高超声速风洞连续变动压舵面颤振试验

季辰; 赵玲; 朱剑; 刘子强; 李锋

doi:10.11729/syltlx20170088

高超声速风洞连续变动压舵面颤振试验

中国航天空气动力技术研究院, 北京 100074

基金项目:

国家自然科学基金项目 11702285

详细信息

作者简介:
季辰(1982-), 男, 江苏南通人, 高级工程师。研究方向:气动弹性试验。通信地址:北京市7201信箱16分箱(100074)。E-mail:jichen167@sina.com

通讯作者:
季辰, E-mail: jichen167@sina.com

中图分类号: V211.74;V215.3⁺4
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出版历程
- 收稿日期: 2017-07-02
- 修回日期: 2017-08-30
- 刊出日期: 2017-12-24

Hypersonic wind tunnel flutter test research on rudder models by continuously varying dynamic pressure

China Academy of Aerospace Aerodynamics, Beijing 100074, China

摘要

摘要: 为了研究舵、翼面高超声速颤振特性，中国航天空气动力技术研究院建立了高超声速风洞连续变动压颤振试验技术。对具有相同结构动力学和气动特性的舵面模型进行颤振试验，试验马赫数为4.95和5.95。试验中缓慢连续增加试验动压直至颤振发生，并由此获得颤振临界参数；采用短时傅里叶变换时频域分析法研究了试验中模型频率随动压变化的耦合特性，分析表明该模型在试验条件下发生了经典弯扭耦合颤振。试验中还采用亚临界试验数据对颤振余度法和阻尼外推法2种颤振边界预测技术进行了研究，2种方法在高超声速颤振试验中都显示了良好的预测精度。研究还表明，动压增加的速率对颤振边界的预测精度影响较小。采用红外热成像技术对模型的气动加热进行了研究，温度场测量显示舵面最高温度出现在舵根部前缘位置，舵前缘和舵面斜面中后部温度也较高；舵轴裸露在流场中的部分由于反射板附面层的影响其气动加热问题并不严重。
- 高超声速 /
- 颤振试验 /
- 颤振边界预测 /
- 气动弹性 /
- 风洞试验
Abstract: In order to study the hypersonic flutter behavior of rudder models, a hypersonic wind tunnel flutter test technique by continuously varying dynamic pressure was developed and experimentally studied in China Academy of Aerospace Aerodynamics. The models with the same structural and aerodynamic design were tested at Mach number 4.95 and 5.95. The flutter critical parameters were obtained by slowly increasing the dynamic pressure until flutter onset. The short-time-fourier-transform time-frequency domain analysis method was used to study the frequency coupling characteristics. The analysis shows that it is the classic flutter that the bending and torsion mode couples as the dynamic pressure increases. Based on the structural dynamic parameter identification method, the damping ratio extrapolation method and the flutter margin method were used to predict the flutter critical parameters with the subcritical data. Both methods show a good prediction accuracy. The results also indicate that the rate of increase of dynamic pressure has small effect on the prediction of the flutter boundary. The temperature field measurements show that the maximum temperature of the model appears at the leading edge of the wing root. The temperatures of the leading edge and the rear part of the slope of the rudder are also relatively high. The temperature of the leading edge of the rudder shaft exposed to the flow field is not high, which might be due to the influence of the reflector surface boundary layer.
- hypersonic /
- flutter test /
- flutter boundary prediction /
- aeroelasticity /
- wind tunnel test

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图 1 数据采集系统

Fig. 1 Data acquisition system

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图 2 红外热像仪

Fig. 2 Infrared thermal imager

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图 3 计算模态振型

Fig. 3 Calculation mode shapes

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图 4 模型纹影图

Fig. 4 Schlieren photograph of the model

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图 5 F5模型高超颤振试验(Ma=4.95)

Fig. 5 F5 hypersonic flutter test (Ma=4.95)

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图 6 F6模型高超颤振试验(Ma=5.95)

Fig. 6 F6 hypersonic flutter test (Ma=5.95)

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图 7 时频谱图(F5, Ma=4.95)

Fig. 7 Time-frequency spectrum (F5, Ma=4.95)

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图 8 时频谱图(F6, Ma=5.95)

Fig. 8 Time-frequency spectrum (F6, Ma=5.95)

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图 9 颤振余度法(F5模型，Ma=4.95)

Fig. 9 Flutter margin method (F5, Ma=4.95)

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图 10 颤振余度法(F6模型，Ma=5.95)

Fig. 10 Flutter margin method (F6, Ma=5.95)

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图 11 阻尼比外插颤振边界(F5模型，Ma=4.95)

Fig. 11 Damping ratio extrapolation (F5, Ma=4.95)

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图 12 阻尼比外插颤振边界(F6模型，Ma=5.95)

Fig. 12 Damping ratio extrapolation (F6, Ma=5.95)

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图 13 颤振发展速率

Fig. 13 Rate of flutter development

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图 14 动压增加速率

Fig. 14 Rate of dynamic pressure increase

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图 15 动压增速与颤振发展速率

Fig. 15 K_FM vs K_q plot

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图 16 F5模型表面温度分布(Ma=4.95)

Fig. 16 Temperature distribution of model F5 (Ma=4.95)

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图 17 F6模型表面温度分布(Ma=5.95)

Fig. 17 Temperature distribution of model F6 (Ma=5.95)

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图 18 F5模型表面温度随时间变化曲线

Fig. 18 The surface temperature of model F5 versus time

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图 19 F5模型表面温度随时间变化曲线

Fig. 19 The surface temperature of model F5 versus time

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表 1 模型模态参数

Table 1 Mode parameters

	Frequency/Hz		Frequency ratio	Damping ratio/(%)
	f₁	f₂	f₂/f₁	ξ₁	ξ₂
FEM	32.3	55.8	1.73	--	--
GVT F5	33.5	55.7	1.66	0.31	0.41
GVT F6	32.1	54.2	1.69	0.33	0.32

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表 2 模型颤振参数

Table 2 Flutter parameters

Model	Mach number	q_{f_m} /(10⁴Pa)	ρ /(kg·m^-3)	f_f/Hz	T_t/K
F5	4.95	4.372	0.1392	40.2	378.2
F6	5.95	4.412	0.1059	39.1	477.3

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表 3 不同方法得到的颤振动压

Table 3 Comparison of flutter dynamic pressures

Model	Measured	Flutter marginfunction		Damping ratio extrapolation
Model	q_{f_m}/(10⁴Pa)	q_{f_f}/(10⁴Pa)	δ/%	q_{f_d}/(10⁴Pa)	δ/%
F5	4.372	4.241	-3.0	4.261	-2.5
F6	4.412	4.305	-2.4	4.352	-1.4

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表 4 颤振动压增速影响

Table 4 Influence of the dynamic-pressure-increasing rate

Ma	K_q /(Pa·s^-1)	K_FM /(s^-4·Pa^-1)	Model
4.95	145	57617	F5-1
4.95	767	58442	F5
4.95	2393	55428	F5-2

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