偏置动量轮控卫星姿态控制

研究了轮控的三轴稳定的偏置动量卫星姿态控制问题。卫星的俯仰回路采用偏置动量轮，滚动／偏航轴上各安装一个反作用飞轮来完成姿态控制。小姿态角下俯仰回路可以单独设计，利用测量的俯仰角来实现其姿态控制；滚动／偏航回路利用滚动信息，采用基于偏航观测器的滑模控制器设计。磁力矩器提供的磁矩与地磁场作用产生的力矩实现了飞轮的动量卸载。对卫星姿态控制系统进行的仿真研究结果表明，所设计的控制方案在飞轮输出力矩工作范围内，可使卫星达到很高的姿态控制精度。     

基于飞轮的偏置动量卫星周期干扰补偿方法

为了提高偏置动量卫星姿态控制精度,针对三轴稳定偏置动量卫星滚动/偏航回路周期干扰力矩补偿问题,提出了一种基于反作用飞轮的滚动/偏航回路周期干扰力矩补偿方案.采用频率分离法,分析了周期干扰力矩对卫星滚动/偏航回路姿态控制精度的影响.设计了周期干扰力矩的飞轮补偿方案,该干扰力矩补偿方案由安装在卫星滚动/偏航轴上的反作用飞轮实现.仿真结果表明,所设计的飞轮补偿方案提高了滚动/偏航轴的控制精度,从而验证了飞轮补偿方案的可行性和有效性.     

利用状态观测器进行三轴稳定卫星偏航姿态误差估计的方法

1.前言为利用地球人造卫星观测地面,要对卫星进行姿态控制,使卫星本体的既定基准轴经常指向地球中心,这样的卫星叫三轴稳定卫星。三轴稳定卫星大致分为有动量飞轮的偏置动量方式和没有动量飞轮的零动量方式。偏置动量方式由于利用角动量引起的滚动/偏航转换,只用地球敏感器检测的滚动误差和俯仰误差作为卫星的姿态误差可控制姿态。另一方面,零动量姿态控制方式就需要经常地检测绕三个轴的姿态误差;滚动误差,俯仰误差和偏航误差。     

航天器姿态控制采用的动量偏置系统

自从上世纪50年代末，空军的Agena卫星采用重力梯度稳定方法以来，动量偏置系统对于航天器的姿态控制与稳定一直占据突出的地位。从那时以来，此类系统已经历了许多发展和改进，而且对于当前的地球同步轨道通信卫星来说，仍是占据主导地位的稳定方法。本文跟踪了Space／Systems Loral公司建造的航天器上所采用的动量偏置系统的一个特别的发展过程，重点讨论三类滚动／偏航控制系统。第一类是在具有动量偏置的早期卫星上使用的WHECON系统，第二类是在80年代中至90年代初的卫星主要采用了Terasaki控制系统，第三类是高斯线性二次型最优(LQG)动量偏置控制系统，该类系统90年代初引入，并在当前飞行的许多卫星上所采用。     

四旋翼无人机的偏航抗饱和与多模式PID控制

四旋翼无人机偏航运动能力比俯仰、滚转运动能力弱,因此在偏航运动上更容易出现执行器饱和.特别在外界干扰下,偏航常常出现饱和现象,对此从实际工程出发提出了一种多模式PID(MMPID)控制器,抑制偏航饱和,并保证在外界干扰情况下具有较好的偏航控制性能.首先,根据四旋翼无人机的动力学模型,设计了基于MMPID的偏航控制器.MMPID算法具有多模式特性,在满足一定条件时采取退饱和控制策略.其次,利用李亚普诺夫稳定理论证明了基于MMPID的偏航控制系统的稳定性.最后,通过四旋翼无人机偏航仿真实验比较了MMPID与变结构PID算法的抗饱和性能与偏航控制性能,结果表明MMPID算法具有明显的控制性能优越性.并且,通过四旋翼原型机实验验证了MMPID偏航控制器的有效性和鲁棒性.实验结果表明:在室内无干扰情况下,基于MMPID的偏航系统基本无执行器饱和现象,最终偏航角误差收敛到±0.05 rad;在室内加入干扰情况下,执行器退出饱和时间约10 s,偏航角误差收敛到±0.08 rad;在室外,偏航退饱和时间约5 s,偏航误差收敛到±0.13 rad.MMPID控制器有效地抑制了四旋翼无人机偏航运动上的执行器饱和,并可实现准确的偏航姿态控制和强鲁棒性.     

