ATTITUDE CONTROL TECHNOLOGY FOR MASS MOMENT NANO-SATELLITE IN LOW EARTH ORBIT 1)

针对气动力矩严重影响低轨纳卫星姿态控制效果的问题,创新性地提出了利用质量矩技术将气动干扰转化为控制力矩的解决方法.由于气动力矩矢量垂直于大气来流速度方向,因而采用质量矩与磁力矩相结合的方式三轴全驱动控制卫星姿态,从而避免系统欠驱动. 建立双执行机构控制方式的姿态动力学模型,并根据各干扰项的影响简化了控制方程.针对气动力不确定、星体参数误差、未知环境影响等复杂干扰,设计了针对理想控制力矩基于干扰观测器的滑模控制器. 为减小滑块附加干扰力矩,研究了理想控制力矩的最优分配策略. 最后, 为双执行机构搭建了半物理仿真平台,结果表明: 姿态机动过程中, 与滑块加速度相关的附加惯性力矩远大于其他干扰项,最优力矩分配策略能够大幅减小快时变的附加干扰, 优化效果明显; 姿态保持过程中,干扰观测器能有效观测系统慢时变干扰, 提高滑模控制律的姿态控制精度,姿态角收敛误差小于$\pm $0.1$^\circ$.最终验证了在低轨纳卫星上利用质量矩技术控制姿态的可行性.     

基于MSCMG大型遥感卫星高精度姿态控制方法

针对大力矩飞轮前馈和闭环反馈补偿复杂、对精度影响敏感性大的问题,提出了基于磁浮控制力矩陀螺闭环补偿的大型遥感卫星高精度姿态控制方法。该方法采用磁悬浮力矩陀螺为控制执行机构,通过变结构反馈补偿控制律设计,建立新的运动补偿控制系统,减小相机和卫星本体耦合效应。基于磁浮力矩陀螺力矩大、反向激励扰动小、精度高的特性,将其应用于对地遥感成像相机运动补偿控制系统中,仿真结果表明,与飞轮前馈补偿相比,姿态稳定度提高了一个数量级,有效提高空间大惯量卫星姿态控制的稳定度,提升相机对地成像质量;研究结果可为甚高精度卫星姿态控制与载荷运动补偿提供参考。     

飞网抛射过程母卫星姿态干扰分析与姿态控制

空间飞网是一种新型的空间碎片或漂浮物回收装置. 飞网抛射过程会对母体卫星产生较大的干扰,影响卫星的姿态稳定,因此需要设计具有抗干扰能力的控制律来保证卫星姿态的稳定. 针对空间飞网直接抛射展开过程,分析了影响飞网抛射效果的主要因素,并建立了飞网发射所产生的干扰力矩模型. 针对卫星的控制要求,建立了卫星姿态控制系统的模型,设计了滑模变结构控制器,并给出数学推导过程. 最后,通过仿真对姿态干扰及控制器控制性能进行了分析研究,结果表明,设计的控制器能够保证卫星的稳定,满足设计指标要求.     

立方星剩磁在轨辨识与主动补偿技术

立方星的姿态测量与控制系统常采用磁测磁控结合偏置动量轮的方案,整星剩磁干扰力矩是影响姿态控制精度的重要因素之一。提出了一种利用磁强计实现剩磁矩在轨辨识与利用磁力矩器实现剩磁矩主动补偿的新方案:基于磁强计输出和卫星姿态动力学建立了剩磁矩在轨辨识模型,并利用采样滤波器(UKF)提高单磁强计条件下的辨识效果;把控制对象简化成线性定常系统,分析了剩磁干扰力矩对姿态的影响数学模型,并针对磁力矩器和磁强计分时工作的特点,基于叠加性原理提出了基于角速度的剩磁矩主动补偿算法。仿真研究表明,在1000 s内剩磁矩在轨辨识精度为0.001 A×m~2量级,主动补偿后,偏航角、滚动角与俯仰角控制误差分别从4.3°、4.6°与2.1°均减少至0.4°以内。提出的方法为类似配置卫星减少剩磁干扰力矩的影响提供了一种新思路。     

