一种偏置动量卫星的三轴姿态控制方法

提出了一种滚动一偏航平面两自由度偏置动量控制方法。在分析某通信卫星平台所受的环境干扰力矩对姿态扰动的基础上,依据Y轴偏置动量大小的影响、构型的角动量包络、动量轮最大角动量变化值和可靠性等参数,与现有的四轮金字塔构型两自由度偏置动量控制进行了比较,其性能更优越。     

Jitter reduction of a reaction wheel by management of angular momentum using magnetic torquers in nano- and micro-satellites

Nowadays, nano- and micro-satellites, which are smaller than conventional large satellites, provide access to space to many satellite developers, and they are attracting interest as an application of space development because development is possible over shorter time period at a lower cost. In most of these nano- and micro-satellite missions, the satellites generally must meet strict attitude requirements for obtaining scientific data under strict constraints of power consumption, space, and weight. In many satellite missions, the jitter of a reaction wheel degrades the performance of the mission detectors and attitude sensors; therefore, jitter should be controlled or isolated to reduce its effect on sensor devices. In conventional standard-sized satellites, tip-tilt mirrors (TTMs) and isolators are used for controlling or isolating the vibrations from reaction wheels; however, it is difficult to use these devices for nano- and micro-satellite missions under the strict power, space, and mass constraints. In this research, the jitter of reaction wheels is reduced by using accurate sensors, small reaction wheels, and slow rotation frequency reaction wheel instead of TTMs and isolators. The objective of a reaction wheel in many satellite missions is the management of the satellite’s angular momentum, which increases because of attitude disturbances. If the magnitude of the disturbance is reduced in orbit or on the ground, the magnitude of the angular momentum that the reaction wheels gain from attitude disturbances in orbit becomes smaller; therefore, satellites can stabilize their attitude using only smaller reaction wheels or slow rotation speed, which cause relatively smaller vibration. In nano- and micro-satellite missions, the dominant attitude disturbance is a magnetic torque, which can be cancelled by using magnetic actuators. With the magnetic compensation, the satellite reduces the angular momentum that the reaction wheels gain, and therefore, satellites do not require large reaction wheels and higher rotation speed, which cause jitter. As a result, the satellite can reduce the effect of jitter without using conventional isolators and TTMs. Hence, the satellites can achieve precise attitude control under low power, space, and mass constraints using this proposed method. Through the example of an astronomical observation mission using nano- and micro-satellites, it is demonstrated that the jitter reduction using small reaction wheels is feasible in nano- and micro-satellites.     

Hardware-In-the-Loop Simulations of spacecraft attitude synchronization using the State-Dependent Riccati Equation technique

A nonlinear control technique pertaining to attitude synchronization problems is presented for formation flying spacecraft by utilizing the State-Dependent Riccati Equation (SDRE) technique. An attitude controller consisting of relative control and absolute control is designed using a reaction wheel assembly for regulator and tracking problems. To achieve effective relative control, the selective state-dependent connectivity is also adopted. The global asymptotic stability of the controller is confirmed using the Lyapunov theorem and is verified by Monte-Carlo simulations. An air-bearing-based Hardware-In-the-Loop Simulator (HILS) is also developed to validate the proposed control laws in real-time environments. The SDRE controller is discretized for implementation of a real-time processor in the HILS. The pointing errors are about 0.2° in the numerical simulations and about 1° in the HILS simulations, and experimental simulations confirm the effectiveness of the control algorithm for attitude synchronization in a spacecraft formation flying mission. Consequently, experiments using the HILS in a real-time environment can appropriately perform spacecraft attitude synchronization algorithms for formation flying spacecraft.     

