火箭推进

超燃冲压发动机的支板在高飞行马赫数下工作时,面临着非常严峻的热环境。对飞行马赫数为8时的支板热环境进行了研究,得到了前缘钝化半径与支板表面温度的关系; 随后,对不同喷注总压,喷孔尺寸和喷注总温下的逆向喷注进行了数值模拟。数值模拟的结果表明,提高喷注总压和增大喷孔尺寸都有助于降低支板表面温度,在喷注总温上升到1 000 K时,逆向喷注仍然具有较好的热防护性能。

The strut of scramjet engine faces a severe thermal environment when operating at high Mach numbers.In this paper, the thermal environment of the strut was studied when the flight Mach number was 8 and the relationship between the passivation radius and the surface temperature of the strut was obtained.Subsequently, numerical simulations were carried out for the reverse injection under different injection total pressures, orifice sizes, and total injection temperatures.The results of numerical simulation show that increasing the total injection pressure and increasing the size of the orifice help to reduce the surface temperature of the strut.When the total injection temperature rises to 1 000 K, the reverse gas injection is still applicable.

引言
1 数值模拟条件
2 支板热环境分析
3 逆向气体喷注热防护
4 结论

图1 燃烧室尺寸和支板位置<br/>Fig.1 Combustion chamber size and strut position

图1 燃烧室尺寸和支板位置
Fig.1 Combustion chamber size and strut position

图2 支板上的测量点分布<br/>Fig.2 Distribution of measurement points on the strut

图2 支板上的测量点分布
Fig.2 Distribution of measurement points on the strut

图3 不同钝化半径下激波锋面与支板前缘的距离<br/>Fig.3 Distance between the shock front and the strut leading edge under different passivation radiuses

图3 不同钝化半径下激波锋面与支板前缘的距离
Fig.3 Distance between the shock front and the strut leading edge under different passivation radiuses

图4 不同钝化半径下的马赫数<br/>Fig.4 Mach number at different passivation radiuses

图4 不同钝化半径下的马赫数
Fig.4 Mach number at different passivation radiuses

图5 不同钝化半径下的测温点温度变化<br/>Fig.5 Temperature variation of measurement points under different passivation radiuses

图5 不同钝化半径下的测温点温度变化
Fig.5 Temperature variation of measurement points under different passivation radiuses

图6 前缘驻点处的热流密度<br/>Fig.6 Heat flux at the stationary point

图6 前缘驻点处的热流密度
Fig.6 Heat flux at the stationary point

图7 不同喷注总压下的流线图<br/>Fig.7 Streamline diagram under different injection total pressures

图7 不同喷注总压下的流线图
Fig.7 Streamline diagram under different injection total pressures

图8 不同喷注总压下激波锋面与支板前缘的距离<br/>Fig.8 Distance between the shock front and the strut leading edge under different total pressures

图8 不同喷注总压下激波锋面与支板前缘的距离
Fig.8 Distance between the shock front and the strut leading edge under different total pressures

图9 不同喷注总压下的静温图<br/>Fig.9 Static temperature under different injection total pressures

图9 不同喷注总压下的静温图
Fig.9 Static temperature under different injection total pressures

$图10 不同喷注总压下的H2O质量分数图<br/>Fig.10 Mass fraction of H2O under different injection total pressures$

图10 不同喷注总压下的H2O质量分数图
Fig.10 Mass fraction of H2O under different injection total pressures

图11 不同喷注总压下的测温点温度变化<br/>Fig.11 Temperature variation of measurement points under different injection total pressures

图11 不同喷注总压下的测温点温度变化
Fig.11 Temperature variation of measurement points under different injection total pressures

图12 不同喷孔尺寸下的流线图<br/>Fig.12 Streamline diagram under different nozzle sizes

图12 不同喷孔尺寸下的流线图
Fig.12 Streamline diagram under different nozzle sizes

图13 不同喷孔尺寸下的静温图<br/>Fig.13 Static temperature under different nozzle sizes

图13 不同喷孔尺寸下的静温图
Fig.13 Static temperature under different nozzle sizes

图14 不同喷孔尺寸下的测温点温度变化<br/>Fig.14 Temperature variation of measurement points under different nozzle sizes

图14 不同喷孔尺寸下的测温点温度变化
Fig.14 Temperature variation of measurement points under different nozzle sizes

图15 喷注总温为1 000 K时的静温图<br/>Fig.15 Static temperature when Tojet=1 000 K

图15 喷注总温为1 000 K时的静温图
Fig.15 Static temperature when Tojet=1 000 K

图16 不同喷注总温下的测温点温度变化<br/>Fig.16 Temperature variation of measurement points under different injection total temperatures

图16 不同喷注总温下的测温点温度变化
Fig.16 Temperature variation of measurement points under different injection total temperatures

[1] TOMIOKA S, MURAKAMI A, KUDO K, et al.Combustion tests of a staged supersonic combustor with a strut[J].Journal of propulsion & power, 2001, 17(2):293-300.
[2] BURNS R A.Development of scalar and velocity imaging diagnostics for supersonic hypermixing strut injector flowfields[D].Texas: University of Texas at Austin,2014.
[3] 刘世杰.超燃冲压发动机支板流场RANS/LES模拟及燃烧过程试验研究[D].长沙:国防科学技术大学,2007.
[4] 苏义.支板超声速混合增强技术及其阻力特性研究[D].长沙:国防科学技术大学,2006.
[5] PIZZARELLI M, NASUTI F, ONOFRI M.Coupled wall heat conduction and coolant flow analysis for liquid rocket engines[J].Journal of propulsion & power, 2012, 29(1):34-41.
[6] ZHONG F Q, CHEN T Y, FAN X J, et al.Three-dimensional heat transfer analysis and optimized design of actively cooled strut for scramjet applications: AIAA2012-3926[R].Reston: AIAA,2012.
[7] 黄拯.高温与超音速条件下发汗冷却基础问题研究[D].北京:清华大学,2015.
[8] 熊宴斌.超声速主流条件发汗冷却的流动和传热机理研究[D].北京:清华大学,2013.
[9] 宁东坡.超声速燃烧室长时间工作稳燃支板性能研究[D].哈尔滨:哈尔滨工业大学,2015.
[10] 孙冰, 郑力铭.超燃冲压发动机支板热环境及热防护方案[J].航空动力学报, 2006, 21(2):336-341.

备注

引言

1 数值模拟条件

2 支板热环境分析

3 逆向气体喷注热防护

4 结论

期刊信息