激波风洞7°尖锥边界层转捩实验研究

栗继伟; 卢盼; 汪球; 赵伟

doi:10.13700/j.bh.1001-5965.2019.0577

激波风洞7°尖锥边界层转捩实验研究

doi: 10.13700/j.bh.1001-5965.2019.0577

栗继伟¹,
卢盼^{1, 2},
汪球^{1, 2, ,},
赵伟^{1, 2}

1.
中国科学院力学研究所高温气体动力学国家重点实验室, 北京 100190
2.
中国科学院大学工程科学学院, 北京 100049

基金项目:

国家自然科学基金 11402275

国家自然科学基金 11727901

国家重点研发计划 2016YFA0401201

详细信息

作者简介:
栗继伟男, 硕士, 工程师。主要研究方向:高超声速边界层转捩、风洞测试技术

汪球男, 博士, 副研究员。主要研究方向:高焓流动气动热

通讯作者:
汪球, E-mail: wangqiu@imech.ac.cn

中图分类号: V411.7
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- 被引次数: 0
出版历程
- 收稿日期: 2019-11-11
- 录用日期: 2019-12-22
- 网络出版日期: 2020-11-20

Experimental study of boundary-layer transition on a 7° sharp cone in shock tunnel

LI Jiwei¹,
LU Pan^{1, 2},
WANG Qiu^{1, 2
, ,},
ZHAO Wei^{1, 2}

1.
State Key Laboratory of High-temperature Gas Dynamics, Institute of Mechanics, Chinese Academy of Sciences, Beijing 100190, China
2.
School of Engineering Science, University of Chinese Academy of Sciences, Beijing 100049, China

Funds:

National Natural Science Foundation of China 11402275

National Natural Science Foundation of China 11727901

National Key R & D Program of China 2016YFA0401201

More Information

Corresponding author: WANG Qiu, E-mail: wangqiu@imech.ac.cn

摘要

摘要:
高超声速边界层转捩对摩阻、传热等有重要影响，飞行器的研制迫切希望能精确预测和控制边界层转捩。在JF8A激波风洞中开展了7°半锥角的高超声速尖锥边界层转捩实验研究，利用响应频率达到1 MHz量级的高频压力传感器对尖锥壁面脉动压力进行了测量，并结合热流测量结果，研究了高超声速尖锥边界层中扰动波的发展过程。实验结果表明：JF8A激波风洞在雷诺数为6.4×10⁶/m状态下核心流的自由流噪声为2.8%；高频脉动压力测量技术能清晰地捕捉转捩过程中的第二模态波及其发展历程，试验状态下模型的第二模态波频率范围为165~206 kHz。当前研究结果能够为高超声速数值方法验证提供数据支撑。
- 激波风洞 /
- 高超声速 /
- 边界层转捩 /
- 尖锥 /
- 第二模态波
Abstract:
Hypersonic boundary layer transition has an important influence on friction drag and heat transfer, and thus accurate prediction and control of boundary layer transition are critical to the development of hypersonic vehicles. In this paper, experimental study on a 7° half-angle sharp cone was conducted in the JF8A shock tunnel to investigate the hypersonic boundary layer transition. The wall fluctuation pressure was measured by the pressure transducer with the response frequency as high as 1 MHz, and the development process of the disturbance wave in hypersonic sharp cone boundary layer was also investigated together with the results of heat flux measurement. The experimental results show that the pressure fluctuation in the free flow is 2.8% under the test condition of Reynolds number equals to 6.4×10⁶/m. Structures and development of the second mode waves in the process of transition can be obtained by the high-frequency fluctuation pressure measurement technique. The characteristic frequency of the second mode wave changes from 165 kHz to 206 kHz under the present test condition. The current research results can provide data support for hypersonic numerical method validation.
- shock tunnel /
- hypersonic /
- boundary layer transition /
- sharp cone /
- the second mode wave

HTML全文

图 1 JF8A高超声速激波风洞

Figure 1. JF8A hypersonic shock tunnel

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图 2 试验典型压力曲线

Figure 2. Typical pressure histories in test

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图 3 尖锥模型及传感器安装示意图

Figure 3. Schematic diagram of sensor installation on sharp cone model

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图 4 原始及滤波后的PCB压力传感器信号

Figure 4. Raw and filtered PCB pressure sensor signals

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图 5 尖锥母线热流分布

Figure 5. Heat-flux distribution on sharp cone generatrix

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图 6 不同位置PCB脉动压力信号

Figure 6. Pressure fluctuation signals of PCB at different positions

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图 7 典型脉动压力功率谱及背景噪声

Figure 7. Typical fluctuation pressure PSD and background noise

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图 8 模型子午线的功率谱结果

Figure 8. PSD along model meridian

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图 9 高、低频脉动压力

Figure 9. High- and low-frequency fluctuation pressure

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表 1 风洞试验状态

Table 1. Test conditions of wind tunnel

状态参数		数值
驻室参数	P₀/MPa	1.2
	H₀ /(MJ·kg^-1)	0.62
	T₀/K	620
自由流参数	ρ_∞/(kg·m^-3)	2.7×10^-2
	u_∞/(m·s^-1)	1 054
	T_∞/K	67.5
	Ma	6.5
	Re/m^-1	6.4×10⁶

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