基于模糊树逆方法的高超飞行器变质心控制

本文针对高超飞行器变质心姿态动力学模型所具有的高度非线性、多变量、强耦合特性,通过合理的简化,得到了一个耦合的非线性动力学系统;根据简化模型的特点,将该模型分为俯仰偏航通道模型和滚动通道模型,通道间耦合视为扰动.应用模糊树模型逼近俯仰偏航通道的逆模型,实现了俯仰偏航通道解耦线性化,对各线性子系统设计闭环控制器,实现了飞行器俯仰偏航通道姿态角的跟踪;应用预测变结构方法保证了滚动通道的快速稳定并且无抖振现象.最后,将该方法应用于高超飞行器姿态系统控制的仿真研究,表明了本文方法的有效性和可行性.     

基于磁传感器的飞行器姿态测量方法

为近程自旋飞行器设计了一种基于地磁测量的姿态测量方法。根据此类飞行器飞行过程中偏航角较小,射程较近时偏航角可以忽略的特点,在偏航角可以忽略的情况下进行姿态角算法分析。由于第一种方法存在姿态角根取舍的问题,首先,利用空间转换矩阵推导出飞行器姿态角的表达式。然后,利用几何分析的方法推导出了飞行器姿态角表达式,并对该算法的误差进行仿真。仿真结果表明:误差在可接受范围。     

基于时间交互偏置影响传播模型的弱连接重叠社区检测

为提高弱连接重叠社区的检测识别性能,提出一种基于时间交互偏置影响传播模型的弱连接重叠社区检测算法。设计针对社区检测图模型分割的目标函数,利用群落结构对处理器负载均衡进行优化,以提高模型求解的效率。基于邻域边缘密度对近似活跃边缘进行重新定义,构建一种影响传播模型以确定用户具有高频率的相互作用,从而提高弱连接用户的识别性能。在此基础上,提出时间交互偏置社区检测方法。实验结果表明,该方法对重叠社区进行检测时具有较高的识别精度和效率。     

矿用全断面硬岩快速掘进机位姿纠偏控制系统设计

介绍矿用全断面硬岩快速掘进机及其姿态测量与控制方法.结合盾构机和悬挂掘进机姿态测量方法,提出基于激光指示仪和新型电子激光标靶的测量系统,对掘进机进行实时三维坐标测量.通过采集的俯仰角、滚动角以及水平偏航角,来推导出掘进机切口、尾部中心的坐标以及位置偏差,经过误差仿真实验表明在本测量系统下,测量精度符合施工要求.PLC根据掘进机水平和垂直偏差量对掘进机进行实时纠偏,并给出了纠偏程序实现流程和友好的人机界面设计,实现了对全断面硬岩快速掘进机姿态的准确测量与稳定控制.     

基于状态空间MPC的无人机无人车联合运动控制

针对四旋翼无人机无人车联合运动缺乏对系统成员姿态约束的问题,研究了一种基于模型预测控制(MPC)的分布式联合运动控制方法.基于虚拟结构法,使用虚拟领航者策略,以虚拟领航者提供参考轨迹及参考速度,分别在各无人器平台上转换成各自所需的预测时域信息,结合推导的各无人器的状态空间模型滚动优化实现预测控制.限定四旋翼高度方向运动状态与偏航角,构造以俯仰角、横滚角与重力加速度乘积为位置运动输入的状态空间模型,将无人机内环姿态控制约束加入位置运动,增强飞行稳定性.改良无人车状态空间模型,增加速度信息得到可提供位置速度追踪的增广状态空间模型,增强运动追踪能力.仿真表明在满足无人器姿态约束条件下,能够保证联合运动的位置速度精度.     

Arm path planning of a space robot with angular momentum

Arm path planning of a space robot with angular momentum is considered in this paper. A space robot changes its attitude by the arm motion and angular momentum of the space robot has the possibility to reduce the attitude change. A path planning method of the arm where the final satellite attitude becomes the same as the initial one is proposed. The method derives an approximate path first based on the attitude change when the arm moves along an infinitely small closed curve and then modifies the path by the Newton method. The amplitude of the arm motion decreases with the magnitude of the angular momentum, which shows that the proposed method utilizes the angular momentum effectively.     

Modification of the exact pole placement method and its application for the control of spacecraft motion

For linear stationary models of the spacecraft motion, a modification of the exact pole placement method is developed that makes it possible to design a modal control for any class of linear stationary systems. In the case of circular orbits, the modified method is used to obtain an analytical solution of the problem of gravitational angular momentum unload problem for inertial actuators of a spacecraft in the gyroscopically coupled roll-yaw channels.     

