高超声速进气道双侧分离建立过程的试验研究

唐啸; 范晓樯; 熊冰; 王良

doi:10.13224/j.cnki.jasp.20240029

高超声速进气道双侧分离建立过程的试验研究

doi: 10.13224/j.cnki.jasp.20240029

国防科技大学高超声速技术实验室，长沙 410073

基金项目: 国家自然科学基金（11872071，12102470）；国防科技大学基金（ZK20-05）

详细信息

作者简介:
唐啸（1990-），男，博士生，主要从事高速飞行器内外流一体化设计及研究。E-mail：Xiaotang11@hotmail.com

通讯作者:
熊冰（1991-），男，副研究员，博士，主要从事高超声速推进技术研究。E-mail：xiongbing10@nudt.edu.cn

中图分类号: V211.48
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出版历程
- 收稿日期: 2024-01-10
- 网络出版日期: 2025-05-28

Experimental research on establishment of dual-separation unstart structure in hypersonic inlet

Hypersonic Technology Laboratory，National University of Defense Technology，Changsha 410073，China

摘要

摘要:
针对广泛存在于高超声速进气道中的机体/唇口双侧分离不起动流态，为了厘清其内部分离区的建立过程，揭示激发双侧分离的前提条件，基于风洞试验分析了增加攻角和堵撤反压两种方式构建该不起动流态的瞬态过程。研究表明：①无论是增加攻角还是堵撤反压过程，首先激发的均是机体侧大尺度分离区，而后才出现唇口侧大规模分离流动；②试验进气道的维持起动能力与双侧分离流态无直接关系，而是受制于喉部壅塞；③只有在机体侧有分离区存在，且其产生的分离激波入射至唇口压缩面特定范围之时，才有可能激发双侧分离流态。
- 流场演化 /
- 不起动流态 /
- 分离流动 /
- 双侧分离区 /
- 高超声速进气道 /
- 风洞试验
Abstract:
In order to clarify the establishment process of the internal separations for the unstart mode of dual-separation on both body and lip (DSBL), and reveal the prerequisite conditions for forming this unstart mode, two transient processes (increasing the angle of attack and blocking and then releasing downstream channel) for inducing DSBL were studied experimentally. The results were as follows: 1) for the two processes, the first induced separations were all located on the body side, and then large-scale separated flow appeared on the lip side; 2) the ability to maintain the start state of the inlet was not related to DSBL flow mode, but rather depended on choking of the throat; 3) DSBL was induced only when there were separations on the body-side, and the generated separation shock impacted to a specific range of the compression surface of the lip.
- flow-field evolution /
- unstart mode /
- separated flow /
- dual-separation /
- hypersonic inlet /
- wind tunnel experiment

HTML

图 1 类X-51A高超声速飞行器的两种进气道不起动形态

Figure 1. Two types of unstart modes of the inlet for X51-A class hypersonic aircraft

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图 2 机体加二维进气道一体化试验模型

Figure 2. Experiment model of aircraft which integrated forebody with a 2-dimensional air inlet

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图 3 试验风洞系统

Figure 3. Wind tunnel system of this experiment

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图 4 起动流态的维持及进入双侧分离不起动纹影图像（Ma₀=4）

Figure 4. Schlieren images of maintenance of start flow mode and the transformation from start to DSBL flow mode （Ma₀=4）

下载: 全尺寸图片幻灯片

图 5 增加攻角构建双侧分离过程的流场演化纹影图像（攻角从18°增加到20°）

Figure 5. Schlieren images of flow field evolution in process of establishing DSBL （increasing angle of attack from 18° to 20°）

下载: 全尺寸图片幻灯片

图 6 增加攻角构建双侧分离过程的流场演化压力变化图（攻角从18°增加到20°）

Figure 6. Changes of pressure of flow field evolution in process of establishing DSBL （increasing angle of attack from 18° to 20°）

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图 7 基础模型和对照模型

Figure 7. Basic model and contrast model

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图 8 激发和维持双侧分离流态的难易程度对比的纹影图像

Figure 8. Schlieren images of comparison of difficulty in inducing and maintaining DSBL flow mode

下载: 全尺寸图片幻灯片

图 9 对照模型起动流态的维持范围（Ma₀=4）

Figure 9. Maintenance of start flow mode of contrast model （Ma₀=4）

下载: 全尺寸图片幻灯片

图 10 节流反压构建双侧分离过程的流场演化（攻角为4°）

Figure 10. Flow field evolution in process of establishing DSBL from blocking and releasing downstream channel （angle of attack of 4°）

下载: 全尺寸图片幻灯片

图 11 节流反压构建双侧分离过程的流场演化（攻角为8°）

Figure 11. Flow field evolution in process of establishing DSBL from blocking and releasing downstream channel （angle of attack of 8°）

下载: 全尺寸图片幻灯片

图 12 激发双侧分离时唇口压缩面上激波的入射范围

Figure 12. Shock wave impact range on compression surface of lip when DSBL is about to be induced

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图 13 进气道内通道的面积变化规律

Figure 13. Variation law of inner channel area of inlet

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表 1 压力测点及位置信息

Table 1. Details of pressure measuring points and their location distribution

压力测点编号	x/l	位置
1	0.406	机体
2	0.451	机体
3	0.5	机体
4	0.543	机体
5	0.584	机体
6	0.607	机体
7	0.659	机体
8	0.597	唇口
9	0.617	唇口
10	0.637	唇口
11	0.657	唇口
12	0.677	唇口
13	0.697	唇口
14	0.714	唇口
高频1	0.706	唇口
高频2	0.633	机体
高频3	0.684	机体

下载: 导出CSV

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