气膜冷却技术在航空燃气涡轮中已经得到了长期的使用。
The film-cooling technique have already got usage of the long-term in the air gas turbine.
对三种不同结构的多斜孔气膜冷却结构进行了数值研究。
Three-dimensional numerical study is conducted to investigate the heat transfer characteristics for the inclined multi-hole film cooling with different configurations.
隔热层烧蚀冷却、气膜冷却是冲压发动机常用的冷却方式。
Adiabatic ablation cooling, film cooling are the main means and have been used widely.
利用液晶测温技术可以很好地反映气膜冷却区域及冷却效果等。
TLC technique can be successfully applied to reflect film cooling effectual area and cooling effect.
利用数值计算的方法研究了非定常尾迹对动叶气膜冷却效率的影响。
Numerical calculations were performed to investigate the effects of unsteady wakes on film cooling.
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.
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.
采用传热传质类比的方法,对多斜孔壁气膜冷却的绝热温比进行了实验研究。
Using the method of heat-mass transfer analogy, experimental research was carried on to investigate adiabatic film effectiveness of inclined multi-hole full-coverage film cooling.
对具有陶瓷隔热涂层的气膜冷却式火焰筒壁面温度和热流提出一种计算方法。
This paper provides a new method suitable for calculating wall temperature and heat flow of ceramic-coated flame tube.
计算的边界条件包括燃气加热,内部对流冷却,气膜冷却和冲击冷却的对流换热。
E. M. The boundary condition consists of convective heat transfer of gas heating, internal cooling, film cooling and impinging cooling.
本文通过数值模拟的方法同时考虑了耦合传热和冷却流通道流动对气膜冷却的影响。
Numerical simulations were used to assess the effect of a conjugate heat transfer and internal coolant flow on film cooling.
采用实验和数值计算相结合的方法,研究了非定常尾迹对叶片头部气膜冷却的影响。
Secondly, the effects of unsteady wakes on film cooling at the leading edge of blade were investigated by experiments and numerical simulations.
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.
根据计算结果重点对前缘气膜冷却复杂三维流动以及整个叶片的气膜冷却特性进行分析。
Analysis of simulation results emphasizes on the complex 3d flow due to cooling near leading edge and blade cooling characteristics.
详细地阐述了对气膜冷却、内部强化换热以及热管冷却等的影响因素,目前的应用状况以及发展前景。
This paper presented influence of factor on film cooling, enhanced heat transfer in tunnel and heat pipe cooling, how to apply and development for these technologies.
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。
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.
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.
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.
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.
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.
通过对几种不同的气膜冷却对流换热系数的计算方法的归纳总结,提出用相对气膜冷却换热系数的方法来计算气膜冷却时的对流换热。
By analyzing the methods of calculating different heat transfer coefficients, the film cooling relative heat transfer coefficients was used to calculate the heat transfer.
研究了基于叶片弦长的主流雷诺数和二次流-主流流量比对冷却效率的影响,并与叶片上圆柱形孔排的气膜冷却效率分布进行了对比。
The influences of re based on the blade chord length and coolant-gas flux ratio are tested, and comparison is made between the effect of the converging-expanding hole rows and cylindrical hole rows.
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。
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.
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.
并从试验中发现了液膜蒸干后气膜段冷却的一些特点。
With it, some characteristics of the gas film cooling after liquid film vaporized-out point have been revealed.
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.
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 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 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.
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