微小卫星联合执行机构的递阶饱和姿态控制方法

随着航天飞行任务复杂程度的提高,微小卫星姿态控制系统对实现大角度姿态机动的快速性及稳定性有着较高的要求。考虑到在实际大角度姿态机动过程中力矩饱和及角速度限制的因素,提出了基于欧拉轴转动的递阶饱和模糊PD姿态控制律,同时采用喷气推力器和反作用飞轮作为联合执行机构为微小卫星姿态机动提供大且精确的控制力矩。与传统PD控制律相比,模糊PD姿态控制系统的未知参数可在线自动整定。数学仿真结果表明,基于欧拉轴转动的递阶饱和模糊PD姿态控制系统能够在125 s内实现50.2°的大角度机动,稳定误差能够控制在0.002°之内。与传统PD控制律相比,此方法具有更高的精度及稳定性。     

三自由度气浮台自动平衡系统动力学建模

为了在地面模拟卫星的姿态运动,三自由度气浮台的回转中心应与重心重合,并且各轴与对应卫星各轴应具有相等的转动惯量,实现1∶1模拟,这样执行机构的控制力矩矢量与实际卫星相同;或两者的角加速度相一致,以模拟卫星在外层空间所受干扰力矩很小的力学环境.当重心相对于回转中心存在漂移时产生静态不平衡;主惯性轴相对于回转轴存在漂移时会产生动态不平衡.该文阐述了三自由度气浮台的平衡步骤以及手动平衡和自动平衡技术,推导了三自由度气浮台自动平衡动力学公式,并对平衡过程进行了数值仿真,仿真结果表明自动平衡调节可以使平台浮起部分的摆动周期延长至约1300 s.     

基于迭代学习干扰观测器的 RLV 容错控制方法

针对重复使用运载器（RLV）等类飞行器再入飞行段存在综合干扰问题,提出了一种基于迭代学习干扰观测器的容错控制方法。首先,根据RLV再入段运动、动力学特性及执行机构故障类型,建立带有执行机构故障的RLV面向控制模型;然后设计了一种基于S型函数的迭代学习观测器的容错控制方法,采用迭代学习干扰观测器完成对综合干扰观测并进行补偿,通过李雅普诺夫定理证明基于干扰观测器设计的改进的自适应控制器能够在有限时间收敛稳定;最后以某型RLV再入段为研究对象进行数值仿真。仿真结果表明,所提出的方法在系统中存在加性、乘性故障时,姿态跟踪能够在3 s内收敛,同时姿态稳态误差在0.01°以内。     

基于扰动观测器的AUV固定时间积分滑模控制方法

针对自主水下航行器（AUV）在参数不确定性和外界干扰下水平轨迹跟踪控制问题，提出一种基于扰动观测器的固定时间积分滑模控制方法。首先将参数不确定性和外界干扰视为复合扰动，设计固定时间扰动观测器对其进行估计。然后在反步法设计框架下，结合固定时间理论和全局积分滑模控制，设计了固定时间积分滑模控制器。轨迹跟踪仿真结果表明，相比于传统滑模控制器，所设计的扰动观测器和控制器可以使位置和姿态的跟踪误差收敛至零域的速度由5.2 s缩短至约2.5 s，并且在没有观测器的帮助下跟踪误差也能收敛至稳定，具有更快的收敛速度和更高的鲁棒性。     

基于自适应滑模的重复使用运载器容错控制

针对重复使用运载器(RLV)等类飞行器存在外界干扰和执行机构故障等情况,提出一种基于姿态跟踪容错控制方法。在正常的运行模式下,姿态跟踪控制采用连续四元数反馈控制器。当系统中出现故障时,飞行器姿态将偏离参考轨迹,此时触发控制系统中滑动模态反应,使系统具有鲁棒性。通过选取适当李雅普诺夫函数,证明了所提出的控制律在存在故障的情况下是渐近稳定的。针对由于传感器干扰滑模面非零而导致的增益渐增,以及控制器性能下降问题,设计了一种具有自适应参数的自适应滑模控制律,使增益能够收敛到合理上界。最后,选取重复使用运载器再入段为对象进行仿真验证。仿真结果表明,采用有自适应滑模参数的控制系统,四元数跟踪误差能够达到10~(-4)量级。     

