欠驱动刚体航天器姿态运动规划的遗传算法

研究欠驱动刚体航天器姿态的非完整运动规划问题．航天器利用3个动量飞轮可以控制其姿态和任意定位,当其中一轮失效,航天器姿态通常表现为不可控．在系统角动量为零的情况下,系统的姿态控制问题可转化为无漂移系统的运动规划问题．基于优化控制理论,提出了求解欠驱动刚体航天器的姿态运动控制遗传算法,并且数值仿真表明：该方法对欠驱动航天器姿态运动的控制是有效的．     

带有两个动量飞轮刚体航天器的姿态非完整运动规划问题

航天器利用三个动量飞轮可以控制其姿态和任意定位.当其中一个动量飞轮失效,在某些特定的情况下,如何控制航天器的姿态问题还没有有效的方法.利用最优控制方法研究了带有两个动量飞轮的刚体航天器姿态优化控制问题.为此考虑系统角动量为零的情况下,将航天器姿态运动方程化为非完整形式约束方程,系统的控制问题可转化为无漂移系统的非完整运动规划问题.通过Ritz近似理论得到求解带有两个动量飞轮航天器姿态的运动规划控制算法.通过数值仿真,表明该方法对航天器姿态运动规划控制是有效的.     

重力梯度卫星大角度姿态机动的变结构控制

针对重力梯度稳定小卫星的大角度姿态机动问题,采用四元数来描述卫星的姿态,通过选择一类滑动流形,设计了变结构控制律,得到了在大角度姿态机动中卫星的姿态角、姿态角速度以及三个反作用飞轮转速的变化规律.理论分析和数值仿真都表明了该控制律具有渐近稳定性和鲁棒性.     

基于DSP的飞行器大角度机动控制实验技术研究

讨论了空间飞行器大角度机动控制实验平台的实现方法,并给出了硬件系统原理框图和软件的设计流程图;采用变结构控制算法,设计了基于反作用飞轮的大角度姿态机动控制器,并进行了不同角度下的闭环姿态机动控制实验,实验结果验证了该实验平台设计的可行性,对空间飞行器大角度姿态机动的研究提供了一个较好的实验平台.     

一种空间飞行器姿态控制非线性模型的预测控制新算法

空间飞行器的姿态控制受到诸如带时延的非线性动态特性、模型和参数的不确定性等因素的影响 ,其控制相当复杂。传统的控制技术 (如PID控制 )对控制对象的过程模型要求较高 ,且不能解决过程控制中非线性、时变、控制输入的约束性等因素的影响 ,其控制所能达到的性能和效率也远不够满足当前飞行器的控制要求。该文将介绍一种新型的基于控制输入的函数空间最优化的模型预测控制算法 ,称为函数空间模型预测控制 (F -MPC)。该法可用于线性和非线性系统 ,对过程模型要求不高 ,能在控制输入约束条件存在的情况下通过在线优化使系统很好地跟踪期望轨迹 ,并且解决了PID控制所遇到的问题。同时 ,将该算法用于空间飞行器的姿态控制仿真 ,仿真结果表明控制效果很好。     

Failure diagnosis of actuator using the EKF and NP theorem

There are many classical methods available for generating maneuver commands for spacecraft, where the nominal control system design consists of three torque inputs. A standard problem for the spacecraft feedback control consists of developing attitude stabilization methods that send the angular velocity to zero, if it exists. If a malfunction occurs on-orbit that eliminates or degrades the capability of torque inputs, reconfiguration of the spacecraft system is required for performing missions with the current condition. In this paper, the model-based failure diagnosis method for actuators of spacecraft, such as reaction wheels, is proposed. After generating residuals of attitude through the extended Kalman filter, failure detection of the actuator is performed. Moreover, the probability of actuator failure is conducted using the Neyman-Pearson theorem in order to recognize a degree of failure. Partial and complete failure cases are considered and successful results of failure diagnosis are presented.     

