质量矩飞行器制导控制问题研究
发布时间：2018-12-14 13:23
【摘要】：导弹机动控制方式从本质上讲均通过调节控制力矩实现,具体实现形式包括两类:调节控制力和控制力臂。传统舵面控制及喷气推力控制均属于调节控制力的范畴,但舵面控制难以解决高速机动过程中的气动烧蚀问题,而喷气推力控制则受限于携带燃料有限且导致固液耦合等缺点。质量矩控制则属于调节控制力臂的范畴,因其执行机构在弹体内部避免了舵面控制气动烧蚀等问题,又因其无需携带额外燃料解决了喷气推力控制携带燃料有限及固液耦合等缺点。鉴于此质量矩控制较传统控制方式优势明显且应用前景广泛,但其独特的控制模式增加了导弹空间运动复杂程度,为制导控制设计带来诸多新的挑战。本文以质量矩飞行器再入机动精确打击为背景,旨在研究双滑块/差动副翼侧滑转弯(skid-to-turn,STT)质量矩飞行器动力学建模、制导控制及其一体化设计问题,从而为质量矩导弹技术的发展提供理论支撑。首先,基于多刚体系统建模方法建立了双滑块/差动副翼STT质量矩飞行器完整空间运动模型。充分考虑质量矩导弹的运动学耦合、惯性耦合及气动惯性交叉耦合等因素,基于动量定理建立了弹体坐标系下弹体质心平动动力学方程,基于动量矩定理建立了弹体坐标系下系统绕弹体质心转动动力学方程,同时建立了再入坐标系下弹体质心平动运动学方程及绕弹体质心转动运动学方程。这些方程完整地描述了质量矩飞行器空间运动机理,揭示了滑块运动与弹体质心平动及绕弹体质心转动的内在联系。与传统舵面控制及喷气控制相比,质量矩飞行器独特的控制机理,即滑块运动与弹体运动之间的内在耦合联系,使得质量矩飞行器空间运动模型较为复杂。其次,针对地面固定目标精确打击问题,分别提出了有/无终端角度约束的有限-r收敛制导律1。1)以弹目距离为参变量描述导弹与目标的相对运动关系,建立了新的制导模型。该制导模型包括两个微分方程,分别描述了视线俯冲运动及视线转弯运动,并且视线俯冲运动微分方程单独解耦。基于该模型的制导律设计既保证了精确性又简化了制导律设计过程。2)基于该制导模型,提出了具有干扰抑制的有限-r收敛制导律,给出了过载形式的制导指令。与传统制导律相比,该制导律理论上保证了视线旋转角速率在弹目距离减小至期望值之前收敛为零。仿真验证了该制导律的正确性。3)基于该制导模型,提出了具有终端角度约束的有限-r收敛制导律,给出了过载形式的制导指令。与传统制导律相比,该制导律理论上保证了视线角偏差及视线旋转角速率在弹目距离减小至期望值之前收敛为零。与比例导引(proportional navigation guidance,PNG)和最优导引(optimal navigation guidance,ONG)的仿真比较结果验证了该制导律的优越性。然后,针对质量矩飞行器姿态控制问题,分别提出了有/无控制输入饱和的有限时间收敛控制律。1)在合理简化基础上建立了质量矩导弹俯仰、偏航和滚转三通道独立控制模型。基于该控制模型,分别设计了俯仰、偏航和滚转通道有限时间收敛姿态控制律。与传统质量矩导弹姿态控制律相比,该控制律理论上保证了姿态角偏差及其变化率有限时间收敛于原点邻域。2)通过引入扩张状态观测器实现对扰动边界的自适应估计,进而设计了俯仰、偏航和滚转通道控制输入饱和的有限时间收敛姿态控制律。与传统质量矩导弹姿态控制律相比,该控制律理论上保证了控制输入饱和情况下姿态角偏差及其变化率有限时间收敛于原点邻域。通过特征点仿真和全弹道仿真验证了所提有/无控制输入饱和的有限时间收敛姿态控制律的正确性。最后,针对地面逃逸目标精确打击问题,分别提出了三通道独立/全状态耦合制导控制一体化设计。1)建立了质量矩飞行器俯仰、偏航和滚转三通道独立制导控制一体化模型。基于鲁棒自适应反演方法设计了一体化控制律,通过鲁棒自适应函数项实现对有界未知扰动的自适应补偿,引入非线性跟踪微分器避免了传统反演控制的“计算膨胀”问题,理论上证明了系统状态跟踪误差及扰动估计误差指数收敛于原点邻域。仿真验证了所提三通道独立制导控制一体化模型及控制律的正确性。2)建立了质量矩飞行器俯仰、偏航和滚装全状态耦合制导控制一体化模型。基于自适应动态面反演方法设计了一体化控制律,通过鲁棒自适应函数项实现对有界未知扰动的自适应补偿,引入动态面技术避免了传统反演控制的“计算膨胀”问题,理论上证明了系统状态跟踪误差、边界层误差及扰动估计误差指数收敛于原点邻域。仿真验证了所提全状态耦合制导控制一体化模型及控制律的正确性。
[Abstract]:The maneuvering control mode of the missile is realized by adjusting the control moment in nature, and the specific implementation form includes two types: adjusting the control force and the control force arm. The control of the conventional rudder surface and the control of the jet thrust belong to the category of the control force, but the control of the rudder surface is difficult to solve the problem of the aerodynamic ablation in the high-speed maneuver, and the jet thrust control is limited by the disadvantages of the limited carrying of the fuel and the coupling of the solid solution. the control of the mass moment belongs to the category of adjusting the control force arm, so that the problem that the rudder surface is controlled by the rudder surface and the like is avoided in the body of the elastic body due to the actuator of the control force arm, and the defect that the jet thrust control carries the limited fuel and the solid solution coupling and the like is solved by the air jet thrust control without carrying the extra fuel. In view of the obvious advantage of this quality moment control and wide application prospect, its unique control mode increases the complexity of the missile space motion and brings many new challenges to the guidance control design. The purpose of this paper is to study the dynamic modeling, guidance control and integrated design of the mass moment of the two-block/ differential aileron sideslip-turn (STT), so as to provide the theoretical support for the development of the mass-moment missile technology. First, a complete space motion model of the two-block/ differential aileron STT mass moment aircraft is established based on the multi-rigid-body system modeling method. Taking full consideration of the factors such as the kinematic coupling, the inertia coupling and the pneumatic inertia cross-coupling of the mass moment missile, the kinetic equation of the elastic body motion in the body coordinate system is established based on the momentum theorem. Based on the momentum moment theorem, the dynamic equations of the body rotation of the system in the body coordinate system are established, and the kinematic equations of the core translation of the elastic body in the re-input coordinate system and the kinematic equations about the core rotation of the elastic body are established. These equations describe the mechanism of the space motion of the mass moment aircraft, and reveal the internal relation between the movement of the slide block and the movement of the elastic body and the rotation of the body of the elastic body. Compared with the conventional rudder surface control and the air-jet control, the unique control mechanism of the mass moment aircraft, that is, the inherent coupling between the movement of the slide block and the body motion, makes the space motion model of the mass moment aircraft more complex. Secondly, aiming at the problem of target fixed target, the finite-r convergent guidance law with/ without terminal angle constraint is proposed. The relative motion between the missile and the target is described by using the target distance as the reference variable, and a new guidance model is established. The guidance model includes two differential equations, respectively describing the line-of-sight subduction motion and the line-of-sight turning motion, and the line-of-sight subduction motion differential equation is separately decoupled. The guidance law design based on the model not only ensures the accuracy but also simplifies the guidance law design process. 