应用分力合成主动振动抑制方法的最优飞行器大角度机动控制策略

 针对当代带大型挠性附件的空间飞行器,提出了分力合成主动振动抑制方法,并且分析了方法的鲁棒性。该方法可以保证挠性飞行器在实现指定的刚体运动的同时,抑制掉对系统影响较大的挠性振动模态,对频率不确定性的鲁棒性使得该方法易于工程实践。对于使用常幅值力矩喷气执行机构的航天器,设计了应用分力合成方法的时间—燃料最优机动控制律,数值仿真结果验证了方法的有效性。     

挠性航天器大角度机动的滑模变结构控制

考虑刚性主体上带有挠性梁的航天器，在建立挠性系统动力学模型的基础上，采用指数趋近率的滑模变结构控制策略进行大角度机动控制，并通过最优控制理论设计弹性稳态器抑制由于刚体运动而激发的弹性振动，数字仿真表明，提出的控制策略在实现旋转机动的同时，有效地抑制了弹性振动。     

基于输入成形的挠性航天器自适应滑模控制

针对大型挠性航天器姿态机动过程中的振动抑制问题,提出一种输入成形(IS)与自适应滑模控制(ASMC)相结合的控制策略.该控制策略利用输入成形抑制标称挠性系统的残余振动,并通过滑模控制保证实际系统在参数不确定性和外部干扰的影响下实现对标称系统的跟踪,解决了输入成形对参数不确定性和外部干扰的敏感性问题.进一步采用自适应技术去除了滑模切换增益对参数不确定性和干扰上界的先验性要求.仿真结果表明,在参数不确定性和外界干扰的影响下,该控制方法能够保证在完成姿态机动的同时抑制航天器的挠性振动.     

刚弹耦合飞行器的机动载荷减缓

针对飞行器刚体/弹性模态耦合作用下的机动载荷减缓问题,运用亚音速非定常空气动力学理论以及分析力学原理,建立了柔性飞机机动载荷分析及控制的气动弹性模型,分析了飞机作纵向机动飞行时结构弹性对机动飞行参数的影响。数值分析结果表明,结构弹性模态对飞机机动过程的影响不容忽视。为了有效减缓机动载荷,运用最优控制理论,采用多组控制面联合偏转的作动方案,设计机动载荷减缓最优控制律,实现了对飞机纵向机动载荷的控制。研究结果表明,提出的机动载荷减缓控制方案是有效的。     

基于反步法的挠性航天器姿态镇定

利用反步法研究了一类挠性航天器的姿态镇定问题,提出一种基于模态观测器的反步控制设计方案.首先,构造挠性模态观测器对挠性模态变量及其变化率进行观测;其次,将角速度看成虚拟控制器,设计虚拟角速度镇定运动学模型与挠性模态变量组成的子系统;最后,利用反步法设计了一种非线性控制器使得角速度能够跟踪虚拟角速度,从而实现姿态镇定的目...     

高超声速飞行器发射段运动模态特征分析

在发射段,吸气式高超声速飞行器和助推火箭构成了一个整体,其运动轨迹属机动轨迹,以此轨迹作为基准状态的运动模态受轨迹机动性的影响.以某吸气式高超声速飞行器和助推火箭组成的整体为研究对象,采用“系数冻结法”,选取机动轨迹跃升段上一个典型的冻结点,建立了线性模型.通过计算分析,得出了飞行器在该点的模态特征,并揭示了轨迹参数和飞行器结构的不对称性对运动模态的影响机理.所得结论对高超声速飞行器的控制器设计和发射轨迹选择有重要意义.     

空空导弹大角度姿态反作用喷气控制

为研究具有大离轴角及越肩发射能力的先进空空导弹初始段敏捷转弯方法,研究了装有反作用喷气控制系统的空空导弹的大角度姿态过失速机动控制律。反作用喷气控制系统用来提供大角度敏捷转弯时大攻角飞行的控制力矩。利用时间尺度分离的方法将导弹的姿态动力学和运动学系统分别看作快子系统和慢子系统。用李亚普诺夫方法设计了慢子系统控制律,利用滑动模态方法设计了快子系统控制律,在该控制律作用下,导弹闭环系统不仅是稳定的而且其动态品质也可以得到保证。分析了控制系统的鲁棒性,结果表明所提控制方法能够有效消除空空导弹大角度姿态机动时转动惯量变化以及各种力矩干扰的影响。最后给出了一个实例来说明姿态控制在空空导弹敏捷转弯中的应用。     

