跨声速叶栅通道中叶片压力面簸箕孔型气膜冷却特性

李杰; 骆剑霞; 朱惠人

doi:10.13224/j.cnki.jasp.2020.08.001

跨声速叶栅通道中叶片压力面簸箕孔型气膜冷却特性

doi: 10.13224/j.cnki.jasp.2020.08.001

计量
- 文章访问数: 687
- HTML浏览量: 137
- PDF量: 611
- 被引次数: 0
出版历程
- 收稿日期: 2020-01-13
- 刊出日期: 2020-08-28

Film cooling performance of fan-shaped film hole on blade pressure side in linear transonic cascade

1.
Research and Design Center,Commercial Aircraft Engine Company Limited,Aero Engine Corporation of China,Shanghai 201108,China
2.
School of Power and Energy,Northwestern Polytechnical University,Xi’an 710072,China

摘要

摘要: 在跨声速叶栅通道内,试验研究了叶片压力面气膜冷却特性,详细地对比分析了在不同主流进口雷诺数(Re=17×105,37×105,57×105)、出口马赫数(Ma=081,091,101)及多个气膜吹风比(M=05～30)条件下的压力面簸箕孔型气膜冷却效率。试验结果表明:主流出口马赫数变化对气膜孔下游冷却效率的分布与具体数值均无影响;而主流进口雷诺数的影响较大。增大主流进口雷诺数使得气膜分离再贴附对应的吹风比相应增大,Re=17×105时在吹风比M=10时出现气膜分离与再贴附现象,而Re=37×105和Re=57×105对应的临界吹风比则分别为20和25;主流进口雷诺数越大,小吹风比下近孔区域的冷却效率越高,而在孔下游区域则相反;大吹风比下,则主流进口雷诺数越大冷却效率越小。
- 跨声速叶栅 /
- 压力面 /
- 气膜孔 /
- 吹风比 /
- 冷却效率
Abstract: Film cooling performance of blade pressure side was tested in a linear transonic cascade. The film cooling effectiveness of the fan-shaped film hole on blade pressure side was measured at different mainstream inlet Reynolds numbers of 17×105, 37×105，57×105, different mainstream exit Mach numbers of 081, 091，101 and different blowing ratios of 05-30. Results showed that the mainstream exit Mach number had no distinct effect on the film cooling effectiveness at downstream film hole; however, the mainstream inlet Reynolds number had a significant effect on the film cooling effectiveness. The blowing ratio corresponding to distinguish the film’s lifting off increased with the added mainstream inlet Reynolds number, the film began to lift off at blowing ratio of 10 at mainstream inlet Reynolds number of 17×105 while these corresponding to mainstream inlet Reynolds number of 37×105 and 57×105 were blowing ratios of 20 and 25. At small blowing ratios, the case with higher mainstream inlet Reynolds number had a lager film cooling effectiveness at the region adjacent to the hole and a lower film cooling effectiveness at downstream area; however, at high blowing ratios, higher mainstream inlet Reynolds number resulted in lower film cooling effectiveness.
- transonic cascade /
- pressure side /
- film cooling hole /
- blowing ratio /
- ,film cooling effectiveness

HTML

参考文献(18)

