NASA C3X叶片前缘气膜冷却的数值模拟

对NASA C3X叶片前缘部分建立了合理简化模型,在此基础上数值研究了流动参数和气膜孔斜角角度对前缘气膜绝热冷却效率的影响。结果表明:①在吹风比为0. 5～2. 5,不同吹风比下,以小吹风比的冷却效率较高,吹风比为0. 75时效率最高;主流雷诺数为100 000和200 000时,大雷诺数比小雷诺数下的冷却效率高;主、射流温度比不同时,冷却效率随温度比增大而增大;主流湍流度对冷却效率影响较小。②取射流孔斜角分别为30°、45°和60°,随着射流角度增加冷却效率降低。     

涡轮叶片前缘复合冷却换热性能的实验研究

本文使用由半圆柱前缘和平直尾部组成的钝物体来模拟涡轮叶片前缘,利用高温风洞实验测试平台,对前缘区域进行了内部射流冲击+外部气膜冷却的实验研究,通过红外热像仪测量了叶片表面的温度场分布,分析了气膜吹风比、前缘位置、主流雷诺数对冷却效率的影响规律。实验中吹风比的变化范围为0.24~1.67,主流雷诺数的变化范围为55 420~111 615。实验结果表明:冷却效率随吹风比的增大而增大,在本文实验参数范围内,最佳吹风比约为0.81;位置2处的冷却效率最高;随着主流雷诺数的增大冷却效率有所提高。     

燃气透平叶片前缘气膜冷却效率的试验研究

为了研究燃气透平叶片前缘气膜冷却的传热特性,建立了叶片可视化测试试验台,对叶片前缘区域的冷却效率进行了试验研究,分析了不同吹风比、不同主流雷诺数对叶片前缘区域冷却效率的影响.结果表明：气膜孔附近的冷却效率随吹风比的增大而提高,气膜孔下游的冷却效率随吹风比的增大而降低;冷却效率最高的区域在吸力面上,最低的气膜冷却效率在压力面上产生;低吹风比时主流雷诺数对叶片冷却效率的影响较小;但在高吹风比时,主流雷诺数对叶片前缘气膜孔附近的冷却效率影响较大.     

扇形孔出口宽度对气膜冷却效率影响

数值模拟了不同出口宽度的扇形孔射流在不同吹风比下的气膜冷却效率,侧重于分析涡流结构和涡量对冷却效率的影响。结果表明,出口宽度增大不但减小了涡量,还改变了涡流方向。出口宽度增大到2.5倍入口直径时,孔中心下游冷气流线由燃气指向壁面,有效改进冷气贴附性。气膜孔出口宽度越大,冷却效率越大。随着吹风比的增加,气膜孔出口宽度对冷却效率影响越明显。     

叶片前缘波纹形气膜孔不同参数冷却特性研究

为了解决涡轮叶片前缘的冷却问题,提出结构简单的波纹形气膜孔的概念。对影响孔冷却特性的4个重要参数(吹风比、倾斜角、扩展角和后倾角),采用4因素4水平下的田口方法进行优化,并采用数值计算的方法将其冷却效率与参考孔型(圆柱形气膜孔)进行对比。其中,固定前缘孔间距30 mm,斜向倾角23°,并与试验保持一致。结果表明：吹风比为1、倾斜角为30°、扩展角为20°和后倾角为0°的波纹形气膜孔结构,壁面平均冷却效率达到0.36,冷却效果最佳;同时,随着吹风比的增大,该气膜孔结构前缘壁面平均冷却效率增长趋势明显,且在吹风比为2时冷却气流依旧具有良好的贴壁性能,壁面平均冷却效率超过圆柱形气膜孔82.96%。     

气膜孔形状对冷却效率影响的数值研究

采用控制容积法对三维定常不可压缩雷诺时均紊流方程(N-S方程)进行了离散,并在吹风比M为0.6和1.2的情况下,利用非结构化网格及两层k-e湍流模型,对气膜孔几何形状对涡轮叶片气膜冷却效率的影响进行了数值模拟,得到气膜孔附近的流场分布.结果表明:圆柱形孔的冷却效率随吹风比的增大而明显降低.前向扩张孔的冷却效率优于圆柱形孔,射流在叶高方向上扩展较广,在侧向孔间区域的气膜冷却效率较高.缩放槽缝孔在不同吹风比下的冷却效率均高于圆柱形孔和前向扩张孔,而且在孔下游较远区域,2个孔之间沿叶高方向的气膜覆盖性较好.缩放槽缝孔和前向扩张孔不同程度地抑制了反向涡旋对的产生,因而提高了射流对壁面的贴附性,增强了壁面的冷却效果.     

