耦合传热下激波对超声速气膜冷却影响

向纪鑫; 李志强; 刘鹏; 王菡

doi:10.13224/j.cnki.jasp.20210413

耦合传热下激波对超声速气膜冷却影响

doi: 10.13224/j.cnki.jasp.20210413

太原理工大学航空航天学院，太原 030000

详细信息

作者简介:
向纪鑫（1991-），男，讲师，博士，主要从事液体火箭发动机热防护研究

中图分类号: V231.3
计量
- 文章访问数: 393
- HTML浏览量: 46
- PDF量: 275
- 被引次数: 0
出版历程
- 收稿日期: 2021-08-02
- 网络出版日期: 2022-10-24

Effect of shock wave on supersonic film cooling under coupled heat transfer

College of Aeronautics and Astronautics， Taiyuan of University of Technology，Taiyuan 030000，China

摘要

摘要:
针对离散孔式超声速平板气膜冷却，在主流区引入楔角形成激波环境，以研究激波与超声速气膜之间的相互作用。通过计算楔角在0°、15°、20°和25°产生的四种激波强度下，超声速气膜与高温壁面的耦合传热。所得结果表明：适当强度的激波能够抑制气膜入射后产生的反向涡旋对，降低主流对气膜的卷吸，增大壁面平均H₂摩尔分数并降低壁面温度。对金属层温度场的分析表明，壁面冷却效果随着激波角的增加而先增加后降低，其中楔角为20°时的流场结构最有利于壁面温度保护。小楔角生成的激波在低冷流马赫数下对冷却效果的改善更明显，大楔角则在高冷流马赫数下更明显，热障涂层（TBC）不影响这种变化趋势；激波的存在削弱了TBC的影响范围。可以揭示超声速气膜在耦合传热条件下的传热机理，为超声速气膜冷却的设计提供参考，或为现有超声速气膜冷却结构的优化提供依据。
- 激波 /
- 超声速气膜冷却 /
- 耦合传热 /
- 表面传热系数 /
- 雷诺时均数值模拟
Abstract:
To study the interaction between shock wave and supersonic film, a model of supersonic plate film cooling with discrete hole was simulated. A wedge was introduced into the mainstream to create shock wave environment. Coupled heat transfer of supersonic film and high temperature wall was studied under four different shock wave intensities, in which wedge angle was 0°, 15°, 20° and 25° respectively. The simulation result showed that the appropriate intensity of shock wave can effectively eliminate the reverse vortex pairs generated with film injection into mainstream, restrain the entrainment of film, increase the average mole fraction of H₂ and also reduce the wall temperature. The analysis of metal temperature field showed that the wall cooling effectiveness first increased and then decreased with the increase of the shock wave angle. The flow field structure with wedge angle of 20° was the most effective to wall temperature protection among all cases. In addition, the cooling effect of coolant at low Mach number could be more easily improved by a smaller wedge angle, whereas the situation was opposite for bigger wedge angles. The results showed that these trends couldn’t be influenced by thermal barrier coating (TBC). The presence of shock waves weakened the scope of TBC’s influence. For revealing the heat transfer mechanism of supersonic film under coupled heat transfer conditions, this method helps to provide a reference for the design of supersonic film cooling, or to provide a basis for the optimization of existing supersonic film cooling structure.
- shock wave /
- supersonic film cooling /
- coupled heat transfer /
- surface heat transfer coefficient /
- Reynolds-averaged numerical simulation

HTML

图 1 计算模型示意图

Figure 1. Schematic of computational model

下载: 全尺寸图片幻灯片

图 2 计算域网格模型

Figure 2. Grids of computational model

下载: 全尺寸图片幻灯片

图 3 仿真结果与实验对比

Figure 3. Comparison of simulation results with experiment data

下载: 全尺寸图片幻灯片

图 4 网格无关性验证

Figure 4. Grid independent verification

下载: 全尺寸图片幻灯片

图 5 Ma_c=1.9时近壁面H₂摩尔分数分布

Figure 5. Distribution of H₂ mole fraction near the wall at Ma_c=1.9

下载: 全尺寸图片幻灯片

图 6 Ma_c=1.9时不同激波强度下的Nu分布

Figure 6. Nu distribution of different shock wave intensities at Ma_c=1.9

下载: 全尺寸图片幻灯片

图 7 Ma_c=1.9时壁面平均H₂摩尔分数沿流向变化

Figure 7. Transverse distribution of H₂ averaged mole fraction on the wall at Ma_c=1.9

下载: 全尺寸图片幻灯片

图 8 Ma_c=1.9壁面平均表面传热系数沿流向变化

Figure 8. Transverse distribution of averaged heat transfer coefficient on the wall at Ma_c=1.9

下载: 全尺寸图片幻灯片

图 9 固体域对称面温度分布（Ma_c=1.9）

Figure 9. Temperature distribution of symmetrical surface in solid region (Ma_c=1.9)

下载: 全尺寸图片幻灯片

图 10 固体域对称面温度分布（Ma_c=1.6）

Figure 10. Temperature distribution of symmetrical surface in solid region （Ma_c=1.6）

下载: 全尺寸图片幻灯片

图 11 固体域对称面温度分布（Ma_c=1.3）

Figure 11. Temperature distribution of symmetrical surface in solid region （Ma_c=1.3）

下载: 全尺寸图片幻灯片

图 12 不同冷流马赫数下的金属层体平均温度

Figure 12. Volume-averaged temperature of metal layer at different coolant Mach numbers