Finite-time control of spacecraft and hardware in the loop tuning

A nonlinear disturbance observer based on a super twisting controller is designed and implemented on the uncertain spacecraft attitude control subsystem simulator. The reaction wheels' angular momentum and their rate saturation are concerned in the controller design. The super twisting algorithm (STA) is devised in a way to make the reaction wheels into rest at the end of the maneuver. A nonlinear-disturbance-observer (NDO) is applied in estimating the external disturbances, unmodeled inertia moment, the eccentricity of rotation and mass center of simulator, and the reaction wheel saturation constraint. The finite-time stability of the closed-loop system is established according to the Lyapunov theory. The simulation and experimental results of this newly designed controller-observer on the spacecraft attitude simulator are compared in uncertain conditions.     

Comments on S. P. Panda “Compensator design for decoupling of multivariable systems by state feedback”

A new approach for microprocessor implementation is introduced for simultaneous control of the attitude and a number of flexure modes, assuming a lumped parameter dynamics model and control actuators mounted on a central rigid body. The switching times of the bang-bang control to be applied are calculated as a function of the current state estimate so as to render the control system performance almost independent of the microprocessor iteration interval. The control law is applicable where the modes are of arbitrarily low frequency so that the current trend of increasing size of spacecraft structures does not impose a limitation. The approach is naturally suited to gas jet actuators, but is also applicable with continuous actuators such as reaction wheels, in which case control torque saturation limits are automatically considered and the effects of bearing stiction reduced by means of the control dither produced during the limit cycle. Attention is restricted at present to control of a single axis, three-axis control with inter-axis coupling requiring a multivariable extension of the principle. At present, initial conditions must be within a certain neighbourhood of the state origin, but developments of the control principle are envisaged which will remove this restriction. Simulation results are presented of a single-axis control with up to three modes and various combinations of frequency.     

Controllability of spacecraft systems in a central gravitationalfield

The configuration space for rigid spacecraft systems in a central gravitational field can be modeled by SO(3)× IR³, where the special orthogonal group SO(3) represents the attitude dynamics and IR³ is for the orbital motion. The attitude dynamics of the spacecraft system is affected by the orbital elements through the well-known gravity-gradient torque. On the other hand, the effects of attitude-orbit coupling can also possibly be used to alter orbital motions by controlling the attitude. This controllability property is the primary issue of this paper. The control systems for spacecraft with either reaction wheels or gas jets being used as attitude controllers are proven in this study to be controllable. Rigorously establishing these results necessitates starting with the formal definitions of controllability and Poisson stability. It is then shown that if the drift vector field of the system is weakly positively Poisson stable and the Lie algebra rank condition is satisfied, controllability can be concluded. The Hamiltonian structure of the spacecraft model provides a natural route of verifying the property of weakly positive Poisson stability. Accordingly, the controllability is obtained after confirming the Lie algebra rank condition. Developing a methodology in deriving Lie brackets in the tangent space of T(SO(3)×IR³), i.e., the second tangent bundle is thus deemed necessary. A general formula is offered for the computation of Lie brackets of second tangent vector fields in TT(SO(3)^m×IRⁿ), in light of its importance in the fields of mechanics, robotics, optimal control, and nonlinear control, among others. With these tools, the controllability results can be proved. The analysis in this paper gives some insight into the attitude-orbit coupling effects and may potentially lead towards new techniques in designing controllers for large spacecraft systems     

Optimal control of large space structures governed by a coupled system of ordinary and partial differential equations

In this paper we consider the problem of optimal regulation of large space structures in the presence of flexible appendages. For simplicity of presentation, we consider a spacecraft consisting of a rigid bus and a flexible beam. The complete dynamics of the system is given by a coupled set of ordinary and partial differential equations. We show that the solution of this hybrid system is defined in a product space of appropriate finite- and infinite-dimensional spaces. We develop necessary conditions for determining the control torque and forces for optimal regulation of attitude maneuvers of the satellite along with simultaneous suppression of elastic vibrations of the flexible beam.     