[1]	FRIC F T,ROSHKO A.Vortical structure in the wake of a transverse jet［J］.Journal of Fluid Mechanics,1994,279(10):1-47.
[2]	MCGOVERN K T,LEYLEK J H.A detailed analysis of film cooling physics:Part 2 compound-angle injection with cylindrical holes［J］.Journal of Turbomachinery,1997,119(1):113-121.
[3]	SEN B,SCHMIDT D L,BOGARD D G.Film cooling with compound angle holes:heat transfer［J］.Journal of Turbomachinery,1996,118(4):807-813.
[4]	GOLDSTEIN R J,ECKERT E R G,BURGGRAF F.Effects of hole geometry and density on three-dinemsional film cooling［J］.International Journal of Heat and Mass Transfer,1974,17(5):595-607.
[5]	GANZERT W,HILDEBRANDT T,FOTTNER L.Systematic experimental and numerical investigations on the aerothermodynamics of a film cooled turbine cascade with variation of the cooling hole shape:Part 1 experimental approach［R］.ASME Paper 2000-GT-295,2000.
[6]	ADAMI P,MARTELLI F,MONTOMOLI F,et al.Numerical investigation of internal crossflow film cooling［R］.ASME Paper GT-2002-30171,2002.
[7]	WANG Zhan,LIU Jianjun,ZHANG Chao.Impacts of geometric parameters pf double-jet film cooling on anti-kidney vortex structure and cooling effectiveness［R］.ASME Paper GT2013-94038,2013.
[8]	朱惠人,许都纯,刘松龄,气膜孔形状对排孔下游冷却效率的影响［J］.航空学报,2002,23（1）:75-78. ZHU Huiren,XU Duchun,LIU Songling.Effects of hole shape on film cooling effectiveness［J］.Journal of Aerospace Power,2002,23（1）:75-78.(in Chinese)
[9]	骆剑霞,朱惠人,张宗卫.涡轮导向叶片换热实验与计算［J］.航空动力学报,2014,29(3):526-531. LUO Jianxia,ZHU Huiren,ZHANG Zongwei.Heat transfer experiment and computation of a gas turbine vane［J］.Journal of Aerospace Power,2014,29(3):526-531.(in Chinese)
[10]	李广超.压力梯度和气膜孔布局对气膜冷却特性影响的研究［D］.西安:西北工业大学,2008. LI guangchao.The effects of the mainstream pressure gradient and various hole geometries on film cooling effectiveness［D］.Xi’an:Northwestern Polytechnical University,2008.(in Chinese)
[11]	SCHWARZ S G,GOLDSTEIN R J.The two-dimensional behavior of film cooling jets on concave surfaces［R］.ASME Paper 88-GT-161,1988.
[12]	SCHWARZ S G,GOLDSTEIN R J,ECKER E R G.The influence of curvature on film cooling performance［R］.ASME Paper 90-GT-010,1990.
[13]	李国庆,朱俊强.涡轮叶片压力面旋转气膜冷却数值模拟［J］.工程热物理学报,2011,32(5):835-838. LI Guoqing,ZHU Junqiang.Numerical simulation of film cooling on pressure side of rotating turbine blade［J］.Journal of Engineering Thermophysics,2011,32(5):835-838.(in Chinese)
[14]	苗强,张发生,戴韧,压力面气膜冷却射流复合角的数值研究［J］.热力透平,2010,39(3):190-194，220. MIAO Qiang,ZHANG Fasheng,DAI Ren.Numerical research on pressure side film cooling at different compound angles［J］.Thermal Turbine,2010,39(3):190-194,220.(in Chinese)
[15]	韩昌,任静,蒋洪德,透平静叶前缘和压力面气膜冷却实验研究［J］.清华大学学报(自然科学版),2014,54(6):769-774. HAN Chang,REN Jing,JIANG Hongde.Experimental investigation of film cooling of the leading edge and pressure side of a turbine vane［J］.Journal of Tsinghua University(Science and Technology),2014,54(6):769-774.(in Chinese)
[16]	张扬,苟金澜,袁新.压力面复合角孔型对气膜冷却的影响［J］.工程热物理学报,2013,34(5):826-829. ZHANG Yang,GOU Jinlan,YUAN Xin.Effects of compound angle on pressure side gill region film cooling［J］.Journal of Engineering Thermophysics,2013,34(5):826-829.(in Chinese)
[17]	LIU Cong,ZHU Huiren,FU Zongyi,et al.The effects of inlet Reynolds number,exit math number of a turbine blade in a linear transonic cascade［R］.ASME Paper GT2015-42888,2015.
[18]	MOFFAT R J.Describing the uncertainties in experimental results［J］.Experimental Thermal and Fluid Sciences,1988,1(1):3-17.