涡轮叶片尾缘气膜冷却耦合传热的数值模拟

采用耦合方法对燃气轮机涡轮叶片尾缘气膜冷却流场进行数值模拟,并与绝热方法下的结果进行对比,得到不同吹风比和不同吸力面厚度下的气膜冷却规律。结果表明:与绝热方法相比,采用耦合方法时气膜冷却效率曲线更平缓,吸力面温度分布均匀,压力面尾部上方区域温度梯度较大;增大吹风比可以减弱吸力面导热对气膜冷却效率的影响,且能有效抑制流体与壁面的分离;随着吸力面厚度的增加,0.6～0.76等温度比线区域内流体温度发生变化,在劈缝出口下游温度升高,在下游的远距离处温度降低。     

新型缩放槽缝孔气膜冷却效率的数值研究

基于控制容积法对三维定常不可压缩N-S方程进行离散,采用分区域非结构化网格及两层k-ε湍流模型,在吹风比M=0.6～1.5的情况下,对新型缩放槽缝形气膜孔进行了详细的平板气膜冷却数值汁算,得到了在喷孔射流下游处及叶高方向上的气膜冷却效率,并研究了其孔长与孔径比L/D对冷却效率的影响.计算结果表明:高吹风比对缩放槽缝孔沿中心线冷却效率的影响较为复杂;缩放槽缝形气膜孔下游的冷却效率并不随吹风比的变化而单调变化,而是在吹风比为1.0时存在最佳值;在孔口下游较远区域,两个相邻孔之间沿叶高方向的气膜覆盖性较好;缩放槽缝孔的冷却效率随着其孔长与孔径比L/D的增大而增大,当L/D>8时,增加的幅度趋缓.     

气膜孔形状对涡轮叶片气膜冷却效果的影响

基于控制容积法对三维定常不可压缩N-S方程进行离散,采用非结构化网格及两层k-ε湍流模型,在吹风比M为0.6和1.2的情况下,数值模拟了气膜孔几何形状对涡轮叶片气膜冷却效果的影响,得到了气膜孔附近的流场分布.所选孔形为圆柱孔、前向扩张孔、开槽前向扩张孔及新型缩放槽缝孔.结果表明:圆柱孔的冷却效率随着吹风比的增加而显著地降低;开槽前向扩张孔的冷却效率优于圆柱孔和前向扩张孔;缩放槽缝孔在不同吹风比下的冷却效率均高于其它3种孔形,缩放槽缝孔和开槽前向扩张孔不同程度地抑制了反向涡旋对的产生,提高了射流对壁面的贴附性,增强了壁面的冷却效果.     

吹风比和唇板厚度对叶片尾缘气膜冷却的影响研究

为了探究吹风比、唇板厚度对叶片尾缘半劈缝冷却结构气膜冷却特性的影响,采用数值模拟方法对比唇板厚度为4,5和3 mm,吹风比Br为0.5,0.8,1.0和1.5条件下叶片尾缘后台阶上的气膜冷却效率。结果表明：在吹风比Br为0.5时,叶片尾缘后台阶上产生的回流区大,冷气向展向扩散范围广,冷气在近劈缝一端向展向覆盖的较好,由于吹风比小,冷气流速慢,动量小,在后台阶远端燃气与冷气掺混量大,导致冷气冷却能力降低;在大吹风比下（Br=1.5）,冷气流速快,冷气从劈缝射出集中覆盖在劈缝下游处,而肋下游冷气覆盖效果差。唇板厚度影响着唇板出口处形成的回流区,增大唇板厚度将导致半劈缝出口气流分离所产生的涡强度变大,促进燃气与冷气的掺混,降低冷却效率,薄唇板会使尾缘气膜冷却效率显著提高。     

Experimental investigations on overall cooling effect of ribbed channel with air bleeds

