横流对叶片前缘冲击-旋流-气膜冷却通道内部传热影响

韩枫; 张舒豪; 陈娇娜; 王凌洋; 陈宏华; 杨凌; 毛军逵

doi:10.13224/j.cnki.jasp.20240415

横流对叶片前缘冲击-旋流-气膜冷却通道内部传热影响

doi: 10.13224/j.cnki.jasp.20240415

韩枫^{1, 2,},
张舒豪^{1, 2},
陈娇娜^{1, 2},
王凌洋³,
陈宏华^{1, 2},
杨凌¹,
毛军逵^1,*, ,

1.
南京航空航天大学能源与动力学院，南京 210016
2.
南京航空航天大学综合能源研究学院，南京 210016
3.
中国科学院微小卫星创新研究院，上海 200120

基金项目: 国家自然科学基金（52476077）；国家科技重大专项（Y2022-Ⅲ-0003-0012）；航空发动机及燃气轮机基础科学中心项目（P2022-A-Ⅱ-007-001）；中央高校基本科研业务费（NS2023010）

详细信息

作者简介:
韩枫（1989-），男，副教授，博士，从事航空发动机热端部件流动与换热研究。E-mail：hanfeng@nuaa.edu.cn

通讯作者:
毛军逵（1976-），男，教授，博士，从事航空发动机热管理和高效热防护研究。E-mail：mjkpe@nuaa.edu.cn

中图分类号: V231.1
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出版历程
- 收稿日期: 2024-06-25
- 网络出版日期: 2024-12-23

Effect of crossflow on internal heat transfer of impingement-swirl-film cooling for leading edge of turbine blade

HAN Feng^{1, 2
,},
ZHANG Shuhao^{1, 2},
CHEN Jiaona^{1, 2},
WANG Lingyang³,
CHEN Honghua^{1, 2},
YANG Ling¹,
MAO Junkui^{1,*
, ,}

1.
College of Energy and Power Engineering，Nanjing University of Aeronautics and Astronautics，Nanjing 210016，China
2.
Integrated Energy Institute，Nanjing University of Aeronautics and Astronautics，Nanjing 210016，China
3.
Innovation Academy for Microsatellites，Chinese Academy of Sciences，Shanghai 200120，China

摘要

摘要:
针对真实涡轮叶片前缘非对称的冲击-旋流-气膜冷却通道，在定冲击进口雷诺数和定总流量两种情况下，通过改变横流比，数值研究了主动横流入流对冷却通道换热效果的影响规律；并拟合获得了两组工况下的努塞尔数无量纲关系式。结果表明：在冲击进口雷诺数不变的情况下，增加横流比会增强冲击靶面的整体换热能力；当雷诺数为20000时，横流比为2获得的努塞尔数比无横流工况时增加了25.8%。在总流量不变的情况下，增加横流比会减弱冲击靶面的整体换热能力；当总流量为0.0236 kg/s时，横流比为2获得的努塞尔数比无横流工况时降低了43.3%。非对称的冲击-旋流-气膜前缘冷却通道内的冷却效果主要由冲击射流决定，相比较而言横流对流冷却影响较小。通过拟合获得的两种情况的努塞尔数关系式，对涡轮叶片前缘非对称结构的研究和优化具有指导意义。
- 前缘 /
- 复合冷却 /
- 旋流冷却 /
- 横流 /
- 无量纲关系式
Abstract:
The asymmetric impingement-swirl-film cooling channel of a realistic turbine blade leading edge was studied. Under two conditions—constant impingement inlet Reynolds number and constant total flow rate, the effect of active crossflow on the heat transfer performance of the cooling channel was numerically investigated by varying the crossflow ratio. Dimensionless Nusselt number correlations at both operating conditions were derived through curve fitting. The results indicated that, under a constant Reynolds number of the impingement inlet, increasing the crossflow ratio enhanced the overall heat transfer performance of the impingement target surface. Specifically, when the Reynolds number was 20000, a crossflow ratio of 2 led to a 25.8% increase in the Nusselt number compared with the no-crossflow case. Conversely, under a constant total flow rate, increasing the crossflow ratio reduced the overall heat transfer performance of the impingement target surface. When the total flow rate was 0.0236 kg/s, a crossflow ratio of 2 resulted in a 43.3% reduction in the Nusselt number compared with the no-crossflow case. The findings demonstrated that the cooling performance in the asymmetric impingement-swirl-film cooling channel was primarily governed by the impinging jet, while the effect of crossflow convection was relatively minor. The fitted Nusselt number correlations for both scenarios can provide a valuable guidance for the design and optimization of asymmetric leading-edge cooling structures in turbine blades.
- leading edge /
- composite cooling /
- swirl cooling /
- crossflow /
- dimensionless correlation