基于李亚普诺夫稳定性理论的压气机失速主动控制方法

针对压气机主动稳定控制方法中模态控制需要大量传感器及高频执行装置的不足,以喷气装置作为失速控制的执行机构,基于李亚普诺夫稳定性理论,利用回溯法设计了采用周向同步喷气的压气机预失速和过失速阶段的失速控制算法。理论分析和仿真结果表明,采用该方法,在预失速阶段,经过约1.1s持续喷气后,扰动的各阶模态的幅值均趋于0,各阶模态的相位均趋于恒定;在过失速阶段,持续约0.1s的喷气控制后,扰动的各阶模态被完全抑制,其各阶幅值趋于0,各阶模态相位趋于恒定;实现了对预失速和过失速的有效控制。由于以平均流量作为反馈输入,该方法只需安装少量传感器,且喷气装置的作动频率不高于50Hz,远低于模态控制方法。此外,采用同步喷气,也降低了执行机构的复杂性。     

升力体机动飞行器气动布局概念设计

开展高超声速机动飞行器的气动布局概念设计研究,提出一种新的升力体机动飞行器气动布局方案,进行了机体的优化设计,并进行了控制舵的匹配设计,研究了飞行器的气动特性和操纵效率问题。研究表明,该方案可以获得较高的配平升阻比及配平攻角,有较高的机动控制效率,是高超声速飞行器实现高机动的潜在可行方案。     

空间飞行器大角度机动飞行的变结构姿态控制

研究了空间飞行器大角度机动飞行的变结构姿态控制。应用四元数来描述姿态运动，以消除大角度机动飞行时欧拉角描述所存在的奇异性。基于Lyapunov方法设计了变结构控制的切换函数，以保证系统的滑动模态．亦即四元数偏差的稳定性。基于所给切换函数，设计了变结构控制器。数值仿真的结果说明了设计方法的有效性。     

Variable structure control and active vibration suppression of flexible spacecraft during attitude maneuver

This paper presents a new approach to vibration reduction of flexible spacecraft during attitude maneuver by using the theory of variable structure control (VSC) to design switching logic for thruster firing and lead zirconate titanate (PZT) as sensor and actuator for active vibration suppression. The spacecraft to be investigated is a hub with a cantilever flexible beam appendage, which can undergo a single axis rotation. The proposed control system includes the attitude controller acting on the rigid hub, designed by variable structure control technique, and the surface-bonded PZT patches for active vibration suppression of flexible appendages, designed by the positive position feedback (PPF) control technique. To avoid chattering, pulse-width pulse-frequency (PWPF) modulation is adopted for the thruster control, which makes the thrusters to be operated in a close to linear manner and also can suppress the relatively large amplitude vibrations excited by, for example, rapid maneuver. However, some residual micro-vibrations still exist due to the switching actions. Upon that, the technique of active vibration control using PZT is turned on to provide further vibration suppression of the residual micro-vibrations and fine tuning of the system performance. By combining the advantages of both these control strategies, an improved performance for vibration control in both the macro-and micro-senses can result. Both analytical and numerical results are presented to show the theoretical and practical merit of this approach.     

Acceleration sensors based modal identification and active vibration control of flexible smart cantilever plate

Some flexible appendages of spacecrafts, such as solar panels, are cantilever plate structures. Thus, vibration problem is unavoidable when there is slewing maneuver or external disturbance excitation. Vibration of such cantilever plate structures includes coupled bending and torsional motion. Furthermore, the low amplitude vibration near the equilibrium point is very difficult to be quickly suppressed due to nonlinear factors of the hardware in the system, which is harmful to stability and attitude control accuracy. To solve these problems, acceleration sensor-based modal identification and active vibration control methods are presented for the first two bending and the first two torsional modes vibration of the cantilever plate. Optimal placements of three acceleration sensors and PZT patches actuators are performed to decouple the bending and torsional vibration of such cantilever plate for sensing and actuating, and identifications are achieved by experiments. A nonlinear control method is presented to suppress both high and low amplitude vibrations of flexible smart cantilever plate significantly. Experimental comparison researches are conducted by using acceleration proportional feedback and the presented nonlinear control algorithms. The experimental results demonstrate that the presented acceleration sensor-based methods can suppress the vibration of cantilever plate effectively.     

基于多目标进化算法的卫星机动路径规划

建立了带挠性帆板卫星姿态控制问题的多目标优化模型,并基于一种多目标优化精英进化算法,规划卫星姿态机动的最优路径,以进一步提高卫星姿态控制系统的控制性能.仿真结果表明,所提方法只需运行一次,便能够有效地规划出一组多样性较好的非支配路径解,供决策者在不同的工作目标下选择,有效地缓解了卫星快速性、姿态稳定度、卫星机动所激发的帆板低振动强度之间的矛盾.      