单框架控制力矩陀螺输出特性分析

广泛用于航天领域的单框架控制力矩陀螺, 具有力矩放大效应的优点,其理论基础为有假设条件的力矩放大原理. 本文不局限于这些假设, 不限定工况,解析单框架控制力矩陀螺的输出特性. 考虑安装基座的运动,得到具有两维输入三维输出的单框架控制力矩陀螺力矩输出模型,提出将输出力矩分解为可调控与不可调控两部分. 为分析单框架控制力矩陀螺的输出特性,定义两个参数, 分别为输出输入力矩比和输出力矩利用率. 研究发现,单框架控制力矩陀螺不恒有力矩放大效应, 也不恒有高效的力矩利用率,两者与其状态密切相关. 最后,以含两个单框架控制力矩陀螺的航天器姿态机动任务为例,对非对角奇异鲁棒操纵控制和优化控制进行仿真,检验了单框架控制力矩陀螺输出特性对控制效果的影响. 同时,根据单框架控制力矩陀螺的三维输出特性, 借助一个单框架控制力矩陀螺的优化控制,实现了航天器的三轴姿态机动. 仿真结果显示, 在优化控制过程中,单框架控制力矩陀螺始终具有力矩放大效应和高效的力矩利用率.      

STOCHASTIC VIBRATION CONTROL BASED ON PIEZOELECTRIC STACK INERTIAL ACTUATOR1)

本文主要研究了基于压电叠堆惯性作动器的随机振动控制方法并分析了控制力时滞对控制效果的影响规律。首先,对压电叠堆惯性作动器的动力学特性进行分析,建立了考虑惯性作动器动力学的主结构的随机动力学模型;接着,提出主结构随机振动 "线性"二次高斯控制(linear quadratic Gaussian,LQG)方法,并探讨了作动器惯性质量和刚度等对控制效果的影响规律;最后,研究了控制力时滞对控制效果的影响,并提出了利用时滞优化控制效果的方法。研究结果表明,压电叠堆惯性作动器对结构的随机振动具有较好的控制效果,当控制时滞时间$\tau = 0.025$s时控制效果可以得到进一步优化。     

Sliding mode maneuvering control and active vibration damping of three-axis stabilized flexible spacecraft with actuator dynamics

This paper presents a dual-stage control system design method for the three-axis-rotational maneuver control and vibration stabilization of a spacecraft with flexible appendages embedded with piezoceramics as sensor and actuator. In this design approach, the attitude control system and vibration suppression were designed separately using a lower order model. Based on the sliding mode control (SMC) theory, a discontinuous attitude control law in the form of the input voltage of the reaction wheel is derived to control the orientation of the spacecraft actuated by the reaction wheel, in which the reaction wheel dynamics is also considered from the real applications point of view. The asymptotic stability is shown using Lyapunov analysis. Furthermore, an adaptive version of the proposed attitude control law is also designed for adapting the unknown upper bounds of the lumped disturbance so that the limitation of knowing the bound of the disturbance in advance is released. In addition, the concept of varying the width of boundary layer instead of a fixed one is also employed to eliminate the chattering and improve the pointing precision as well. For actively suppressing the induced vibration, modal velocity feedback and strain rate feedback control methods are presented and compared by using piezoelectric materials as additional sensors and actuators bonded on the surface of the flexible appendages. Numerical simulations are performed to show that rotational maneuver and vibration suppression are accomplished in spite of the presence of disturbance torque and parameter uncertainty.     

基于分布式合成双射流的飞行器无舵面三轴姿态控制飞行试验

将自主可控的合成双射流激励器集成于常规布局飞行器中, 进行了三轴无舵面控制飞行试验, 验证了分布式合成双射流对飞行器巡航时的无舵面姿态调控能力. 对合成双射流激励器进行改进, 设计了分布式三轴姿态控制合成双射流激励器, 滚转环量控制激励器分别安装于两侧机翼翼尖处后缘, 射流出口靠近压力面; 偏航反向合成双射流控制激励器分别安装于靠近两侧机翼翼尖20%弦长处, 上、下沿展向均匀布置; 俯仰环量控制激励器安装于V尾下的平尾后缘, 射流出口靠近压力面. 针对巡航速度为30 m/s的飞行器, 进行了三轴姿态控制飞行试验, 结果表明: 分布式合成双射流实现了飞行器巡航时的三轴无舵面姿态操控; 横航向控制存在耦合, 滚转环量控制激励器实现了飞行器的双向滚转操控, 能产生的最大滚转角速度达16.87°/s, 偏航反向合成双射流控制激励器实现了飞行器的双向偏航操控, 能产生的最大偏航角速度达9.09°/s; 俯仰环量控制激励器实现了飞行器的纵向控制, 能产生的最大俯仰角速度达7.68°/s.      