2) Based on the guidance model, a limited-r convergence guidance law with interference suppression is proposed, and the guidance instruction in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the rotation angle rate of the line of sight converges to zero before the target distance is reduced to the expected value. The validity of the guidance law is verified by the simulation. 3) Based on the guidance model, a limited-r convergence guidance law with terminal angle constraint is proposed, and the guidance command in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the line-of-sight angle deviation and the line-of-sight rotation angle rate converge to zero before the target distance is reduced to the expected value. The superiority of the guidance law is verified by the simulation results of the proportional guidance (PNG) and the optimal navigation guide (ONG). In this paper, a finite-time convergence control law with/ without control input saturation is proposed for the attitude control of the mass moment aircraft, and a three-channel independent control model of the pitching, yaw and rolling of the mass moment missile is established on the basis of a reasonable simplification. Based on the control model, the attitude control law with limited time for pitching, yaw and rolling is designed. in comparison with that attitude control law of the traditional mass moment missile, the control law theoretically ensure that the attitude angle deviation and the rate of change of the attitude angle converge to the origin neighborhood. The yaw and roll channel controls the limited time-convergence attitude control law of the input saturation. Compared with the conventional attitude control law of the mass moment, the control law theory ensures that the attitude angle deviation and the change rate finite time in the control input saturation condition converge to the origin neighborhood. The correctness of the finite-time convergence attitude control law with/ without control input saturation is verified by the characteristic point simulation and the full-trajectory simulation. in that end, the integrated design of three-channel independent/ full-state coupled guidance control is put forward, and the integrated model of independent guidance control for the pitch, yaw and roll of the mass moment is established. The integral control law is designed based on the self-adaptive inversion method of the Rurod, the self-adaptive compensation of the unknown disturbance is realized through the self-adaptive function of the Rurod, the problem of the 鈥淐alculate expansion鈥，
本文编号：2378684
[Abstract]:The maneuvering control mode of the missile is realized by adjusting the control moment in nature, and the specific implementation form includes two types: adjusting the control force and the control force arm. The control of the conventional rudder surface and the control of the jet thrust belong to the category of the control force, but the control of the rudder surface is difficult to solve the problem of the aerodynamic ablation in the high-speed maneuver, and the jet thrust control is limited by the disadvantages of the limited carrying of the fuel and the coupling of the solid solution. the control of the mass moment belongs to the category of adjusting the control force arm, so that the problem that the rudder surface is controlled by the rudder surface and the like is avoided in the body of the elastic body due to the actuator of the control force arm, and the defect that the jet thrust control carries the limited fuel and the solid solution coupling and the like is solved by the air jet thrust control without carrying the extra fuel. In view of the obvious advantage of this quality moment control and wide application prospect, its unique control mode increases the complexity of the missile space motion and brings many new challenges to the guidance control design. The purpose of this paper is to study the dynamic modeling, guidance control and integrated design of the mass moment of the two-block/ differential aileron sideslip-turn (STT), so as