复合角度气膜冷却叶片的数值模拟

采用Realizable k-ε紊流模型,并结合Simplec算法和有限元法对体积进行离散,研究了静止叶栅前缘射流孔在2种复合角度α=30°、β=45°和α=135°、β=45°,不同吹风比M=0.5、M=1.0、M=1.5,主流温度为T∞=293 K时的压力面和吸力面温度场,并分析了典型工况下的气膜冷却效率.结果表明:相同复合角度、不同吹风比的压力面和吸力面冷却效率曲线变化趋势一致;前缘复合角度射流对整个吸力面的冷却效率有较大影响,随着吹风比的增大,冷却效率提高;当前缘复合角度为α=30°、β=45°时,压力面冷却效率随着吹风比的增大而提高,而当前缘复合角度为α=135°、β=45°时,随着吹风比的增大,压力面的冷却效率降低.     

入射角度对缩放槽缝孔气膜冷却效果的影响

基于有限体积法对三维定常不可压缩N-S方程进行离散,采用两层k-ε湍流模型,在吹风比M为0.5、1.0、1.5和2.0的情况下,数值研究了入射角度(α=25°、45°和60°)对缩放槽缝孔气膜冷却效果的影响,对不同入射角度的气膜冷却整体效果进行了对比分析。结果表明:在任何吹风比的情况下,α=25°喷射时的冷却效率高于其它喷射角的冷却效率,并且随着吹风比的增大,小角度喷射优于其它喷射角的趋势也越来越大;入射角度为45°和60°的气膜孔沿孔排下游的冷却效率在下降过程中重新升高然后又继续下降,60°喷射角的上升趋势略大于45°角的上升趋势;大角度喷射时,在气膜孔下方生成了强度较强的反向涡旋对,两个旋涡之间的距离较近,冷却气流的附壁性较差,冷却效率较低。小角度喷射时,所生成的反向涡旋对与大角度喷射相比尺度较小、强度较弱,冷气射流对主流的阻碍作用比较小,冷却效率较高。     

涡轮静叶复合角度气膜冷却孔排布置优化的数值研究

采用Realizable k-ε两方程模型,通过改变射流入射角度和吹风比,对具有8排复合角度射流孔的涡轮静叶片表面温度场进行数值研究,确定各个孔排的最佳吹风比;对复合角度孔排的布置进行调整,得到较好的气膜冷却方案。结果表明:在相同的吹风比下,气膜有效覆盖范围随入射角α的减小而增大,展向倾角β的引入可增大气膜展向的覆盖范围;在大吹风比下,角度为α=30°、β=45°时,气膜贴壁性较好,角度为α=60°(α=45°)、β=45°时,压力面前半部气膜出现脱壁现象;在滞止线两侧采用单一角度叉排对吹式孔排布置可对前缘形成更有效的气膜保护。     

压力面气膜冷却射流复合角的数值研究

基于冷气喷射模型的验证结果,对复合角分别为0°、30°和60°三种条件下的叶片压力面前部单排孔喷射的气膜冷却特性进行了三维环形叶栅数值模拟,详细分析了在不同吹风比条件下的叶片气膜冷却效率特征。分析结果表明:Coolinh/Bleed冷气喷射模型得出的预测结果可靠。复合角使射流孔附近孔间区域冷却效率值升高。低吹风比下,复合角不能改善展向气膜冷却效率分布的均匀性;高吹风比下,复合角使展向气膜冷却效率更加均匀分布,且可减弱冷却射流脱离壁面的程度。但是,复合角不一定能增强冷却孔下游的整体气膜冷却效果。     

旋转对涡轮叶片气膜冷却影响的数值模拟

采用数值模拟的方法对旋转涡轮叶片表面的气膜冷却效率进行了研究,同时对涡轮静叶栅和动叶片在不同的旋转速度下分别进行计算,分析不同转速、吹风比和冷却气流喷射角度对气膜冷却的影响.计算结果表明,旋转使压力面气膜冷却效率降低,转速越高,气膜冷却效率越低;在吸力面冷却孔下游附近区域,叶片旋转对气膜冷却效率的影响不大,但叶片旋转使离冷却孔较远处的吸力面冷却效率升高.同时,在旋转状态下,靠近叶顶区域的叶片表面气膜冷却效率升高.     

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.     