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图 1 典型叶片冷却结构

Figure 1. Typical turbine blade cooling structure

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图 2 研究模型几何结构

Figure 2. Model geometry

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图 3 边界条件

Figure 3. Boundary conditions

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图 4 网格示意图

Figure 4. Schematic diagram of mesh grids

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图 5 网格数量无关性验证

Figure 5. Grid-independence validation

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图 6 验证模型示意图

Figure 6. Schematic diagram of validated model

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图 7 验证模型$\overline {Nu} $比较

Figure 7. $\overline {Nu} $ comparison of validated model

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图 8 不同横流比下局部Nu分布云图（Re=20000）

Figure 8. Local Nu distributions with different crossflow ratio （Re=20000）

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图 9 不同横流比下冲击靶面特征线Nu分布（X/d=1.5, Re=20000）

Figure 9. Nu distribution of characteristic line of impact target with different crossflow ratio （X/d=1.5, Re=20000）

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图 10 不同横流比下冲击靶面$\overline {Nu} $（Re=20000）

Figure 10. $\overline {Nu} $ of impact target with different crossflow ratio （Re=20000）

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图 11 速度矢量图（X/d=1.5, Re=20000, m₂/m₁=0）

Figure 11. Velocity vector diagram （X/d=1.5, Re=20000, m₂/m₁=0）

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图 12 不同横流比下速度分布图（X/d=1.5, Re=20000）

Figure 12. Velocity distributions with different crossflow ratio （X/d=1.5, Re=20000）

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图 13 不同横流比下局部Nu分布云图（m=0.0236 kg/s）

Figure 13. Local Nu distributions with different crossflow ratio （m=0.0236 kg/s）

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图 14 不同横流比下冲击靶面特征线Nu分布（X/d=1.5, m=0.0236 kg/s）

Figure 14. Nu distribution of characteristic line of impact target with different crossflow ratio （X/d=1.5, m=0.0236 kg/s）

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图 15 不同横流比下冲击靶面$\overline {Nu} $ （m=0.0236 kg/s）

Figure 15. $\overline {Nu} $ of impact target with different crossflow ratio（m=0.0236 kg/s）

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图 16 不同横流比下第3个冲击孔中心截面速度流线图（m=0.0236 kg/s）

Figure 16. Velocity streamline diagram of the central section of the third impact holes with different crossflow ratio（m=0.0236 kg/s）

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图 17 射流中心线速度分布曲线

Figure 17. Velocity distribution curve of jet center line

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图 18 无量纲关系式和数值模拟结果对比（以冲击进口Re和横流比为变量）

Figure 18. Comparison of dimensionless relation and numerical simulation results （Re and crossflow ratio are used as variables）

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图 19 无量纲关系式和数值模拟结果对比（以总流量和横流比为变量）

Figure 19. Comparison of dimensionless relation and numerical simulation results （total flow and crossflow ratio are used as variables）

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表 1 结构几何参数

Table 1. Geometrical parameters of structure mm

几何参数	数值
冲击孔直径 d	10
气膜孔直径 d′	3
冲击隔板和吸力面延长线交点与吸力面下端点距离 l₁	73
压力面和吸力面延长线交点与压力面上端点距离 l₂	54
冲击隔板-吸力面交接面曲率半径和冲击隔板-吸力面交接面曲率半径 r₁	5
叶片前缘曲率半径 r₂	12.5
冲击孔偏移距离 e	15
气膜孔间距 s	15
冲击隔板与前缘3排气膜孔进口中心点垂直距离 i₁, i₂, i₃	20, 25, 30

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表 2 不同雷诺数下具体工况

Table 2. Specific conditions under different Reynolds numbers

Re	m₂/(kg/s)	m₂/m₁
10⁴	0.00787	1，2，3，4
2×10⁴	0.0157	0.5，1，1.5，2
3×10⁴	0.0236	0.33，0.67，1，1.33
4×10⁴	0.0315	0.25，0.5，0.75，1

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表 3 不同流量下具体工况

Table 3. Specific working conditions under different flow rates

m/（kg/s）	0.0157	0.0236	0.03148
m₂/m₁	0.05/0.1/0.33/0.5/0.75/1/1.5/2/2.5/3

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