Adaptive sliding mode fault-tolerant control for satellite attitude tracking system

In this paper, an adaptive modified sliding mode control approach is developed for attitude tracking of a nano-satellite with three magnetorquers and one reaction wheel. A sliding variable is chosen based on finite-time convergence of the nano-satellite attitude tracking error and avoiding the singularity of the control signal. The control gain of the proposed method is developed adaptively to reduce the tracking error and improve the closed-loop control performance. The sliding variable and adaptive parameter are also employed in the reaching phase of the control law to decrease the chattering phenomenon. In addition, the finite-time convergence of attitude variables in the presence of actuator faults, inertia uncertainty, and external disturbances is proved using the extended Lyapunov theorem. The simulations are conducted to evaluate the performance of the proposed method according to different evaluation criteria. Monte Carlo simulations are also used to survey the reliability of the system in the presence of the mentioned condition.     

灵敏小卫星能量/姿态一体化控制研究

研究应用变速控制力矩陀螺群(VSCMGs, Variable Speed Control Moment Gyros)作为主执行机构,完成灵敏小卫星多目标快速姿态机动时的能量/姿态一体化控制问题.在考虑执行机构饱和、机动任务要求和敏感器测量与跟踪能力受限等情况下,设计了非线性的姿态/能量一体化控制器,对由于频繁的姿态机动引起的姿态四元数的漂移进行了整定.设计了VSCMGs的操纵律.按是否接近框架构型奇异,合理分配了操纵任务,并设计了相应的操纵方法.对采用金字塔构型的VSCMGs进行了较为严格的仿真,结果表明卫星在机动中达到了快速和稳定的要求,同时能够满足能量控制要求和VSCMGs转子转速的平衡.     

基于退步控制的姿态跟踪控制律的设计

退步控制是一种有效设计级联系统控制律的方法,而基于单位四元数的刚体姿态方程就是一种典型的级联系统。文章针对恒值干扰力矩作用下的刚体姿态跟踪控制问题,基于退步控制方法设计出一种含积分项的刚体姿态控制律,控制律中引入积分项是为了消除恒值干扰力矩产生的恒值偏差;并通过在李亚普诺夫函数中引入交叉项证明了控制律的全局稳定性;最后通过数值仿真验证了控制律。     

零动量轮卫星姿态控制系统研究

针对零动量轮卫星控制系统解耦和进动抑制问题,提出了一种新的进动抑制方案。分析了仅靠飞轮控制的零动量轮卫星系统的稳定性。将偏置动量轮卫星的进动和章动抑制原理引入零动量控制系统,并针对零动量轮卫星自身特性,对磁控制规律进行改进,解决了不解耦情况下零动量轮卫星的进动抑制问题。仿真结果验证了所提方案的可行性和有效性。     

火星环绕器环火轨道的角动量管理方法

针对中国首次火星探测的任务需求和规划，结合环绕器质量特性和布局构型以及大椭圆轨道特性，分析了环绕器在环火飞行阶段所受的重力梯度力矩和光压力矩规律.空间干扰力矩累积导致飞轮转速上升，影响姿态机动能力，需要喷气卸载.为减少喷气卸载次数，在满足系统任务的基础上，设计了姿态偏置方案，利用光压力矩和重力梯度力矩相互抵消减少角动量累积.通过仿真表明该方法可以将空间干扰力矩累积量减少70%，节约卸载喷气的燃料消耗，延长航天器寿命.     

基于动量轮的航天器全姿态捕获技术

全姿态捕获是指航天器丢失姿态基准而需恢复正常姿态或者定向新姿态目标时的一种控制过程.以往卫星全姿态捕获控制过程一般采用喷气推进系统或磁力矩器作为执行机构,这些方法需要消耗燃料或捕获时间长.提出了一种基于动量轮的全姿态捕获方法,采用的部件为卫星的常规配置,可以实现对任意目标的定向,该方法克服了以往方法的缺陷,在轨验证结果表明该方法有效、工程可操作性强.     

