轮叶片内部冷却特性直接影响到叶片的壁面温度分布，必须对其展开详细的研究。
The temperature distribution of the nozzle guide vanes are affected greatly by the flow and heat transfer of the internal cooling passages.
本文以航空发动机涡轮叶片冷却为应用背景，对一种基于彻体力场下热驱动理论的新型冷却技术展开了首次研究，并从机理上验证了此新型冷却技术的强化冷却效果。
In this thesis, a new cooling technique, which is based on the theory of thermally driven, was studied at the background of turbine blade's cooling design in aero-engine.
采用数值模拟的方法，研究了涡轮叶片尾缘斜劈缝气膜冷却的流场特性，及其对叶背面尾缘温度分布的影响。
Numerical methods are used to investigate the film cooling effect of turbine guide vane's trailing edge.
本文通过水模拟实验，对涡轮叶片气膜冷却时射流的流场进行了研究。
The flow field of film-cooled turbine vane is investigated in this paper with the help of water simulation.
为了探讨在不同径向角下动量比和湍流度对圆柱形气膜孔流动和换热的影响规律，采用数值模拟方法，研究了涡轮叶片的冷却效率。
The cooling effectiveness of cylindrical gas-film cooling holes was numerically investigated, which was determined at different radial angles, momentum ratio and turbulent intensity.
本文在涡轮叶片新型超级冷却技术机理研究的基础上，设计了旋转条件下带冷却通道的新型超级冷却与涡轮叶片粗糙肋强化冷却对比实验。
A new turbine blade's cooling technique, which was based on the theory of thermally driven, was studied further. Two kinds of experimental models with internal cooling tunnels were designed.
本文在涡轮叶片新型超级冷却技术机理研究的基础上，设计了旋转条件下带冷却通道的新型超级冷却与涡轮叶片粗糙肋强化冷却对比实验。
A new turbine blade's cooling technique, which was based on the theory of thermally driven, was studied further. Two kinds of experimental models with internal cooling tunnels were designed.
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