Time-optimal reorientation for rigid satellite with reaction wheels

This article introduces a time-optimal reorientation manoeuvre controller with saturation constraints on both reaction wheels’ torques and angular momentum. The proposed control scheme consists of two parts. The first part is an open-loop time-minimum reorientation trajectory generated by the Legendre pseudospectral method. Actuator dynamics, saturations on control torques and angular momentums of reaction wheels are taken into account in generating the open-loop optimal trajectory. The second part is a closed-loop tracking control law to track the optimised reference trajectory based on attitude error dynamics with reaction wheel dynamics. Numerical simulations show that reaction wheel dynamics play an important role in attitude manoeuvres. The proposed controller performs better for rest-to-rest reorientation manoeuvre than other existing methods.     

Controllability of nonlinear time-varying systems: applications to spacecraft attitude control using magnetic actuation

Nonlinear controllability theory is applied to the time-varying attitude dynamics of a magnetically actuated spacecraft in a Keplerian orbit in the geomagnetic field. First, sufficient conditions for accessibility, strong accessibility and controllability of a general time-varying system are presented. These conditions involve application of Lie-algebraic rank conditions to the autonomous extended system obtained by augmenting the state of the original time-varying system by the time variable, and require the rank conditions to be checked only on the complement of a finite union of level sets of a finite number of smooth functions. At each point of each level set, it is sufficient to verify escape conditions involving Lie derivatives of the functions defining the level sets along linear combinations over smooth functions of vector fields in the accessibility algebra. These sufficient conditions are used to show that the attitude dynamics of a spacecraft actuated by three magnetic actuators and subjected to a general time-varying magnetic field are strongly accessible if the magnetic field and its time derivative are linearly independent at every instant. In addition, if the magnetic field is periodic in time, then the attitude dynamics of the spacecraft are controllable. These results are used to show that the attitude dynamics of a spacecraft actuated by three magnetic actuators in a closed Keplerian orbit in a nonrotating dipole approximation of the geomagnetic field are strongly accessible and controllable if the orbital plane does not coincide with the geomagnetic equatorial plane.     

A special case of spacecraft optimal attitude control

The problem of optimal control of a spacecraft reorientation from an arbitrary initial attitude to a prescribed angular attitude is studied. The reorientation time is given. The case when the quadratic norm of the angular velocity vector of the spacecraft is minimized is studied. Using the necessary optimality conditions in the form of Pontryagin??s maximum principle and the quaternion method for spacecraft motion control, an analytical solution of this problem is obtained. Formal equations are derived and expressions for the optimal control program are obtained. For a dynamically symmetric spacecraft, the angular velocity is obtained in the analytical form. Results of the mathematical simulation of the spacecraft dynamics under the optimal control are presented that demonstrate the practical usefulness of the proposed algorithm for the spacecraft attitude control.     

Momentum unloading excessive reaction-wheel system of a spacecraft

Wide use of electromechanical actuators in attitude control systems of spacecraft is intimately connected with improvement of these actuators using unloading methods—relieving excessive momentum. As applied to electromechanical actuators of the type of reaction wheels, the problem of unloading momentum for the case of excessive flywheel system is studied. The key feature of the work is the use of arbitrary parameters in the general solution of the undefined system of linear algebraic equations as additional control parameters. For the minimum excessive flywheel system and magnetorquers of the unloading system creating the additional external moment, control algorithms are synthesized which guarantee asymptotic stability of the zero solution to model equations describing the flywheel motion. The performance of the proposed algorithms and specific features of the process of unloading momentums of the flywheels are studied, as exemplified by the controlled motion of the spacecraft in stabilization of the regime of three-axial orbital orientation.     

Decentralised finite-time attitude synchronisation and tracking control for rigid spacecraft

The problem of finite-time attitude synchronisation and tracking for a group of rigid spacecraft nonlinear dynamics is investigated in this paper. First of all, in the presence of environmental disturbance, a novel decentralised control law is proposed to ensure that the spacecraft attitude error dynamics can converge to the sliding surface in finite time; then the final practical finite-time stability of the attitude error dynamics can be guaranteed in small regions. Furthermore, a modified finite-time control law is proposed to address the control chattering. The control law can guarantee a group of spacecraft to attain desired time-varying attitude and angular velocity while maintaining attitude synchronisation with other spacecraft in the formation. Simulation examples are provided to illustrate the feasibility of the control algorithm presented in this paper.     