改善航天器反作用轮扰动实验模型参数的辨识方法

反作用轮系统是影响航天器姿控系统精度的主要扰动源之一.建立反作用轮扰动模型的目的是预测扰动对航天器产生的影响,并采取相应的控制方法和隔离系统.基于反作用轮的扰动实验模型,通过对反作用轮扰动实验数据的分析,确定出反作用轮扰动实验模型中的参数:谐波数和幅值系数,并在此基础上提出了能量补偿法,最后进行了数值仿真.结果表明,谐波数的辨识精度不超过0.04%,当采用振幅谱法计算幅值系数时,误差高达15.5%;而用能量补偿法,其幅值系数的精度不超过1.1%.可见能量补偿法提高了幅值系数的辨识精度.本文研究为改善航天器姿态控制精度和稳定度奠定了一定的基础.      

“风云一号” B 卫星姿态控制系统

“风云一号”气象卫星姿态控制系统采用了三轴稳定对地定向的主动控制方案。已发射了两颗（Ａ、Ｂ）卫星,其中Ｂ星是在Ａ星主控系统的基础上,增加了一个完整的备份系统,采取一系列冗余措施并设计全方位姿态重新捕获的故障对策。经飞行试验及在轨故障应急处理证明,系统设计是完善和成功的,其中反作用飞轮控制、偏置动量控制、磁章动进动和飞轮卸载控制、全方位姿态重新捕获方案在中国是首次采用,均取得了较好的飞行效果,为长寿命卫星姿态控制系统的设计积累了宝贵经验。     

万向磁悬浮动量轮研究

目前大部分空间飞行器三轴姿态控制采用的是飞轮系统 ,其精度较高 ,而采用磁悬浮轴承支承的动量轮比普通的滚珠轴承动量轮有更高的姿控性能。在此基础上发展的动量矩方向可改变的万向磁悬浮动量轮可以单独实现三轴较高精度的主动姿态控制 ,具有重要的实用价值。同时 ,万向磁悬浮动量轮的研究也为集成能源与姿态控制系统的研究提供了技术支持。文意介绍了万向磁悬浮动量轮在国际上研究和发展的情况 ,概括了其中的关键技术 ,并给出了一种初步的结构设计实例。     

Landing of hopping rovers on Irregularly-shaped small bodies using attitude control

A hopping rover that is driven only by internal or external attitude actuators is an ideal mobility approach for surface exploration of small solar system bodies. Without thrust control and grasping mechanisms, a hopping rover is mechanically simple to design and less prone to mechanical failures, but faces challenges during soft landing. It may rebound from the surface, causing deviations from its original landing site. In this paper, landing of a hopping rover on the surface of an asteroid is investigated, and a strategy using only attitude control to shorten the landing distance is proposed. Based on rigid body impact dynamics, the edge impact configuration is investigated in detail. The factors that affect the impact states of a cube-shaped hopping rover are studied. Then, controlled edge landing is analyzed, in which the post-impact velocity of the hopping rover is changed by controlling its attitude prior to impact. Three guidance schemes are developed, followed by attitude profile generation and finite time stable attitude control. Finally, simulations are performed on an ideal flat surface and uneven terrain. The results show that controlled edge landing can effectively reduce the landing distance and settling time, compared with uncontrolled landing. This study on hopping motion on the surface of an irregular-shaped asteroid with attitude control, can provide a reference for hopper trajectory plan in future asteroid surface explorations.     

CubeSats detumbling using only embedded asymmetric magnetorquers

After deployment from a rocket, a CubeSat is detumbled using magnetorquer rods bringing the norm to the point where the reaction wheels take over to reduce the angular velocity to null. Therefore, utilizing reaction wheels for satellite detumbling at higher initial velocities is vital but they are heavy and occupy significant space on a spacecraft having challenging control. To address this challenge, this paper features a disruptive approach that conducts the control only by the PCB-integrated magnetorquers with various geometries using a diverse non-unity track width ratio. The trace widths are parametrized such that the optimal torque to power dissipation ratio is investigated. The optimizations are then simulated for various geometric distributions and validated through comprehensive measurement setups that establish a framework for selecting the best-case coil configuration according to mission requirements. The detumbling rates of multiple asymmetric coil configurations are compared with the embedded designs in published literature and state of the art. It is found that the proposed asymmetric embedded magnetorquers can detumble the vehicle at high initial angular velocities. Lastly, the simulation results of thermal analysis are validated for selecting the application-specific optimal coils configuration. At the end, the proposed system is compared with the embedded magnetorquers available in the literature and commercial attitude control systems.     