An experimental investigation on overall heat transfer performance of a rectangular channel, in which one wall has periodically placed oblique ribs to enhance heat exchange and cylindrical film holes to bleed cooling air, has been carried out in a hot wind tunnel at different mainstream temperatures, hot mainstream Reynolds numbers, coolant Reynolds numbers and blowing ratios. To describe the cooling effect of combined external coolant film with the internal heat convection enhanced by the ribs, the overall cooling effectiveness at the surface exposed in the mainstream with high temperature was calculated by the surface temperatures measured with an infrared thermal imaging system. The total mass flow rate of cooling air through the coolant channel was regulated by a digital mass flow rate controller, and the blowing ratio passing through the total film holes was calculated based on the measurements of another digital-type mass flow meter. The detailed distributions of overall cooling effectiveness show distinctive peaks in heat transfer levels near the film holes, remarkable inner convective heat transfer effect over entire channel surface, and visible conductive heat transfer effect through the channel wall; but only when the coolant Reynolds number is large enough, the oblique rib effect can be detected from the overall cooling effectiveness; and the oblique bleeding hole effect shows the more obvious trend with increasing blowing ratios. Based on the experimental data, the overall cooling effectiveness is correlated as the functions of Re_m (Reynolds number of hot mainstream) and Re_c (Reynolds number of internal coolant flow at entrance) for the parametric conditions examined.     

Turbine endwall film cooling with combustor-turbine interface gap leakage flow: Effect of incidence angle

This paper is focused on the film cooling performance of combustor-turbine leakage flow at off-design condition. The influence of incidence angle on film cooling effectiveness on first-stage vane endwall with combustor-turbine interface slot is studied. A baseline slot configuration is tested in a low speed four-blade cascade comprising a large-scale model of the GE-E³Nozzle Guide Vane (NGV). The slot has a forward expansion angle of 30 deg. to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5 × 10⁵ and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet). The blowing ratio of the coolant through the interface gap varies from M = 0.1 to M = 0.3, while the blowing ratio varies from M = 0.7 to M = 1.3 for the endwall film cooling holes. The film-cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique. The results show that with an increasing blowing ratio the film-cooling effectiveness increases on the endwall. As the incidence angle varies from i = +10 deg. to i = ?10 deg., at low blowing ratio, the averaged film-cooling effectiveness changes slightly near the leading edge suction side area. The case of i = +10 deg. has better film-cooling performance at the downstream part of this region where the axial chord is between 0.15 and 0.25. However, the disadvantage of positive incidence appears when the blowing ratio increases, especially at the upstream part of near suction side region where the axial chord is between 0 and 0.15. On the main passage endwall surface, as the incidence angle changes from i = +10 deg. to i = ?10 deg., the averaged film-cooling effectiveness changes slightly and the negative incidence appears to be more effective for the downstream part film cooling of the endwall surface where the axial chord is between 0.6 and 0.8.     

Research on cooling effectiveness in stepped slot film cooling vane

As one of the most important developments in air cooling technology for hot parts of the aero-engine,film cooling technology has been widely used.Film cooling hole structure exists mainly in areas that have high temperature,uneven cooling effectiveness issues when in actual use.The first stage turbine vanes of the aero-engine consume the largest portion of cooling air,thereby the research on reducing the amount of cooling air has the greatest potential.A new stepped slot film cooling vane with a high cooling effectiveness and a high cooling uniformity was researched initially.Through numerical methods,the affecting factors of the cooling effectiveness of a vane with the stepped slot film cooling structure were researched.This paper focuses on the cooling effectiveness and the pressure loss in different blowing ratio conditions,then the most reasonable and scientific structure parameter can be obtained by analyzing the results.The results show that 1.0 mm is the optimum slot width and 10.0 is the most reasonable blowing ratio.Under this condition,the vane achieved the best cooling result and the highest cooling effectiveness,and also retained a low pressure loss.     

Investigation of the Conjugate Heat Transfer and Flow Field for a Flat Plate with Combined Film and Impingement Cooling

Film cooling combined with internal impingement cooling is one of the most effective technologies to protect the gas turbine vanes and blades from the hot gas. In this study, conjugate heat transfer CFD study was undertaken for a flat plate with combined film cooling and impingement cooling. An experiment on conjugate heat transfer of a flat plate with combined film and impingement cooling was performed to validate the code. Then the effects of several parameters including Biot number, blowing ratio, film hole shape and impingement hole diameter on the overall cooling effectiveness were numerically studied. The results show that for a specific combined cooling scheme and a given blowing ratio, the coolant potential can be reasonably allocated to the internal and the external cooling to achieve the overall cooling effectiveness. As the blowing ratio increases, the overall cooling effectiveness trends to reach a maximum value. For different film hole geometrical, the maximum values of the overall cooling effectiveness at high blowing ratio approximate to the same value. At a given mass flow rate of coolant, the increase of the impingement hole diameter leads to the reduction of the overall cooling effectiveness.     

