Large-angle attitude tracking maneuver of a spacecraft

Discusses how to transfer an attitude tracking problem with time variable boundary condition into a fixed boundary problem by means of line of sight coordinate, rotational quaternion and relative error feed back signals, and the control law of attitude tracking designed based on the quasi Euler signals and describes the simulation of a forewarning satellite monitoring a low orbital spacecraft to prove the correctness of the design method.

Chaotic attitude and reorientation maneuver for completely liquid-filled spacecraft with flexible appendage

The present paper investigates the chaotic attitude dynamics and reorientation maneuver for completely viscous liquid-filled spacecraft with flexible appendage. All of the equations of motion are derived by using Lagrangian mechanics and then transformed into a form consisting of an unperturbed part plus perturbed terms so that the system＇s nonlinear characteristics can be exploited in phase space. Emphases are laid on the chaotic attitude dynamics produced from certain sets of physical parameter values of the spacecraft when energy dissipation acts to derive the body from minor to major axis spin. Numerical solutions of these equations show that the attitude dynamics of liquid-filled flexible spacecraft possesses characteristics common to random, non- periodic solutions and chaos, and it is demonstrated that the desired reorientation maneuver is guaranteed by using a pair of thruster impulses. The control strategy for reorientation maneuver is designed and the numerical simulation results are presented for both the uncontrolled and controlled spins transition.

Attitude maneuver of liquid-filled spacecraft with a flexible appendage by momentum wheel

Attitude maneuver of liquid-filled spacecraft with an appendage as a cantilever beam by momentum wheel is studied. The dynamic equations are derived by conserva- tion of angular momentum and force equilibrium principle. A feedback control strategy of the momentum wheel is ap- plied for the attitude maneuver. The residual nutation of the spacecraft in maneuver process changes with some chosen parameters, such as steady state time, locations of the liq- uid container and the appendage, and appendage parame- ters. The results indicate that locations in the second and fourth quadrants of the body-fixed coordinate system and the second quadrant of the wall of the main body are better choices for.placing the liquid containers and the appendage than other locations if they can be placed randomly. Higher density and thicker cross section are better for lowering the residual nutation if they can be changed. Light appendage can be modeled as a rigid body, which results in a larger residual nutation than a flexible model though. The resid- ual nutation decreases with increasing absolute value of the initial sloshing angular height.

Adaptive variable structure control based on backstepping for spacecraft with reaction wheels during attitude maneuver

An adaptive variable structure control method based on backstepping is proposed for the attitude maneuver problem of rigid spacecraft with reaction wheel dynamics in the presence of uncertain inertia matrix and external disturbances. The proposed control approach is a combination of the backstepping and the adaptive variable structure control. The cascaded structure of the attitude maneuver control system with reaction wheel dynamics gives the advantage for applying the backstepping method to construct Lyapunov functions. The robust stability to external disturbances and parametric uncertainty is guaranteed by the adaptive variable structure control. To validate the proposed control algorithm, numerical simulations using the proposed approach are performed for the attitude maneuver mission of rigid spacecraft with a configuration consisting of four reaction wheels for actuator and three magnetorquers for momentum unloading. Simulation results verify the effectiveness of the proposed control algorithm.

Robust Adaptive Attitude Maneuvering and Vibration Reducing Control of Flexible Spacecraft Under Input Saturation

A robust adaptive control scheme is proposed for attitude maneuver and vibration suppression of flexible spacecraft in situations where parametric uncertainties,external disturbances,unmeasured elastic vibration and input saturation constraints exist. The controller does not need the knowledge of modal variables but the estimates of modal variables provided by appropriate dynamics of the controller. The requirements to know the system parameters and the bound of the external disturbance in advance are also eliminated by adaptive updating technique. Moreover,an auxiliary design system is constructed to analyze and compensate the effect of input saturation,and the state of the auxiliary design system is applied to the procedure of control design and stability analysis. Within the framework of the Lyapunov theory,stabilization and disturbance rejection of the overall system are ensured. Finally,simulations are conducted to study the effectiveness of the proposed control scheme,and simulation results demonstrate that the precise attitude control and vibration suppression are successfully achieved.

Robust Adaptive Attitude Maneuvering and Vibration Reducing Control of Flexible Spacecraft with Prescribed Performance

A robust adaptive control scheme with prescribed performance is proposed for attitude maneuver and vibration suppression of flexible spacecraft,in which the parametric uncertainty,external disturbances and unmeasured elastic vibration are taken into account simultaneously.On the basis of the prescribed performance control(PPC)theory,the prescribed steady state and transient performance for the attitude tracking error can be guaranteed through the stabilization of the transformed system.This controller does not need the knowledge of modal variables.The absence of measurements of these variables is compensated by appropriate dynamics of the controller which supplies their estimates.The method of sliding mode differentiator is introduced to overcome the problem of explosion of complexity inherent in traditional backstepping design.In addition,the requirements of knowing the system parameters and the unknown bound of the lumped uncertainty,including external disturbance and the estimate error of sliding mode differentiator,have been eliminated by using adaptive updating technique.Within the framework of Lyapunov theory,the stability of the transformed system is obtained.Finally,numerical simulations are carried out to verify the effectiveness of the proposed control scheme.