基于电磁力与力矩的航天器运动控制的研究进展

电磁力与力矩是一种非接触式的相互作用力与力矩.将通电线圈或磁铁放在磁场中会受到电磁力与力矩的作用.多个通电线圈或磁铁相互作用时，在彼此构成的磁场中也会受到电磁力与力矩的作用.作为一种非接触的力与力矩，电磁力与力矩在航天器的姿态与相对位置控制系统中已有了广泛和深入的应用.本文主要介绍近年来由航天器上电磁装置受到地磁场作用产生电磁力矩在姿态控制中的应用，和带有电磁机构的多航天器间相互作用控制多航天器间相对位姿的应用等.     

多星近距离绕飞观测任务姿轨耦合控制研究

针对多星近距离绕飞观测任务,建立了相对姿态轨道动力学模型,分别考虑了在椭圆、空间圆绕飞轨道上观测卫星的两种期望三角形编队构型,以观测卫星视线始终指向目标为期望姿态,采用基于四元数和角速度误差反馈的比例 微分控制律以及一种改进的基于人工势场法的制导方法相结合,对相对姿态及轨道进行控制。仿真结果表明：在控制律的作用下,绕飞过程中各观测卫星均能够有效地跟踪期望相对姿态和期望相对轨道;在空间圆绕飞轨道构型中,各观测卫星从初始同一位置出发后,在任意时刻3颗观测卫星构成的编队构型始终为正三角形,且正三角形的边长从零逐渐增大,最终等于期望正三角形构型的边长。     

一种抑制反作用轮低速摩擦对卫星姿态扰动的方法

在现代卫星的姿态控制系统中,反作用轮得到了广泛应用。但是当反作用轮的转速过零时,摩擦力矩会对卫星的姿态产生较大影响。本文采用基于特征模型的黄金分割自适应控制方法,建立了包括反作用轮在内的卫星系统的特征模型,并由此设计了控制律。仿真结果表明,该方法可以有效抑制反作用轮低速摩擦对卫星姿态的扰动,从而可以提高卫星姿态控制精度。     

基于周期变化惯量积的主动章动控制方法研究

偏置动量卫星姿态控制过程中会产生章动运动。若星体俯仰轴与其它轴存在固定的惯量积,则可以利用偏置动量轮有效地克服章动。针对转动惯量周期性变化的偏置动量卫星,提出一种通过在动量轮的控制律中引入时变的滚动轴信息,利用动量轮克服滚动轴／偏航轴章动影响的方法。通过数学仿真验证了该算法的有效性。     

动量轮微振动机理及仿真

动量轮是卫星等航天器姿态控制和精度保持的关键机械部件,其微振动严重影响卫星姿态稳定度和成像精度。动量轮的非均匀、非连续几何构形和旋转效应会引起结构系统的参数激励和载荷激励。针对具有非均匀力学特征参数的动量轮结构系统动力学模型,通过分析动力学方程中各矩阵参数的扰动,进行动量轮微振动机理的研究。仿真和试验结果表明:动量轮结构系统内部存在基频和高频激励,其中基频主要来自支点动载荷,高频来自轴承碾压作用;轮缘的局部振动会随转速形成前后行波。      

零动量卫星轮控系统可重构性设计研究

针对卫星姿态控制系统设计中常见的动量轮配置问题,总结了星上常用的5种轮控系统构型,首先对各种构型从角动量包络、干扰力矩下动量轮的饱和情况、系统功耗和可靠性等方面进行了可重构性分析和比较,在此基础上提炼了动量轮可重构性设计准则,可为轮控系统的设计提供参考依据。