Unsteady numerical investigation of back‐step three‐dimensional slots on film cooling effectiveness

A numerical investigation was carried out to study the aerodynamics and cooling effects of a trailing edge. For a greater understanding and to learn more details, an unsteady numerical model has been proposed based on a steady model. The unsteady numerical simulation was conducted under different blowing ratios (0.5, 2.0) to show their effects on film cooling effectiveness. The computational results show that the turbulence intensity downstream of a trailing edge outlet tends to be enhanced by the unsteady flow effect. Film cooling effectiveness of an unsteady model is weaker than that obtained by a steady model. When the blowing ratio is 0.5, the mixing and intersecting of the main flow and secondary flow is rapid and fierce; with the increase of blowing ratio up to 2.0, the secondary flow plays a dominant role on the flow characteristics near the outlet of the trailing edge. The mixing and intersecting of the main flow and secondary flow become smooth. The unsteady computational results agree better with the experiment results than those of steady computation. © 2008 Wiley Periodicals, Inc. Heat Trans Asian Res; Published online in Wiley InterScience ( www.interscience.wiley.com ). DOI 10.1002/htj.20236     

不同吹风比下双出口孔射流气膜冷却数值模拟计算

为了获得吹风比对新型气膜冷却孔冷却效率的影响规律,利用Fluent软件求解Navier-Stokes方程,对吹风比分别为0.5、1.0、1.5和2.0时单入口-双出口孔射流冷却效率进行了数值模拟计算,得到了不同吹风比下的流场和冷却效率.结果表明：吹风比对冷却效率有很大影响;随着吹风比的提高,不同次孔方位角下的冷却效率变化规律也不相同;当次孔方位角γ=30°时,吹风比为1.0时的冷却效率最高;当γ=45°时,冷却效率随着吹风比提高而提高;当γ=60°时,冷却效率随着吹风比提高而降低;在研究高吹风比对气膜冷却效率的影响时,γ=45°最佳.     

透平叶顶气膜冷却效果数值研究

采用三维数值模拟方法,研究了GE E3发动机第一级透平动叶叶顶间隙内的气膜流动与换热特性,评估了气膜吹风比M分别为0.5、1.0和1.5时,对叶顶换热系数以及冷却效率的影响.计算结果表明:叶顶气膜冷却空气改变了叶顶泄漏流动特性,随着吹风比的增加,叶顶间隙内的泄漏流动区域不断缩小,从而导致叶顶间隙泄漏量不断减小;随着气膜冷却吹风比的增大,叶顶平均换热系数逐步降低;在M=1时,冷却效果最佳.     

几何结构对后台阶三维缝隙换热影响的研究

针对涡轮叶片尾缘冷却结构特点,建立了后台阶三维缝隙结构气膜冷却特性试验台;测量了缝隙中心和肋中心下游气膜换热系数的局部分布;研究了不同几何结构参数对换热系数的影响规律,其中缝宽/肋宽比b1/b2的变化参数是0.67、1.0及1.5,缝高/唇厚比H/d的变化参数是0.5、1.0及2.0。试验结果表明:缝后中心线与肋后中心线的换热系数随轴向距离的增加其总的趋势都是减小;换热系数随着缝高/唇厚比的减小,曲线变得比较平缓;换热系数在不同的缝宽/肋宽比时,曲线基本整体地上下移动,但吹风比不同Bc不同时,曲线移动幅度区别较大。     

动叶端壁离散孔气膜冷却的数值研究

基于某F级燃气轮机第一级动叶栅的数值模拟,以实现动叶端壁气膜冷却全覆盖为目标,分析定常下动叶端壁的流动与传热特征,拟综合考虑端壁二次流结构特征与级间封严冷气泄漏流的影响,将端壁划分为四个具有不同流动传热特征的区域,并据此设计了叶根端壁仅13孔数的离散气膜孔布置方案各区域采取不同的冷却方式根据不同吹风比下的研究结果发现:吹风比为0.75时端壁冷却有效度均值在0.2以上,实现了全端壁冷却的目标;前缘附近端壁冷却效果受吹风比影响显著,吹风比在0.75以上时冷却有效度达到0.5以上;除近压力面区域,气膜冷却效果随吹风比的增大而提高。     

尾迹对气膜冷却影响的三维非定常数值模拟

采用数值计算的方法，对带有前导静叶的涡轮动叶气膜冷却进行三维非定常数值计算，研究了静叶非定常尾迹对动叶气膜孔周围流场及绝热冷却效率的影响，同时比较了在吹风比M=0．8和M=1．5时动叶表面气膜冷却效果的变化。结果表明，静叶尾迹使动叶压力面和吸力面上冷却气流的流场发生改变，降低了叶片表面的气膜冷却效果。另外，吹风比从0．8增大到1．5时，静叶非定常尾迹对动叶气膜冷却效果的影响降低。