Development of forebody asymmetric vortex control based on alternating synthetic jet and the verification on model free flight
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摘要: 为推进前体非对称涡流动控制方法在飞行器大迎角飞行控制方面的应用，提出并发展了一种基于双合成射流的前体非对称涡控制技术。研发了一套机载型双合成射流控制装置及模型自由飞验证机，通过风洞半自由飞及模型自由飞实验，验证了利用前体非对称涡控制技术实现尾旋改出和大迎角姿态控制的可行性；同时，依靠飞行测控系统和机载压力测量系统，实现飞行器姿态及前体表面压力的同步测量，可对前体非对称涡控制效能进行有效评估。风洞半自由飞实验结果表明：在60°迎角下，双合成射流可有效控制前体非对称涡相对位置，产生偏航力矩，实现大迎角航向操纵。在模型自由飞实验中，该技术可在常规方向舵失效的迎角下实现尾旋改出，可控尾旋角速度达到173 (°)/s；依靠该技术，验证机可在大迎角飞行时进行快速偏航操控，由控制输入到偏航角速度改变的时滞小于0.5 s。Abstract: In order to apply the forebody vortex flow control method to high Angle of Attack (AoA) flight control of the aircraft, an asymmetric vortex control technique based on the Alternating Synthetic Jet (ASJ) flow was proposed and developed. A set of airborne alternating synthetic jet control device and a free flight verification model aircraft were developed. The feasibility of using the forebody vortex control method to realize tail spin out and high Angle of Attack attitude control was verified by semi-free flight in the wind tunnel and model free flight in open airspace. Meanwhile, by means of the flight measurement and control system and the airborne pressure measurement system, the aircraft attitude, vortex position and body surface pressure can be measured synchronously, which can effectively evaluate the efficiency of the vortex control technology. Wind tunnel semi-free flight test results show that the alternating synthetic jet can effectively control the relative position of the forebody vortices at 60° Angle of Attack, which can generate yaw moment and realize heading control at high Angle of Attack. In the flight test, the technology can realize the change of tail spin under the failure of conventional rudder, and the controllable tail spin angular velocity can reach 173 (°)/s. Based on this technology, the verification model aircraft can perform fast yaw control when flying at high Angle of Attack, and the time delay from control input to yaw angular velocity change is less than 0.5 seconds.
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图 2 基于双合成射流的前体非对称涡控制验证机系统构成
Figure 2. The system of the verification model aircraft based on the ASJ forebody asymmetric vortex control technology
图 4 机载压力测量系统及飞行测控系统安装实物图
Figure 4. Airborne pressure measurement and flight control system
图 5 前体非对称涡控制验证机风洞半自由飞实验平台示意图
Figure 5. Schematic diagram of semi-free flight apparatus of model plane based on the ASJforebody asymmetric vortex control technology
图 6 前体非对称涡控制验证机风洞半自由飞实验平台
Figure 6. Semi-free flight apparatus of model plane based on the ASJ forebody asymmetric vortex control technology
图 7 前体非对称涡产生偏航力矩原理示意
Figure 7. Principle of yaw moment generated by forebody asymmetric vortex
图 8 不同输入信号下特征截面的平均速度分布
Figure 8. The average velocity distribution results of the characteristic section of the forebody under different input signals
图 9 不同前体非对称涡模态下空间流场、侧向力方向、表面压力分布
Figure 9. Comparison of flow field, side force and surface pressure under different forebody asymmetric vortex modes
图 10 风洞半自由飞输入信号、特征截面周向压力积分与偏航角加速度
Figure 10. The input signal, the circumferential pressure integral of the characteristic section and the yaw angle acceleration of the semi-free flight in the wind tunnel
图 11 双向尾旋改出机动的姿态–航迹图
Figure 11. Flight attitude and flight path diagram of bidirectional spin recovery maneuver
图 12 双向尾旋改出机动过程的控制输入与姿态角
Figure 12. The control input and attitude angle during bidirectional spin recovery maneuver process
图 13 双向尾旋改出机动过程的特征截面周向压力积分与偏航角加速度
Figure 13. The circumferential pressure integral of the characteristic section and yaw acceleration during bidirectional spin recovery maneuver process
图 14 大迎角航向机动控制中的输入信号、迎角和偏航角
Figure 14. Input signal, angle of attack and yaw angle in high angle of attack heading maneuver control
图 15 大迎角航向机动控制中的输入信号与偏航角速度
Figure 15. Input signal and yaw rate in high angle of attack heading maneuver control
表 1 基于双合成射流的前体非对称涡控制验证机总体参数
Table 1. The parameters of the verification model aircraft based on the ASJ forebody asymmetric vortex control technology
参数名称 参数值 翼展 1.2 m 实际起飞重量 3.5 kg 偏航转动惯量（Izz） 0.525 kg·m2 
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表 2 双合成射流激励器控制参数
Table 2. The parameters of the Alternating Synthetic Jet
参数名称 参数值 激励器质量 57.1 g 激励器尺寸 直径36 mm，长90 mm 喷口面积（2个孔合计） 8 mm2 激励波形 方波 激励频率 175 Hz 供电电压 12 V 最大功耗 1.3 W 射流最大时均速度 17 m/s
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表 3 机载压力测量系统参数
Table 3. The parameters of the airborne pressure measurement
参数名称 参数值 重量 80 g 尺寸 120 mm × 50 mm × 20 mm 测压精度 0.01 FS 量程 ± 500 Pa
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表 4 侧向力与特征截面周向压力积分的Pearson相关性系数
Table 4. Pearson correlation coefficient between side force and circumferential pressure integral of characteristic section
迎角范围 相关性系数 0°～15° 0.64 20°～35° 0.94 40°～65° 0.92 70°～85° 0.69
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