冲角和孔排间距对端壁气膜冷却效率的影响

燃气轮机在变工况运转时透平叶栅和级的特性对燃机总体性能影响极大,而叶栅端壁气膜冷却效率是关键因素。为了提高端壁气膜冷却效率,通过优化气膜孔间距排列的方法,在叶栅端壁20%、50%、90%轴向弦长处和距前缘-10%轴向弦长端壁处布置单排带复合角度的圆柱形气膜冷却孔,运用CFD(计算流体动力学)方法对冲角(10°、0°、-10°)在不同吹风比(1、1.5、2)条件下端壁气膜冷却效率进行对比分析。结果表明:采用气膜孔非等距排列方式能有效缓解因横向压力梯度变化引起的马蹄涡在压力侧的阻隔作用,压力侧冷却效率较高;高吹风比的冷却射流会出现抛射冷却,能有效抑制冷却射流脱离壁面,壁面平均冷却效率提高;主流正冲角有利于提高端壁吸力侧气膜冷却效率,压力侧变化不大。     

入射角对气膜冷却效果影响的数值研究

采用Realizablekε-湍流模型对吹风比M=1.0时平板单圆孔不同入射角度的气膜冷却流场进行了三维定常数值模拟,得出了沿程速度、气膜冷却效率及壁面努塞尔数的分布,并对流场的流动和传热特性进行了详细分析.结果表明：随着入射角度的减小,射流在沿程方向的影响区域有所增加,最大速度点的位置逐渐下移,但并不是按入射角度减小成线性递减;在各种射流角度下,最大冷却效率均出现在气膜孔下游附近的区域,并沿程逐渐降低;当5〈x/d〈20时,冷却效率在入射角度α=10°时最大,α=70°时最小;射流下游壁面的Nu在分离点附近出现峰值,吹风比相同时,α=10°时Nu最大,α=70°时最小;在流动区域内存在反向涡旋对,小角度射流时气膜良好的贴壁性抑制了反向涡旋对的抬升和发展,其换热效果在展向的影响范围相对较小.     

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。     

双叉排孔脉动射流气膜冷却效率的实验与数值模拟研究

采用高精度红外热像仪测量了平板绝热气膜冷却效率,比较了双叉排孔和单排孔气膜冷却效率,分析了吹风比(M=0.65,1.0,1.5)和脉动频率(St=0,0.01,0.015,0.025)以及孔间作用对气膜冷却效率的影响,结合数值计算得到的瞬态流场和温度场分析了脉动射流气膜冷却下的流动传热机理。结果表明：在稳态射流工况下,单排孔的气膜冷却效率随着吹风比的增加而减小,双叉排孔的气膜冷却效率却随着吹风比的增加而增大;在脉动射流时,单排和双叉排孔的气膜冷却效率在低吹风比下低于稳态射流,在高吹风比下,脉动射流对气膜冷却效率的影响减小,且低频脉动射流气膜冷却效率略高于稳态射流。     

基于快响应压敏漆和高温磷光热图的气膜冷却非定常实验测量技术综述

涡轮内部流动及冷却射流的湍流特性决定了气膜冷却具有强烈的非定常性,精确获取高时空分辨率的气膜冷却换热特性对揭示其内在机理尤为重要。本文针对气膜冷却的非定常测试需求,聚焦于气膜冷却中涉及的壁面气膜动态覆盖以及叶片表面温度测量,从测量原理、测试系统及实际应用方面依次介绍了两种气膜冷却非定常实验测量技术。基于快响应压敏漆能够测量涡轮叶片、叶身及端壁的非定常气膜冷却,进而利用高温磷光热图评估近发动机工况下高温叶片表面的动态温度分布。     

Effect of rotation on leading edge region film cooling of a gas turbine blade with three rows of film cooling holes

Effect of rotation on detailed film cooling effectiveness distributions in the leading edge region of a gas turbine blade with three showerhead rows of radial-angle holes were measured using the Pressure Sensitive Paint (PSP) technique. Tests were conducted on the first-stage rotor blade of a three-stage axial turbine at three rotational speeds. The effect of the blowing ratio was also studied. The Reynolds number based on the axial chord length and the exit velocity was 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and exit Mach number was 0.1 and 0.3, respectively. The film cooling effectiveness distributions were presented along with the discussions on the influences of rotational speed, blowing ratio, and vortices around the leading edge region. Results showed that different rotation speeds significantly change the film cooling traces with the average film cooling effectiveness in the leading edge region increasing with blowing ratio.     

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。     

具有新型双层壁结构的透平叶片流热耦合研究

多通道壁面射流冷却结构是一种新型的燃气透平动叶内部冷却结构,具有消耗冷气少、压力损失小等优点。本文构建了简化的壁面射流冷却叶片与GE-E3冷却结构叶片模型,采用流热耦合方法对比研究了其流动与换热特性。结果表明,壁面射流冷却通道内的狭小空间抑制了横流的产生,冷气在冷却通道中形成了流向涡;前缘冷气流道中的大量冷气流经吸力侧冷却区,并从出口压力更小、面积更大的尾缘排出,使得前缘气膜孔出流的冷气流量和动量较小,冷气在叶片外表面的气膜覆盖特性更好;离心力的影响导致前缘冷气流道中叶根处的压力较低,叶根附近的气膜孔出现燃气主流入侵现象。相比于GE-E3叶片,壁面射流冷却叶片的前缘温度和温度梯度都较小,因此多通道壁面射流冷却在前缘具有更优异的冷却特性。     

Film cooling performance and heat transfer over an inclined film-cooled surface at different convergent angles with respect to highly turbulent mainstream