下载: 全尺寸图片幻灯片

表 1 固体域材料参数

Table 1. Material parameters of solid region

材料厚度/mm 导热系数/（W/（m·K））

8YSZ（陶瓷基体） 0.3 2.16
NiCoCrAlY（黏结层） 0.2 4.00
Cooper（金属平板） 5.0 387.60

下载: 导出CSV

表 2 不同工况下的入口参数

Table 2. Inlet parameters of different conditions

工质 Ma p^*/kPa T^*/K M

主流空气 3.2 4000 1600
冷却剂工况1 H₂ 1.9 750 400 0.3
工况2 H₂ 1.6 500 400 0.25
工况3 H₂ 1.3 350 400 0.17

下载: 导出CSV

参考文献(18)

[1]	葛绍岩,刘登瀛,徐靖中,等.气膜冷却[M].北京: 科学出版社,1985.
[2]	商圣飞,向树红,姜利祥,等. 不同孔型对高超声速逆喷流气膜冷却影响[J]. 航空动力学报,2020,35(8): 1612-1621. SHANG Shengfei,XIANG Shuhong,JIANG Lixiang,et al. Numerical simulation of supersonic gaseous film cooling[J]. Journal of Aerospace Power,2020,35(8): 1612-1621. (in Chinese)
[3]	王建,孙冰,魏玉坤,等. 超声速气膜冷却数值模拟[J]. 航空动力学报,2008,23(5): 865-870. WANG Jian,SUN Bing,WEI Yukun,et al. Numerical simulation of supersonic gaseous film cooling[J]. Journal of Aerospace Power,2008,23(5): 865-870. (in Chinese)
[4]	GROSDEMANGE H. Advanced nozzle extension demonstration for Vulcain Mark Ⅱ application[R]. AIAA-1999-2537,1999.
[5]	FUKUSHIMA Y,NAKATUZI H,KISHIMOTO K,et al. Development status of LE-7A engine[R].AIAA-1997-2817,1997.
[6]	丁兆波,孙纪国,路晓红,等. 国外典型大推力氢氧发动机推力室技术方案综述[J]. 导弹与航天运载技术,2012(4): 27-30. doi: 10.3969/j.issn.1004-7182.2012.04.006 DING Zhaobo,SUN Jiguo,LU Xiaohong,et al. Review on technical schemes of foreign large LOx/LH₂ thrust chamber[J]. Missiles and Space Vehicles,2012(4): 27-30. (in Chinese) doi: 10.3969/j.issn.1004-7182.2012.04.006
[7]	彭威,姜培学. 直通道和弯曲通道中超声速气膜冷却研究[J]. 航空动力学报,2008,23(3): 406-409. PENG Wei,JIANG Peixue. Investigation of supersonic film cooling in flat and convex channels[J]. Journal of Aerospace Power,2008,23(3): 406-409. (in Chinese)
[8]	KONOPKA M,MEINKE M,WOLFGANG S,et al. Large-eddy simulation of shock/cooling-film interaction[J]. Angewandte Chemie,2013,45(18): 2962-2965.
[9]	KONOPKA M,MEINKE M,SCHRODER W,et al. Large-eddy simulation of high Mach number film cooling with shock-wave interaction[J]. Progress in Flight Physics,2013,5(6): 309-326.
[10]	PENG Wei,JIANG Peixue. Influence of shock wave on supersonic flim cooling[J]. Journal of Spacecraft and Rockets,2009,46(1): 67-73. doi: 10.2514/1.38458
[11]	SUN Xiaokai,NI Hang,PENG Wei,et al. Influence of shock wave impinging region on supersonic film cooling[J]. Chinese Journal of Aeronautics,2021,34(5): 452-465. doi: 10.1016/j.cja.2020.12.012
[12]	倪航,王明军,彭威,等. 冷却流分段注入对超声速气膜冷却的影响[J]. 工程热物理学报,2021,42(8): 2090-2096. NI Hang,WANG Mingjun,PENG Wei,et al. Effect of segmented cooling-stream injection on supersonic film cooling[J]. Journal of Engineering Thermophysics,2021,42(8): 2090-2096. (in Chinese)
[13]	孙冰,王太平,张佳. 离散孔结构超声速气膜冷却数值模拟[J]. 航空动力学报,2017,32(5): 452-465. SUN Bing,WANG Taiping,ZHANG Jia. Numerical simulation of discrete holes supersonic gaseous film cooling[J]. Journal of Aerospace Power,2017,32(5): 452-465. (in Chinese)
[14]	ZHANG Jia,SUN Bing. Experiments on film cooling with sonic injection into a supersonic flow[J]. Journal of Aerospace Power,2015,30(5): 1084-1091.
[15]	曹学强. 热障涂层材料[M].北京:科学出版社,2007.
[16]	HYDE C R,SMITH B R,SCHETZ J A,et al. Turbulence measurements for heated gas slot injection in supersonic flow[J]. AIAA Journal,1990,28(9): 1605-1614. doi: 10.2514/3.25259
[17]	PENG Wei,JIANG Peixue. Effect of shock wave on supersonic film cooling with a slotted wall[J]. Applied Thermal Engineering,2014,62(1): 187-196. doi: 10.1016/j.applthermaleng.2013.09.013
[18]	ZUO J,ZHANG S,QIN J,et al. Effects of shock waves interaction on hydrocarbon fueled supersonic film cooling with combustion[J]. Aerospace Science and Technology,2021,113(6): 106693.1-106693.14.