to provide the theoretical support for the development of the mass-moment missile technology. First, a complete space motion model of the two-block/ differential aileron STT mass moment aircraft is established based on the multi-rigid-body system modeling method. Taking full consideration of the factors such as the kinematic coupling, the inertia coupling and the pneumatic inertia cross-coupling of the mass moment missile, the kinetic equation of the elastic body motion in the body coordinate system is established based on the momentum theorem. Based on the momentum moment theorem, the dynamic equations of the body rotation of the system in the body coordinate system are established, and the kinematic equations of the core translation of the elastic body in the re-input coordinate system and the kinematic equations about the core rotation of the elastic body are established. These equations describe the mechanism of the space motion of the mass moment aircraft, and reveal the internal relation between the movement of the slide block and the movement of the elastic body and the rotation of the body of the elastic body. Compared with the conventional rudder surface control and the air-jet control, the unique control mechanism of the mass moment aircraft, that is, the inherent coupling between the movement of the slide block and the body motion, makes the space motion model of the mass moment aircraft more complex. Secondly, aiming at the problem of target fixed target, the finite-r convergent guidance law with/ without terminal angle constraint is proposed. The relative motion between the missile and the target is described by using the target distance as the reference variable, and a new guidance model is established. The guidance model includes two differential equations, respectively describing the line-of-sight subduction motion and the line-of-sight turning motion, and the line-of-sight subduction motion differential equation is separately decoupled. The guidance law design based on the model not only ensures the accuracy but also simplifies the guidance law design process. 2) Based on the guidance model, a limited-r convergence guidance law with interference suppression is proposed, and the guidance instruction in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the rotation angle rate of the line of sight converges to zero before the target distance is reduced to the expected value. The validity of the guidance law is verified by the simulation. 3) Based on the guidance model, a limited-r convergence guidance law with terminal angle constraint is proposed, and the guidance command in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the line-of-sight angle deviation and the line-of-sight rotation angle rate converge to zero before the target distance is reduced to the expected value. The superiority of the guidance law is verified by the simulation results of the proportional guidance (PNG) and the optimal navigation guide (ONG). In this paper, a finite-time convergence control law with/ without control input saturation is proposed for the attitude control of the mass moment aircraft, and a three-channel independent control model of the pitching, yaw and rolling of the mass moment missile is established on the basis of a reasonable simplification. Based on the control model, the attitude control law with limited time for pitching, yaw and rolling is designed. in comparison with that attitude control law of the traditional mass moment missile, the control law theoretically ensure that the attitude angle deviation and the rate of change of the attitude angle converge to the origin neighborhood. The yaw and roll channel controls the limited time-convergence attitude control law of the input saturation. Compared with the conventional attitude control law of the mass moment, the control law theory ensures that the attitude angle deviation and the change rate finite time in the control input saturation condition converge to the origin neighborhood. The correctness of the finite-time convergence attitude control law with/ without control input saturation is verified by the characteristic point simulation and the full-trajectory simulation. in that end, the integrated design of three-channel independent/ full-state coupled guidance control is put forward, and the integrated model of independent guidance control for the pitch, yaw and roll of the mass moment is established. The integral control law is designed based on the self-adaptive inversion method of the Rurod, the self-adaptive compensation of the unknown disturbance is realized through the self-adaptive function of the Rurod, the problem of the 鈥淐alculate expansion鈥，
本文编号：2378684
本文链接：https://www.wllwen.com/shoufeilunwen/gckjbs/2378684.html
