高超声速飞行器前缘主动冷却影响因素

罗世彬; 庙智超; 宋佳文

doi:10.7527/S1000-6893.2022.27023

航空学报 >

2022 , Vol. 43 >Issue 12: 627023 - 627023

DOI: https://doi.org/10.7527/S1000-6893.2022.27023

临近空间高速飞行器数值模拟技术专栏

高超声速飞行器前缘主动冷却影响因素

罗世彬 ,
庙智超 ,
宋佳文

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中南大学航空航天学院, 长沙 410083

收稿日期: 2022-02-06

修回日期: 2022-03-07

网络出版日期: 2022-03-22

基金资助

翼型、叶栅空气动力学重点实验室基金（614220121020114）；湖南省重点研发计划（2021GK2011）

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Influencing factors of active cooling at leading edge of hypersonic vehicles

LUO Shibin ,
MIAO Zhichao ,
SONG Jiawen

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School of Aeronautics and Astronautics, Central South University, Changsha 410083, China

Received date: 2022-02-06

Revised date: 2022-03-07

Online published: 2022-03-22

Supported by

Foundation of National Key Laboratory of Science and Technology on Aerodynamic Design and Research(614220121020114); Key R&D Projects of Hunan Province (2021GK2011)

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摘要

主动冷却技术是目前高超声速飞行器发展的重点方向，对不同的主动冷却方式进行组合可实现优势互补，能为承受热流密度极高的飞行器前缘提供有效的热防护。以高超声速飞行器前缘气膜-发散组合冷却结构为研究对象，建立CFD数值计算模型，研究来流攻角为0°、4°和12°时对组合冷却效果的影响，对不同前缘上楔角构型的组合冷却效果进行分析。结果表明，攻角的出现使前缘模型上下半段的温差增大，上下壁面的温差最高可达639.2 K，攻角的改变通过影响外壁面压力的分布来影响结构中冷却剂的流量分配。增大前缘上楔角会使冷却剂向多孔介质下游输运的距离减小，外壁面温度与上楔角之间呈现近似线性增长的趋势。

关键词： 高超声速飞行器; 气膜冷却; 发散冷却; 来流攻角; 前缘楔角

本文引用格式

罗世彬 , 庙智超 , 宋佳文 . 高超声速飞行器前缘主动冷却影响因素[J]. 航空学报, 2022 , 43(12) : 627023 -627023 . DOI: 10.7527/S1000-6893.2022.27023

Abstract

The active cooling technology is the key development direction of hypersonic vehicles, and the combination of different active cooling technologies can achieve complementary advantages and provide effective thermal protection for the leading edge of vehicles with high heat flux. This study examines the film-transpiration combined cooling structure on the leading edge of hypersonic vehicles. The CFD numerical calculation model is established to study the influence of angles of attack of 0°, 4°, and 12° on the combined cooling effect, and the combined cooling effect of configurations with different upper wedge angles on the leading edge is analyzed. The results show that the angle of attack increases the temperature difference between the upper and lower halves of the leading edge model. The maximum temperature difference between the upper and lower walls is 639.2 K. The change in the angle of attack affects the flow distribution of the coolant in the structure by influencing the distribution of outside wall pressure. Increasing the upper wedge angle of the leading edge will reduce the distance of the coolant to the downstream of the porous medium. There is an approximate linear growth trend of the outside wall temperature with the increase of the upper wedge angle.

Key words： hypersonic vehicle; film cooling; transpiration cooling; angle of attack; wedge angle of leading edge

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