Experiments have been performed to study and obtain the adiabatic-wall film cooling effectiveness and the heat transfer over a film-cooled surface that is made inclined at various angles with respect to a highly turbulent flow. The film-cooled air is injected from a tangential slot. The normal temperature distributions were measured to infer the flow structure and the rate of mixing of film jet with the freestream. The freestream turbulence intensity is controlled to range from 1.0% to 26.4%, the inclination or the convergent angle of the film-cooled surface ranges from 0° to 20°, the blowing parameter from 0.5 to 2.0. It is found that the mixing of the film jet with the freestream is significantly enhanced by both the freestream turbulence intensity and the convergent angle of the film-cooled surface, which leads to the decrease in the film cooling effectiveness and the increase in the heat transfer when the inclination angle of the film-cooled surface is not large. This is attributed to the two competition mechanisms of impinging effect and the stabilization due to acceleration of the mainstream. The normal temperature distribution at several locations along the flow direction is also measured and used to infer the flow structure of the mixing of film jet with the mainstream. More detailed discussion is presented. Correlations for both the film cooling effectiveness and the heat transfer under the film-cooled surface have been very successful and are provided.     

Investigation of the Influence of Inclination Angle and Diffusion Angle on the Film Cooling Performance of Chevron Shaped Hole

The film cooling performance of chevron holes with different inclination angles and exit lateral diffusion angles has been studied experimentally and numerically. The inclination angles include 35° and 55°. The exit lateral diffusion angles include 20° and 25°. The film cooling effectiveness, heat transfer coefficient and discharge coefficient were measured on a flat plate model by transient liquid crystal measurement technique under four blowing ratios. The results show that the large inclination angle reduces the film cooling effectiveness. The influence of diffusion angle has two aspects: the large diffusion angle leads to mainstream ingestion and decreases film cooling effectiveness at M=1.0 and 1.5; however, the large diffusion angle increases the film cooling effectiveness at high blowing ratio of 2.0, because the larger hole exit area decreases the normal momentum component of the film jet. The large inclination angle decreases the heat transfer coefficient in the right downstream region at M=0.5 and 1.0. The large diffusion angle enhances the heat transfer in the right downstream of the holes in M=0.5～1.5 conditions. The chevron hole with large inclination angle generally has the highest discharge coefficient.     

Two-phase flow simulation of mist film cooling with deposition for various boundary conditions

Air film cooling is a conventional cooling technique that has been successfully used for gas turbine hot-section components, such as combustor liners, combustor transition pieces, and turbine vanes and blades. However, the increased benefit seems to approach a limit. This paper investigates the film cooling effectiveness considering mist injection. All the studies for various boundary conditions are conducted numerically, including the effects of droplet size, the flow rates of droplet injection, and the coolant air. Film cooling is also affected by the interaction between deposition and mist injection. A deposition configuration is located near the film hole with an inclination angle of 35°. Results show that the combined effect of injection and deposition is to weaken the film cooling effectiveness, especially upstream of x/d?=?19. For the coolant air at a low speed, the mist injection cannot provide better cooling protection than without the mist injection.     

The interior heat transfer characteristics of gas turbine blade due to sparse film cooling holes

The film cooling technique is one of the most useful cooling methods. At present, the midchord region of gas turbine blades in an aeroengine often adopt a sparse film cooling technique and impingement cooling technique at the same time. So the interior heat transfer characteristics on the inner side of blades due to the sparse film cooling holes have become a very complicated and interesting problem. In this paper, the heat transfer characteristics of impingement‐cooling have been investigated experimentally. Through lots of experimental data, the effect of flow parameters and geometric parameters on heat transfer characteristics has been studied. Correlation equations obtained show good agreement with experimental data. © 2005 Wiley Periodicals, Inc. Heat Trans Asian Res, 34(3): 197–207, 2005; Published online in Wiley InterScience ( www.interscience.wiley.com ). DOI 10.1002/htj.20052     

Effect of bulk flow pulsations on film cooling with compound angle holes

Experiments are conducted to investigate the effects of bulk flow pulsations on film cooling with compound angle holes. A row of five film cooling holes is considered with orientation angles of 0°, 30°, 60°, and 90° at a fixed inclination angle of 35°. Static pressure pulsations are produced by an array of six rotating shutter blades, which extend across the span of the exit of the wind tunnel test section. The pulsation frequency is fixed at 36 Hz, but changes in the time-averaged blowing ratios of 0.5, 1.0 and 2.0 produce three different coolant Strouhal numbers, 3.6, 1.8 and 0.9, respectively. Detailed film cooled boundary layer temperature distributions are measured by a cold wire and the adiabatic film cooling effectiveness by thermochromic liquid crystal (TLC). The boundary layer temperature surveys show that pulsations induce large disruptions to the boundary layer temperature distribution and the film coverage. As the orientation angle increases, the injectant concentration spreads further into the spanwise direction because of pulsations than the steady case. With pulsations the adiabatic film cooling effectiveness value decreases regardless of the orientation angle. The amount of reduction, however, depends on the orientation angle in such a way that the larger the orientation angle is, the smaller the reduction is.