气膜孔形状对涡轮叶片气膜冷却影响的研究进展

气膜冷却是航空发动机叶片上采用的冷却方式之一,气膜孔结构对冷却效率影响非常显著。通过对不同形状孔射流气膜冷却回顾,指出了圆柱孔射流冷却的有害涡流动结构。论述了几何结构和气动参数对气膜冷却特性的影响,提出了一种高效气膜冷却孔结构——双出口气膜孔。利用商业软件对双出口射流的冷却效率进行了数值模拟。结果表明,双出口孔射流时,形成的涡结构有利于冷气贴附在壁面。最后给出了圆柱孔和双出口孔射流冷却效率对比结果,无论在平板上还是在叶片前缘,双出口孔射流冷却效率都明显高于圆柱孔射流冷却效率。     

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.     

涡轮叶片表面全气膜冷却传热实验研究

对全气膜覆盖的涡轮导向叶片的表面进行了详细的传热实验研究,重点研究了不同质流比和不同雷诺数对当地气膜冷却效率和换热系数的影响.实验结果表明:质流比(冷气质量流量和主流质量流量的比值)的变化会显著地影响叶片表面温度场的分布,从而影响叶片表面的换热和冷却效率;在同一质流比下,雷诺数对气膜冷却效率的影响则相对较小.全气膜冷却能够有效地降低叶片的热负荷,该结果可作为在工程实际中的参考.     

气膜孔几何结构对涡轮叶片气膜冷却影响的研究进展

综述了近年来涡轮叶片气膜孔几何结构对气膜冷却特性影响的研究成果,介绍了影响气膜冷却效果的因素.总结了气膜孔结构对叶片前缘、叶片端壁以及对平板气膜冷却影响的研究现状.阐述了气膜孔结构对气膜冷却传热特性的影响.最后指出进一步优化气膜孔结构,综合考虑气膜孔尺寸、长度、间隔、形状以及相对透平叶片取向对气膜冷却的影响和新型气膜孔的研究.将是今后工作的重点.     

气膜孔堵塞对叶片吸力面气膜冷却的影响

为阐明热障涂层工艺造成的气膜孔堵塞对气膜冷却的影响机理,采用数值模拟方法研究了叶片吸力面在气膜孔堵塞比为0.2、0.5和0.8,吹风比为0.5、1.0、1.5和2.0时的气膜冷却效率变化.结果表明:堵塞比越大,气膜冷却效率下降幅度越明显,孔下游气膜覆盖面积越小;相比无堵塞情况,堵塞比为0.8时展向平均气膜冷却效率退化为...     

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.     

涡轮叶片前缘气膜冷却的数值研究

对前缘带有2排冷却孔的高压涡轮叶栅进行了气膜冷却数值计算,在吹风比分别为0、0.7、1.1、1.5的情况下得到了叶片型面的静压分布,并将计算结果与实验数据进行了对比研究.此外,还详细分析了冷却孔出口附近区域的流场特性.结果表明,吹风比对叶片型面的压力分布影响不大,只是在冷却孔附近有较为明显的变化.同时,由于冷气射流的注入,在冷却孔后出现了一对旋转方向相反的肾形涡,此时增大吹风比不会对旋涡产生明显的影响,但对旋涡下游流场的影响比较明显.     

带前缘对吹孔涡轮导向叶片气膜冷却特性实验

为获得涡轮导向叶片气膜冷却特性,在叶栅风洞中运用红外热成像技术进行了带前缘对吹孔涡轮导向叶片的气膜冷却特性实验。叶片前缘布置5排复合角气膜孔形成对吹孔结构,其特点是叶片高度方向的上下两部分气膜孔径向角都偏向中截面。吸力面和压力面分别布置5排和16排圆形孔。测试的叶栅入口雷诺数为1.2×105,2.4×105和3.6×105,吹风比为1.0,1.5和2.0。实验结果表明：从前缘对吹孔出流的冷气向吸力面和压力面中截面聚集,导致中截面区域气膜覆盖效果增强;吹风比为1.0时,前缘和压力面中截面换热系数低;随吹风比增加中截面换热增强,压力面和吸力面高换热区域沿流向变长;雷诺数为1.2×105时,压力面气膜覆盖呈发散状;雷诺数为2.4×105和3.6×105时,压力面气膜覆盖宽度沿流向先变窄后变宽。     