Variable structure attitude maneuver and vibration control of flexible spacecraft

A dual-stage control system design method is presented for the three-axis-rotational maneuver and vibration stabilization of a spacecraft with flexible appendages embedded with piezoceramics as sensor and actuator. In this design approach, the attitude control and the vibration suppression sub-systems are designed separately using the lower order model. The design of attitude controller is based on the variable structure control (VSC) theory leading to a discontinuous control law. This controller accomplishes asymptotic attitude maneuvering in the closed-loop system and is insensitive to the interaction of elastic modes and uncertainty in the system. To actively suppress the flexible vibrations, the modal velocity feedback control method is presented by using piezoelectric materials as additional sensor and actuator bonded on the surface of the flexible appendages. In addition, a special configuration of actuators for three-axis attitude control is also investigated: the pitch attitude controlled by a momentum wheel, and the roll/yaw control achieved by on-off thrusters, which is modulated by pulse width pulse frequency modulation technique to construct the proper control torque history. Numerical simulations performed show that the rotational maneuver and vibration suppression are accomplished in spite of the presence of disturbance torque and parameter uncertainty.

Maneuver and vibration reduction of flexible spacecraft using sliding mode/command shaping technique

A generalized scheme based on the sliding mode and component synthesis vibration suppression (CSVS) method has been proposed for the rotational maneuver and vibration suppression of an orbiting spacecraft with flexible appendages. The proposed control design process is twofold: design of the attitude controller followed by the design of a flexible vibration attenuator. The attitude controller using only the attitude and the rate information for the flexible spacecraft (FS) is designed to serve two purposes: it forces the attitude motion onto a pre-selected sliding surface and then guides it to the state space origin. The shaped command input controller based on the CSVS method is designed for the reduction of the flexible mode vibration, which only requires information about the natural frequency and damping of the closed system. This information is used to discretize the input so that minimum energy is injected via the controller to the flexible modes of the spacecraft. Additionally, to extend the CSVS method to the system with the on-off actuators, the pulse-width pulse-frequency (PWPF) modulation is introduced to control the thruster firing and integrated with the CSVS method. PWPF modulation is a control method that provides pseudo-linear operation for an on-off thruster. The proposed control strategy has been implemented on a FS, which is a hub with symmetric cantilever flexible beam appendages and can undergo a single axis rotation. The results have been proven the potential of this technique to control FS.

Flexible Satellite Attitude Control via Adaptive Fuzzy Linearization

The adaptive fuzzy control is combined with input-output linearization control to constitute the hybrid controller. The control method is then applied to the attitude maneuver control of the flexible satellite. The basic control structure is given. The rules of the controller parameter selection, which guarantee the attitude stabilization of the satellite with parameter uncertainties, have been analyzed. Simulation results show that the precise attitude control is accomplished in spite of the uncertainty in the system.

Exponential Time-varying Sliding Mode Control for Large Angle Attitude Eigenaxis Maneuver of Rigid Spacecraft

An eigenaxis maneuver strategy with global robustness is studied for large angle attitude maneuver of rigid spacecraft. A sliding mode attitude control algorithm with an exponential time-varying sliding surface is designed, which guarantees the sliding mode occurrence at the beginning and eliminates the reaching phase of time-invariant sliding mode control. The proposed control law is global robust against matched external disturbances and system uncertainties, and ensures the eigenaxis rotation in the presence of disturbances and parametric uncertainties. The stability of the control law and the existence of global siding mode are proved by Lyapunov method. Furthermore, the system states can be fully predicted by the analytical solution of state equations, which indicates that the attitude error does not exhibit any overshoots and the system has a good dynamic response. A control torque command regulator is introduced to ensure the eigenaxis rotation under the actuator saturation. Finally, a numerical simulation is employed to illustrate the advantages of the proposed control law.