柔性航天器振动主动抑制及姿态控制

针对柔性航天器柔性附件振动主动抑制及姿态高精度快速稳定问题，研究了一种输入成形器（IS）-自适应有限时间干扰观测器（FDO）-有限时间积分滑模控制器综合的设计方法。首先，基于柔性模态的频率及阻尼信息，获得能够有效抑制柔性振动的输入成形器形式，并与系统参考输入进行卷积，得到期望参考输入；其次，基于航天器动力学模型，设计一种新型的自适应有限时间干扰观测器，避免了综合干扰上界必须已知的约束，且保证干扰估计误差有限时间收敛至零，实现对干扰及残余振动影响的快速精确估计；最后，基于观测器的估计值，设计多变量有限时间积分滑模控制器，保证对期望参考输入的高精度快速跟踪控制，并进行严格的稳定性证明。仿真结果表明，该综合设计策略能够保证柔性附件振动抑制75%，姿态稳定度达到10^-4数量级。     

A composite control scheme for attitude maneuvering and elastic mode stabilization of flexible spacecraft with measurable output feedback

This paper treats the question of attitude maneuver control and elastic mode stabilization of a flexible spacecraft based on adaptive sliding mode theory and active vibration control technique using piezoelectric materials. More precisely, a modified positive position feedback (PPF) scheme is developed to design the PPF compensator gains in a more systematical way to stabilize the vibration modes in the inner loop, in which a cost function is introduced to be minimized by the feedback gains subject to the stability criterion at the same time. Based on adaptive sliding mode control theory, a discontinuous attitude control law is derived to achieve the desired position of the spacecraft, taking explicitly into account the mismatched perturbation and actuator constraints. In the attitude control law, an adaptive mechanism is also embedded such that the unknown upper bound of perturbation is automatically adapted. Once the controlled attitude control system reaches the switching hyperplane, the state variables can be driven into a small bounded region. An additional attractive feature of the attitude control method is that the structure of the controller is independent of the elastic mode dynamics of the spacecraft, since in practice the measurement of flexible modes is not easy or feasible. The proposed control strategy has been implemented on a flexible spacecraft. Both analytical and numerical results are presented to show the theoretical and practical merit of this approach.     

中继卫星H∞姿态稳定控制研究

 主要研究中继卫星H_∞回路成形姿态稳定控制问题。中继卫星具有弱阻尼的柔性模态频率和晃动模态频率,并且模态频率具有不确定性,同时系统增益也存在有较大的变化。首先利用拉格朗日方程建立了该卫星的动力学方程,描述了系统存在的结构不确定性,然后设计了H_∞回路成形控制器。在设计过程中,首先提出了控制系统所需要满足的鲁棒稳定性和鲁棒性能指标;随后利用H_∞回路成形理论设计了卫星的姿态稳定控制器,利用ν间隔度量方法对所设计的控制器进行了降阶处理,并且对所设计的控制器进行了鲁棒稳定性和鲁棒性能分析;最后通过数学仿真证明了所设计的控制器的有效性。     

Adaptive control and stabilization of elastic spacecraft

This work treats the question of large angle rotational maneuver and stabilization of an elastic spacecraft (spacecraft-beam-tip body configuration). It is assumed that the parameters of the system are completely unknown. An adaptive control law is derived for the rotational maneuver of the spacecraft. Using the adaptive controller, asymptotically decoupled control of the pitch angle of the space vehicle is accomplished, however this maneuver causes elastic deformation of the beam connecting the orbiter and tip body. For the stabilization of the zero dynamics (flexible dynamics), a stabilizer is designed using elastic mode velocity feedback. In the closed-loop system including the adaptive controller and the stabilizer, reference pitch angle trajectory tracking and vibration suppression are accomplished. Simulation results are presented to show the maneuver capability of the control system     

内装式空射火箭箭机分离RCS姿态控制方案研究

采用反作用喷气控制系统(RCS)对内装式空中发射运载火箭箭机分离阶段实施姿态控制的方案进行了研究.根据不同发射条件下的箭机分离运动,确定了RCS启控时运载火箭的飞行状态,设计了发动机点火前的运载火箭俯仰运动程序.在此基础上,给定标准弹道方案,并确定了RCS推力器组合方案.采用脉冲调宽调频(PWPF)调制方法,设计出适合...     

Detumbling and reorientation maneuvers and stabilization of NASASCOLE system

The questions of rotational maneuver and vibration stabilization of the NASA Spacecraft Control Laboratory Experiment (SCOLE) system is considered. The mathematical model of the SCOLE system includes the rigid body dynamics as well as the elastic dynamics representing transverse and torsional deformations of the elastic beam connecting the orbiter and end body (reflector). For the rotational maneuver, a new control law (orbiter control law) is derived using an orbiter input torque vector. Detumbling and reorientation maneuvers of the SCOLE system are accomplished using this control law; however, this excites the elastic modes of the beam. The orbiter control law asymptotically linearizes the flexible dynamics. Using the linearized model, a linear feedback control law is designed for vibration suppression. An observer is designed for estimating the state variables using sensor outputs which are also used for the synthesis of the control law. Simulation results are presented to show that in the closed-loop system detumbling and reorientation maneuvers can be accomplished and the effect of control and observation spillover is insignificant