Mathematical Modeling and Parameter Identification of a Planar Servo-Pneumatic Test Facility. Part I: Mathematical Modeling and Computer Simulation

High quality multi-axis test facilities used for testing heavy loads and large structures of industrial equipment are usually simulated, designed and controlled based on reduced model equations neglecting the inertia properties of the actuators. The design and control of servo-pneumatic test facilities used for testing small and light structures must take into account extended test facility models including the various inertia properties of the actuators. In this paper (Part I) an extended test facility model is presented including the various inertia properties and joints of the actuators. These extended model equations are represented in a form well suited to be directly implemented in control algorithms based on exact linearization techniques for real time control. This is done by stepwise projecting the inertia properties of the various actuator housings and actuator pistons down to the common mass of the test table and payload. The resulting extended model equations have the same form as the reduced model equations. They only include more complex system matrices and vector functions. These compact model equations turn out to be suitable for an efficient nonlinear controller design of these test facilities. Computer simulations and associated laboratory experiments show the necessity to use extended model equations in case of testing small and light structures. In Part II of this paper [1] the inertia parameters of the planar test facility will be identified in laboratory experiments.     

Chaotic pitch motion of an asymmetric non-rigid spacecraft with viscous drag in circular orbit

We study the pitch motion dynamics of an asymmetric spacecraft in circular orbit under the influence of a gravity gradient torque. The spacecraft is perturbed by a small aerodynamic drag torque proportional to the angular velocity of the body about its mass center. We also suppose that one of the moments of inertia of the spacecraft is a periodic function of time. Under both perturbations, we show that the system exhibits a transient chaotic behavior by means of the Melnikov method. This method gives us an analytical criterion for heteroclinic chaos in terms of the system parameters which is numerically contrasted. We also show that some periodic orbits survive for perturbation small enough.     

Nonlinear adaptive control for quadrotor flying vehicle

In this work, we deal with autonomous tracking and disturbance rejection problem of quadrotor vehicle flying in uncertain environment. The vehicles kinematic and modeling error uncertainties are associated with external disturbance, inertia, mass, and nonlinear aerodynamic forces and moments. The proposed method integrate the techniques from adaptive control and robust control theory. Robust and adaptive control algorithms for translational and orientation tracking are derived using Lyapunov method. It is shown in our analysis that the altitude, position, and attitude tracking errors are bounded and their bounds asymptotically converge to zero in Lyapunov sense. Simulation results on a commercial quadrotor flying vehicle are given to demonstrate the effectiveness of theoretical arguments for real world application.     

非自旋航天器混沌姿态运动及其参数开闭环控制

研究万有引力场中受大气阻力且存在结构内阻尼的非自旋航天器在椭圆轨道上平面天平动的混沌及其参数开闭环控制问题．在建立数学模型的基础上确定出现混沌的必要条件并数值验证混沌的存在性，提出非线性振动系统混沌运动的参数开闭环控制并应用于控制航天器的混沌姿态运动．     

Fault-tolerant sliding mode attitude control for flexible spacecraft under loss of actuator effectiveness

In this paper, a novel fault-tolerant attitude control synthesis is carried out for a flexible spacecraft subject to actuator faults and uncertain inertia parameters. Based on the sliding mode control, a fault-tolerant control law for the attitude stabilization is first derived to protect against the partial loss of actuator effectiveness. Then the result is extended to address the problem that the actual output of the actuators is constrained. It is shown that the presented controller can accommodate the actuator faults, even while rejecting external disturbances. Moreover, the developed control law can rigorously enforce actuator-magnitude constraints. An additional advantage of the proposed fault-tolerant control strategy is that the control design does not require a fault detection and isolation mechanism to detect, separate, and identify the actuator faults on-line; the knowledge of certain bounds on the effectiveness factors of the actuator is not used via the adaptive estimate method. The associated stability proof is constructive and accomplished by the development of the Lyapunov function candidate, which shows that the attitude orientation and angular velocity will globally asymptotically converge to zero. Numerical simulation results are also presented which not only highlight the ensured closed-loop performance benefits from the control law derived here, but also illustrate its superior fault tolerance and robustness in the face of external disturbances when compared with the conventional approaches for spacecraft attitude stabilization control.