应用变速控制力矩陀螺的航天器姿态自适应控制

为提高敏捷挠性航天器在轨连续机动的快速性和高稳定性,应用变速控制力矩陀螺(variable speed control moment gyroscopes,VSCMGs)作为姿态控制执行机构,提出了一种将观测器与自适应控制结合的姿态控制律与VSCMGs复合操纵律。考虑到机动过程中挠性模态及精确惯量不可知,采用模态观测器和转动惯量估计器对不可测的状态或参数进行辨识,辨识结果用于精确估计前馈补偿力矩,利用Lyapunov分析方法证明了闭环控制系统的稳定性。鉴于VSCMGs实际使用的力矩分配能力、避奇异能力、轮速平衡能力与末态框架角定位能力,分别设计了加权伪逆操纵律与3种对应的零运动。基于雅可比矩阵条件数提出了末态框架角的优选方法,给出了VSCMGs零运动在机动过程不同阶段的部署方案。结果表明:通过连续姿态机动数值仿真验证了所提算法的有效性;VSCMGs在连续机动过程中平滑切换模式,在不同的机动阶段实现了相应功能。模态观测值和惯量估计值在多次机动后收敛至真值附近,经过参数辨识后的控制器使航天器在机动末端更快更稳地达到指向精度要求。

航天器敏捷机动控制技术发展及展望

对航天器敏捷控制技术的发展现状进行综述,分析了航天器敏捷机动在路径规划、高敏捷高稳定度控制等方面面临的技术挑战;从航天器输入激励方面给出了不同路径规划对姿态敏捷机动性能的影响及适用范围。分析了单级路径跟踪控制、执行机构精准输出控制、多级复合控制等方法对提升航天器敏捷机动高稳定控制性能的技术途径。航天器敏捷机动控制作为一项系统性技术,需要从结构动力学、控制架构、测量与执行部件研制等多方面开展综合分析与设计。针对航天器敏捷机动控制技术发展需求及所面临的挑战,从智能自主控制架构设计、自主决策与协同规划技术、高性能部件研制技术等方面,给出敏捷航天器控制技术的发展建议。

近距空战中目标飞机战术机动数学模型建模与仿真

作为空空导弹设计与验证的主要手段,数字仿真技术在制导系统性能评估、抗干扰性能评估、空战对抗仿真等领域发挥着重要的作用。为解决空空导弹数字仿真试验中目标飞机机动样式单一,不能反映近距格斗过程中目标姿态的变化问题,梳理了飞机机动数学模型在导弹总体性能评估与对抗推演等方面的应用背景需求。对六自由度飞机机动简化数学模型的建设进行了分析和论证。构建了飞机作战典型机动动作库、机动动作指令生成器,实现了以较少的参数刻画不同的机动动作。在三自由度质点运动模型基础上,考虑机动过载、目标攻角、横滚角变化,建立目标机动数学模型。根据所建模型,对几种典型机动教学模型进行数字仿真。仿真结果表明,目标机动轨迹与预期吻合,能够满足导弹总体设计、对抗推演对数字仿真的需求。

充液挠性航天器时变滑模控制及振动抑制

针对存在外界未知干扰、参数不确定问题的刚-液-柔多体耦合航天器姿态控制进行了研究.将液体燃料的晃动等效为球摆模型,挠性附件假设为欧拉-伯努利梁,建立了多体耦合航天器动力学方程.首先,设计了积分滑模干扰观测器,使其能够在有限的时间范围内对控制系统的集总扰动实现准确估计;其次,以此干扰观测器为基础,设计了一种时变滑模控制方法,该控制运用双曲正切函数;最后,结合光滑整形技术,设计出零振动指令光滑器,从而抑制液体晃动和挠性附件振动.数值仿真结果表明:本文所设计的控制方案具有可行性和有效性,多模态充液挠性航天器在姿态机动过程中所引起的残余振动可以得到有效抑制.

贵州省公众海姆立克急救法知信行和自我效能现状及影响因素分析

目的了解贵州省公众海姆立克急救法知信行和自我效能现状并分析其影响因素。方法于2023年9-10月,采用便利抽样法,应用自行编制的海姆立克急救法知信行和自我效能问卷对贵州省2887名公众进行调查并统计分析。结果公众海姆立克急救法知信行和自我效能总体得分(78.55±11.30)分,知识得分(15.68±4.99)分、态度得分(40.80±4.69)分、行为得分(9.03±1.11)分、自我效能得分(13.04±4.19)分。回归分析结果显示,公众海姆立克急救法知信行和自我效能得分的共同影响因素是本科及以上文化程度、居住在农村、健康状况一般、不了解海姆立克急救法和未参加过海姆立克急救法培训(P<0.05)。结论公众对海姆立克急救法知识和自我效能处于中等水平。相关部门需加大对本科及以下文化程度、居住在农村、健康状况一般、不了解海姆立克急救法和未参加过海姆立克急救法培训的群体开展知信行和自我效能的培训力度,切实提升公众自救互救能力。

Adaptive Integral-type Sliding Mode Control for Spacecraft Attitude Maneuvering Under Actuator Stuck Failures

