Experimental study on fine thermal measurement of high-speed aircraft wing rudder gapin shock wave tunnel

WU Ningning; KANG Honglin; LUO Jinling

doi:10.7638/kqdlxxb-2018.0156

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Experimental study on fine thermal measurement of high-speed aircraft wing rudder gapin shock wave tunnel

更新时间：2026-05-28

- Experimental study on fine thermal measurement of high-speed aircraft wing rudder gapin shock wave tunnel
- Vol. 37, Issue 1, Pages: 133-139(2019)
- 作者机构：
  
  北京空天技术研究所,北京,100074
- 作者简介：
- 基金信息：
- DOI：10.7638/kqdlxxb-2018.0156
  CLC： V211.3
- Published：2019
- 稿件说明：
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吴宁宁, 康宏琳, 罗金玲. 高速飞行器翼舵缝隙激波风洞精细测热试验研究[J]. 空气动力学学报, 2019, 37(1): 133-139.

WU Ningning, KANG Honglin, LUO Jinling. Experimental study on fine thermal measurement of high-speed aircraft wing rudder gapin shock wave tunnel[J]. 2019, 37(1): 133-139.
吴宁宁, 康宏琳, 罗金玲. 高速飞行器翼舵缝隙激波风洞精细测热试验研究[J]. 空气动力学学报, 2019, 37(1): 133-139. DOI： 10.7638/kqdlxxb-2018.0156.

WU Ningning, KANG Honglin, LUO Jinling. Experimental study on fine thermal measurement of high-speed aircraft wing rudder gapin shock wave tunnel[J]. 2019, 37(1): 133-139. DOI： 10.7638/kqdlxxb-2018.0156.

摘要

高速飞行器的气动控制翼舵面，为了转动灵活，在弹体和翼舵面之间存在缝隙。缝隙的存在会导致高速热气流进入，在舵轴根部产生强分离再附区域，形成高热、高压、高剪切严酷热环境，对飞行器的热防护提出了很高要求。由于影响翼舵缝隙流动的因素十分复杂，缝隙内热环境的准确预测非常困难。目前传统的激波风洞缝隙测热试验受限于薄膜热流传感器2 mm直径，只能在分离再附区布置有限测点，无法捕捉到热流峰值，导致计算与试验存在较大偏差。本文根据缝隙分离再附区热环境特点，针对精细测量的可行性，从传感器选取、测点布置方案、测量及数据后处理等方面进行了详细分析，提出了分布式热电偶精细测量方法，实现了采用点测热达到面测热的效果。针对简化的圆柱弹身加舵面的模型，完成翼舵缝隙精细测热试验，获得了翼舵干扰区峰值热流。试验研究了不同缝隙高度、舵偏角、迎角对翼舵干扰区热环境的影响规律，试验结果表明：翼舵缝隙对弹身干扰主要集中在舵轴干扰区。舵轴干扰区热环境随着缝隙高度的增加而增强，随着舵偏角和迎角的增大而增大。同时，试验结果与CFD计算结果对比表明，两者基本吻合。

Abstract

The pneumatic control wing rudder surface of a high-speed aircraft has a gap between the projectile and the wing rudder surface for flexibility. The existence of the gap leads to the entry of high-speed hot airflow as well as the strong separation and reattachment area at the root of the rudder shaft with high heat

high pressure

and high shear thermal environment. These severe environments require high thermal protection of the aircraft. Since the factors affecting the flow of the rudder gap are very complicated

accurate prediction of the thermal environment inside the gap is very difficult. At present

the traditional shock wave test of the wind tunnel is limited by the 2 mm diameter of the thin film heat flow sensor. Only limited measurement points can be arranged in the separation and reattachment area

and the peak of the heat flow cannot be captured

resulting in large differences between the calculation and the test. According to the characteristics of the thermal environment of the gap separation and re-attachment area

this paper analyzes the feasibility of fine measurement regarding sensor selection

measurement point layout scheme

measurement

and data post-processing. Moreover

a distributed thermocouple fine measurement method is proposed to realize the use of spot heat measurement for the effect of surface heat. Aiming at the simplified model of the cylindrical body and the rudder surface

the fine thermal test of the rudder gap is completed

and the peak heat flow in the rudder interference zone is obtained. The effects of different gap heights

rudder angles

and angles of attack on the thermal environment of the rudder interference zone are studied experimentally. The experimental results show that the rudder gap interference on the missile body is mainly concentrated in the rudder axle interference zone. The thermal environment of the rudder axis interference zone increases with increasing gap height

rudder angle

and angle of attack. The comparison between the test results and the CFD calculation results shows that they are basically consistent

and the peak heat flow difference is within 10%.

关键词

Keywords

references

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