挠性航天器鲁棒后步滑模姿态跟踪及主动振动控制

针对挠性航天器姿态跟踪及振动抑制问题,提出一种双回路鲁棒控制方法.首先,采用滑模控制与后步法设,计了姿态跟踪控制器,基于Lyapunov方法分析系统的渐近稳定性,并从实际应用角度考虑了反作用飞轮的动态特性;其次,为抑制挠性结构的振动,采用压电智能材料作为敏感器和作动器,设计了应变速率反馈补偿器.仿真结果表明,所提方法在保证完成姿态跟踪任务的同时,能有效抑制挠性附件的振动.     

四轴飞行器姿态控制系统设计

由于四轴飞行器系统具有不稳定、非线性、强耦合等特性,所以姿态控制在飞行器完成飞行任务的过程中尤为 重要。本文着重对飞行器姿态控制算法进行研究。首先对飞行器建立合理的坐标系,根据角度传感器所测得的角度,得到以 四元数表示的姿态转换矩阵。根据空气动力学原理,牛顿第二定律,对飞行器建立动力学模型,得到四个独立通道的控制输入 量,该控制输入量可以通过控制四轴飞行器各个方向的加速度来对飞行器进行姿态控制。     

Time-varying exponential stabilization of the position and attitudeof an underactuated autonomous underwater vehicle

The paper considers feedback stabilization of the position and attitude of an autonomous underwater vehicle (AUV) with a reduced number of actuators. A nonlinear model describing both the dynamics and the kinematics of an AUV is studied. The paper shows that previous results on attitude stabilization of a spacecraft can be applied to exponentially stabilize both the position and attitude of an AUV using only four, possibly three, actuators. Simulation results are presented     

深空探测柔性太阳帆航天器动力学建模与姿态控制

太阳帆利用太阳辐射压力提供太空航行的必要动力,由于具有理论上的无限速度和无需消耗任何燃料等优势,被认为是完成未来深空探测任务的有效技术途径之一.柔性太阳帆航天器的动力学模型包括多体动力学、刚柔耦合动力学和太阳辐射光压模型,复杂的动力学特性导致其姿态控制设计具有很强的挑战性.针对带有控制杆的柔性太阳帆航天器,本文采用拉格朗日方程和有限元法,给出了面向控制的解析式动力学模型.所推导的刚柔耦合动力学模型,刻画了太阳帆航天器的本质动力学特性,即双框架控制杆的短周期运动,姿态与柔性太阳帆的耦合效应,以及在太阳光压恢复力矩下的姿态静稳定性和长周期运动.基于带控制杆的太阳帆航天器的双时间尺度特性,提出了双回路控制结构,用于实现航天器俯仰轴和偏航轴的姿态控制.将内回路设计为PD控制器,用于实现质心位置的调整.将外回路设计为PID控制器,用于阻尼姿态运动,并实现在平衡太阳光压力矩下的姿态保持.从而将柔性太阳帆航天器的复杂姿态控制问题转化为两个低阶子问题,实现了在不同频带上的控制设计.仿真结果验证了动力学建模和姿态控制设计方法的有效性.     

Synchronization and Tracking of Multi‐Spacecraft Formation Attitude Control Using Adaptive Sliding Mode

This paper presents the finite‐time attitude synchronization and tracking control method of undirected multi‐spacecraft formation with external disturbances. First, a modified adaptive nonsingular fast terminal sliding mode surface (ANFTSMS) is designed by introducing a user‐defined function, both of which avoid the singularity problem and continuous sliding surface, and, therefore, can freely adjust relative weighting between angular velocity error and attitude error adaptively, such that the controller can provide sufficient maneuvers and precision. This provides designers with a new technique to adjust and improve formation control performance. Second, by applying the ANFTSMS associated with adaptation, two proposed decentralized ANFTSM‐controllers provide finite‐time convergence, robustness to disturbance, and chattering free for continuous design. Finally, simulation results validate the proposed algorithms.