A fault tolerant control （FTC） design technique against actuator stuck faults is investigated using integral-type sliding mode control （ISMC） with application to spacecraft attitude maneuvering control system. The principle of the proposed FTC scheme is to design an integral-type sliding mode attitude controller using on-line parameter adaptive updating law to compensate for the effects of stuck actuators. This adaptive law also provides both the estimates of the system parameters and external disturbances such that a prior knowledge of the spacecraft inertia or boundedness of disturbances is not required. Moreover, by including the integral feedback term, the designed controller can not only tolerate actuator stuck faults, but also compensate the disturbances with constant components. For the synthesis of controller, the fault time, patterns and values are unknown in advance, as motivated from a practical spacecraft control application. Complete stability and performance analysis are presented and illustrative simulation results of application to a spacecraft show that high precise attitude control with zero steady-error is successfully achieved using various scenarios of stuck failures in actuators.

An adaptive attitude algorithm based on a current statistical model for maneuvering acceleration

A current statistical model for maneuvering acceleration using an adaptive extended Kalman filter（CS-MAEKF） algorithm is proposed to solve problems existing in conventional extended Kalman filters such as large estimation error and divergent tendencies in the presence of continuous maneuvering acceleration. A membership function is introduced in this algorithm to adaptively modify the upper and lower limits of loitering vehicles＇ maneuvering acceleration and for realtime adjustment of maneuvering acceleration variance. This allows the algorithm to have superior static and dynamic performance for loitering vehicles undergoing different maneuvers. Digital simulations and dynamic flight testing show that the yaw angle accuracy of the algorithm is 30% better than conventional algorithms, and pitch and roll angle calculation precision is improved by 60%.The mean square deviation of heading and attitude angle error during dynamic flight is less than3.05°. Experimental results show that CS-MAEKF meets the application requirements of miniature loitering vehicles.

Heteroclinic bifurcations in attitude maneuver of coupled slosh-spacecraft with flexible appendage

In this paper, the chaotic dynamics in an attitude transition maneuver of a slosh-spacecraft coupled with flexible appendage in going from minor axis to major axis spin under the influence of dissipative effects due to fuel slosh and a small flexible appendage constrained to only torsional vibration is investigated. The slosh-spacecraft coupled with flexible appendage in attitude maneuver carrying a sloshing liquid is considered as multi-body system with the sloshing motion modeled as a spherical pendulum. The focus in this paper is that the dynamics of the liquid and flexible appendage vibration are coupled. The equations of motion are derived and transformed into a form suitable for the application of Melnikov’s method. Melnikov’s integral is used to predict the transversal intersections of the stable and unstable manifolds for the perturbed system. An analytical criterion for chaotic motion is derived in terms of system parameters. This criterion is evaluated for its significance to the design of spacecraft. The dependence of the onset of chaos on quantities such as body shape and magnitude of damping values, fuel fraction and torsional vibration frequency of flexible appendage are investigated. In addition, we show that a spacecraft carrying a sloshing liquid, after passive reorientation maneuver, will end up with periodic limit motion other than a final major axis spin because of the intrinsic non-linearity of fuel slosh. Furthermore, an extensive numerical simulation is carried out to validate the Melnikov’s analytical result.

Spacecraft attitude maneuver control using two parallel mounted 3-DOF spherical actuators

A parallel configuration using two 3-degree-of-freedom（3-DOF） spherical electromagnetic momentum exchange actuators is investigated for large angle spacecraft attitude maneuvers.First, the full dynamic equations of motion for the spacecraft system are derived by the NewtonEuler method. To facilitate computation, virtual gimbal coordinate frames are established. Second,a nonlinear control law in terms of quaternions is developed via backstepping method. The proposed control law compensates the coupling torques arising from the spacecraft rotation, and is robust against the external disturbances. Then, the singularity problem is analyzed. To avoid singularities, a modified weighed Moore-Pseudo inverse velocity steering law based on null motion is proposed. The weighted matrices are carefully designed to switch the actuators and redistribute the control torques. The null motion is used to reorient the rotor away from the tilt angle saturation state. Finally, numerical simulations of rest-to-rest maneuvers are performed to validate the effectiveness of the proposed method.

航天器姿态机动规划技术研究进展

以航天器多约束姿态控制为背景,围绕姿态机动路径的复杂约束处理问题,对姿态机动规划技术的规划模型、方法研究现状与未来发展趋势进行了论述、分析与归纳。首先介绍了航天器姿态机动规划模型,随后梳理并总结了航天器姿态机动规划技术发展历程及现状,从方法逻辑、设计思想等方面对姿态机动规划方法进行分类和关键技术分析,进一步根据技术发展脉络和未来航天任务需求,提出了航天器姿态机动规划技术的若干发展趋势。