欠驱动航天器双飞轮—单喷气姿态最优控制原理及方法

欠驱动航天器的姿态控制能够增强航天器的可靠性.本文针对欠驱动航天器姿态控制, 从喷气姿态阻尼的角动量等效原理出发, 推导脉宽调制公式, 得到燃料消耗最小时给定姿态、非给定姿态两种情况下的喷气最优组合方案.同时, 为了实现喷气全局最优, 提出欠驱动飞轮姿态控制策略, 实现了运动航天器机动至预期姿态.进一步分析欠驱动飞轮航天器的姿态控制原理及稳定性, 提出了共面双飞轮-单喷气的配置方案, 通过双飞轮组合稳定航天器的角速度, 使得航天器到达预期姿态机动时燃料全局最省.结合绕两个旋转轴的姿态机动路径规划方法, 通过姿态机动时序关系的实时分配可实现航天器姿态机动与稳定控制.最后, 通过航天器姿态控制仿真和对比分析, 发现共面双飞轮-单喷气的欠驱动姿态阻尼及姿轨控制方案能够在较少硬件配置下实现对航天器的姿态控制, 且消耗燃料最少.     

Distributed observer‐based leader‐follower attitude consensus control for multiple rigid bodies using rotation matrices

This paper addresses the distributed observer‐based leader‐follower attitude consensus control problem for multiple rigid bodies. An intrinsic distributed observer is proposed for each follower to estimate the leader's trajectory, which is only available to a subset of followers. The proposed observer can guarantee that the estimated attitude evolves on rotation matrices all the time, and it provides us with a simple way to design the attitude consensus control law. The dynamics of rigid bodies and distributed observer are both modeled directly on rotation matrices, so that the singularity and ambiguity can be avoided. Furthermore, adopting the idea of disturbance observer on vector space, a gyro bias observer on the rotation matrices is proposed. Based on the distributed observer, three types of attitude consensus control law are proposed, which are respectively on the basis of full‐state, biased angular velocity, and external disturbance combined with biased angular velocity. Finally, the SimMechanics experiments are provided to illustrate effectiveness of the proposed theoretical results.     

Three-dimensional locomotion control of single-legged robot: Resonance hopping and running direction

This study proposes a locomotion controller for a single-legged robot. The locomotion controller comprises five parts: virtual spring, height control, forward velocity control, body attitude control, and angular momentum control. First, we propose an effective method called a virtual spring to generate a springy force using a linear actuator. Two virtual springs are adopted to compute the energy exchange and to compensate the energy loss during hopping. A simple and intuitive method is proposed to implement resonance oscillation in terms of energy loss. A height controller is proposed based on the resonance oscillation using a virtual spring. In addition, a running direction controller, which has never been resolved in previous studies, is proposed. This novel controller can remove the unexpected angular momentum about the yaw direction during running. All of the proposed algorithms and methods are validated through dynamic simulations.     

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.     

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．     

A Singularity-free Steering Law of Roof Array of Control Moment Gyros for Agile Spacecraft Maneuver

In this paper, a singularity-free steering law for single gimbal control moment gyros (CMGs) is addressed for agile spacecraft. The geometrical array considered particularly in this work is a roof array due to the simplicity of singularity envelope. A feasible angular momentum chart which can provide a singularity-free bound is employed. The chart allows a guaranteed maximum torque output and angular momentum at any time without concerning the geometrical singularity of the array. Furthermore, a new deterministic allocation algorithm, called half-leading steering logic, of gimbal angular rates, is also suggested instead of the well-known pseudo-inverse technique to meet control torque commands required and to keep away from the singularity. It is noted that a momentum vector recovery to the initial direction is also an important task for the CMG array to overcome the singularity and for the reliable operation of CMGs. An optimization technique is addressed to restore the gimbal vectors back to their original angular position after the attitude reorientation mission. The techniques proposed are demonstrated using illustrative numerical simulations.

     

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.     

On the stabilization of a communication satellite without measuring its angular velocity

The motion control about the center of mass of the communication satellite Yamal-200 is considered. Due to possible failure of the angular velocity sensors, an algorithm for maintaining a prescribed attitude without measuring the angular velocity was developed for this satellite. The algorithm ensures the gyroscopic stabilization of the satellite motion about the local vertical. In the course of control, the onboard computer stores projections of the flywheels?? angular momentum vector on the axes of the inertial basis; the control system tries to preserve the angular momentum magnitude and direction in the inertial space based on the measurements of the angular momentum. The workability of the proposed stabilization algorithm was confirmed by flight tests.