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。     

叶片前缘表面气膜冷却的实验研究

燃气轮机叶片前缘处有很多高换热区,其中叶片前缘表面处是最先与高温燃气接触的位置,存在承受较高的热负荷的问题,目前较为常见的解决办法是气膜加冲击冷却。通过实验对比不同吹风比、不同曲率位置气膜冷却排孔的温度分布,得出以下结论:冷却效果以出气孔为起点向下游不断减弱,并且在不同的排孔处,下游减弱的趋势不同;当吹风比M1时,吸力面气膜冷却有效度随吹风比的增加而逐渐增加;当吹风比M1时,气膜冷却有效度随吹风比的增加而减小;下游曲率的增加可以有效地增加气膜冷却效果。     

Experimental Investigation on the Effect of Mainstream Turbulence on Full Coverage Film Cooling Effectiveness for a Turbine Guide Vane

Experiments have been performed to investigate the effect of mainstream turbulence on the three-dimensional distribution of the full coverage film cooling effectiveness for two enlarged actual twisted vanes with cylindrical or shaped holes. The film cooling effectiveness was measured by transient liquid crystal technique at mainstream turbulence intensities of 2%, 9% and 15%. The mass flow rate ratios range from 5.5% to 12.5%. There are 3, 8 and 7 rows of film holes on the suction side, leading edge and pressure side, respectively. Results show that for the cylindrical hole vane the high mainstream turbulence intensity decreases the film cooling effectiveness in the top region and down region of pressure side in the low mass flow rate ratio of 5.5%, while the effect is opposite in the high mass flow rate ratio of 12.5%. The film cooling effectiveness in the middle region of pressure side decreases obviously with the increase of the turbulence at the low mass flow rate ratio of 5.5%, while the influence of increasing turbulence weakens gradually with the increase of mass flow rate ratio. Moreover, the high mainstream turbulence improves the film cooling effectiveness in the further downstream of the holes on suction side at the high mass flow rate ratio of 12.5%. For the shaped hole vane, the increase of mainstream turbulence decreases the film cooling effectiveness at all mass flow rate ratios. This study reveals the influence rule of the mainstream turbulence on the film cooling effectiveness in the different regions of the three-dimensional vane surface. The results would guide the designs of engineering heat transfer with application in gas turbine blade/vane cooling.     

Film Cooling Effectiveness Superposition Calculation of Double-Row Injection Holes on Turbine Vane

Film cooling effectiveness superposition of double-row injection holes on the turbine vane was studied by infrared temperature measurement experiment. The Sellers superposition method and a modified Sellers method were adopted for dustpan-shaped hole and cylindrical hole. Numerical simulations were implemented to analyze the film superposition mechanism. It is found that the Sellers method is more accurate on the suction side than the pressure side. Injection film of the two types of holes exhib...     

Effects of Trenched Film Hole Configurations on the Endwall Film Cooling and Suction Side Phantom Cooling

To maximize the turbine thermal efficiency, modern gas turbine's inlet temperature is significantly augmented within the past few decades. To prolong the lifespan of gas turbines, many efficient cooling techniques have been proposed and applied in the endwall cooling schemes. However, conventional discrete film hole does not take effect at the leading edge nearby region. In this research, how the trenched film hole configurations affects the endwall cooling and phantom cooling characteristics were deeply studied by using a verified approach. Steady 3D Reynolds-averaged Navier-Stokes(RANS) governing equations together with the shear stress transport(SST) k-w turbulence model have been solved. Firstly, results indicate that trenched film holes greatly influence the cooling effectiveness at leading edge nearby region compared to normal case. Nevertheless, suction side phantom cooling is hardly influenced by the trenched film holes. Secondly, the case with a smaller trench width obtains higher endwall cooling effectiveness, particularly at upstream region. More importantly, the cases with W=3D achieve large cooling effectiveness at leading edge nearby region with little influence by trench depth. Additionally, majority of trenched film holes coolant flow is driven towards middle passage. Therefore, the suction side phantom cooling is unaffected by the trenched film holes.     

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。     

低压涡轮导向器气动性能的实验研究

为了考查低压涡轮导向器的气动性能,在低速环形叶栅风洞上对导向器的实验叶栅进行了吹风实验。测量了由栅前至栅后共6个横截面的气动参数分布以及叶片和上、下端壁表面的静压场。实验结果表明,主流区的损失并不太大,主要的损失发生在上、下端壁附近,尤其是上端壁附近的损失非常大。引起叶栅损失的主要根源在于叶栅子午面较大的扩张,进一步降低叶栅的损失应从改进叶栅通道的子午面型线出发。     

Experimental Investigation of Louver Cooling Scheme on Gas Turbine Stator

An experimental investigation has been conducted to investigate the film cooling performance of the louver scheme over the surface of a gas turbine stator using a transient thermochromic liquid crystal technique. A two-dimensional airfoil cascade has been employed during the study. The exit Reynolds number based on the true chord length is 1.5E5 and the exit Mach number is 0.23. Two rows of an axially oriented louver scheme are distributed on both suction and pressure sides in a staggered arrangement. The effect of hole location on the cooling performance is investigated for each row individually; then row interaction is investigated at four different blowing ratios ranging from 1 to 2 and a 0.9 density ratio. The detailed local performance distribution and the lateral-averaged normalized performance are presented over both sides of the vane in terms of heat transfer coefficient and cooling effectiveness. The louver scheme provides a better cooling performance compared with the similar cylindrical scheme of the same base diameter at the same cooling amount. The blowing ratio does not influence significantly the performance for the louver scheme due to the considerable decrease in the jet momentum that impedes the jet lift-off at exit. The location of the scheme exit has a high impact on the cooling performance as it affects the development of the boundary layer. The double injection on the pressure side provides a superior effectiveness due to the blockage of the mainstream by the coolant injected from the first row. The louver scheme provides higher net heat flux reduction, which suits the cooling capacity needed for the next generation of gas turbines.     

Experimental Study on Impingement and Film Overall Cooling Characteristics of Turbine Shroud

Based on the variable characteristics of the actual operating conditions of the turbine shroud and the purpose of improving the cooling effect of the turbine shroud,this paper builds a test system of the impingement-film cooling shroud with two gas inlet angles(90°,167°).The effects of film cooling hole arrangement,gas inlet angle,blowing ratio(0.7,1.0,1.5,2.0,2.5,3.0)and temperature ratio(1.2,1.3,1.4,1.5,1.6)on the cooling characteristics of the impingement-film cooling shroud were experimentally studied by infrared temperature measurement technology,especially the effects of gas inlet angle and temperature ratio.The results showed that the film covering effect of the film cooling hole vertical or the same direction of the high-temperature gas incoming flow is better than the film covering effect of the reverse direction with the incoming flow,and the optimal arrangement of film cooling holes can improve the cooling effectiveness of the shroud.Compared with 90°intake gas,the film coverage area on the shroud surface of the 167°intake gas is expanded,and the surface average overall cooling effectiveness is increased by 1.03%to 12.6%.The overall cooling effectiveness of turbine shroud increases with the increase of blowing ratio,which increases the flow rate and pressure of cooling gas,and the corresponding increase rate is between 1.04%and 9.96%.However,the increase in the temperature ratio increases the mainstream heating capacity,resulting in a decrease in the cooling effectiveness of the shroud,with a maximum reduction rate of 11.04%.     

Investigation of Interacting Mechanism between Film Cooling and Internal Cooling Structures of Turbine Blade

This paper presents three-dimensional numerical simulations with the established realizable k-ε model to clarify the underlying and interacting mechanisms between the film cooling and the internal cooling. On the one hand, the effects of three different internal cooling channels, i.e., smooth channel, continuous ribbed channel, and truncated ribbed channel, on the film cooling effectiveness and the discharge coefficients are investigated. On the other hand, the influences of three different film...     

透平静叶表面多排孔气膜冷却特性试验研究

为研究燃气轮机透平静叶表面的气膜冷却特性,搭建了平面叶栅气膜冷却试验台,采用红外成像技术测量了带有多排扇形气膜孔的透平静叶表面气膜冷却有效度和传热系数,分析了不同吹风比下静叶压力面和吸力面不同孔排下游气膜冷却有效度和传热系数的分布规律.结果 表明:随着吹风比的增大,静叶压力面各排孔后气膜冷却有效度增大,静叶吸力面孔排1下游气膜冷却有效度减小,吸力面孔排2和吸力面孔排3下游气膜冷却有效度先增大后减小,吸力面孔排4下游气膜冷却有效度增大;大吹风比时,气膜出流使得静叶表面大部分区域展向平均传热系数相比无气膜时小幅增大.     

涡轮动叶压力面顶部喷冷对气动性能的影响

本文采用中心差分格式和多区网格技术,对燃气轮机第一级动叶喷射冷气流场进行了全三维N-S方程数值求解。分析了在端壁不同角度冷气喷射条件下,端壁区域的冷却效率、温度场的分布规律,以及不同角度冷气喷射对叶片端壁区域流场的影响。结果表明,在动叶压力面顶部喷射的冷气很好地覆盖了上端壁表面,起到了很好的冷却效果,喷射方向与端壁夹角较大时,端区的能量损失系数较小,冷却效率较高,与端壁夹角较小时,对端壁冷气覆盖效果较好。