Attitude Motion Of Spacecraft Research Articles

Hybrid Nonlinear Passivity-Based Control Approach to Magnetic-Impulsive Spacecraft Attitude Regulation

This paper proposes two novel passivity-based control design frameworks for hybrid nonlinear time-varying dynamical systems involving an interacting amalgam of continuous-time and discrete-time dynamics whose dynamical properties evolve periodically over time. In this regard, the hybrid Kalman–Yakubovich–Popov conditions are employed in tandem with the passivity theorem to construct a three-step control design algorithm to render closed-loop dynamics stable. The proposed control architectures are subsequently exploited to regulate the attitude motion of spacecraft with magnetic and impulsive modes of operation. Practical considerations involved in implementing the proposed hybrid algorithms are then discussed in detail. Simulation results show significant improvement in the performance of the attitude control system in terms of system response, robustness, and the required magnetic and impulsive control usage as compared to a hybrid linear approach.

DEEP LEARNING FOR SPACE DEBRIS REMOVAL

The advances in deep learning have revolutionized the field of artificial intelligence. These advances, as well as new tasks and requirements in space exploration, have led to an increased interest in these deep learning methods among space scientists and practitioners. The problems of controlling the attitude and relative motion of spacecraft are considered for both traditional and new missions, such as contactless space debris removal. Both supervised and reinforcement learning is used to solve such problems based on various architectures of artificial neural networks, including convolutional ones. The possibility of using deep learning together with methods of control theory is analyzed to solve the considered problems more efficiently. The difficulties that limit the application of these methods for space applications are highlighted. The necessary research directions for solving these problems are indicated.

Experimental Validation of Constrained Spacecraft Attitude Planning via Invariant Sets

This paper experimentally validates the invariant-set motion planner (ISMP) for the spacecraft attitude motion planning problem. Three novel results are presented that enable the experimental implementation: i) a method for gridding quaternions from the keep-in cone, ii) a method for scaling the invariant sets to enforce angular velocity constraints, and iii) a method for scaling the sets used by the ISMP to ensure their positive invariance despite this torque constraint. The ISMP is experimentally validated through three experimental scenarios. In the first scenario, the spacecraft must perform a re-orientation maneuver that caused it to move toward a keep-out cone. The ISMP manages the momentum of the spacecraft to prevent it from overshooting into the keep-out cone. In the second scenario, the spacecraft performs a slalom maneuver to avoid a pair of keep-out cones. The ISMP must reverse the momentum of the spacecraft to transition from avoiding the first keep-out cone to the second. The final scenario is an unrealistically difficult scenario designed to stress-test the capabilities of the ISMP where the spacecraft must escape from a maze of keep-out cones. These results demonstrate the ability of the ISMP to control the spacecraft attitude while enforcing state and input constraints.

Minimum-Time Spacecraft Attitude Motion Planning Using Objective Alternation in Derivative-Free Optimization

This work presents an approach to spacecraft attitude motion planning which guarantees rest-to-rest maneuvers while satisfying pointing constraints. Attitude is represented on the group of three dimensional rotations. The angular velocity is expressed as weighted sum of some basis functions, and the weights are obtained by solving a constrained minimization problem in which the objective is the maneuvering time. However, the analytic expressions of objective and constraints of this minimization problem are not available. To solve the problem despite this obstacle, we propose to use a derivative-free approach based on sequential penalty. Moreover, to avoid local minima traps during the search, we propose to alternate phases in which two different objective functions are pursued. The control torque derived from the spacecraft inverse dynamics is continuously differentiable and vanishes at its endpoints. Results on practical cases taken from the literature demonstrate advantages over existing approaches.

Coupled Relative Orbit and Attitude Control Augmented by the Geomagnetic Lorentz Propulsions

An electrostatically charged spacecraft is subject to the Lorentz force and torque when moving in the geomagnetic field. By active modulation of the surface charge, the induced Lorentz force and torque can be used for orbit and attitude control, respectively. Because of the electromagnetic mechanism, the orbit and attitude motion of such spacecraft are naturally coupled, and the induced Lorentz force and torque are both instantaneously underactuated. Thus, other kinds of actuation are required to render the system fully actuated and controllable. A hybrid control system is then proposed in this paper for Lorentz-augmented spacecraft relative orbit and attitude control. The relative orbit motion is controlled by the combination of the Lorentz force and the thruster-generated control force, and the relative attitude motion is governed by the combination of the Lorentz torque and the magnetic torque. By using the adaptive backstepping control method, a closed-loop control scheme is designed for the hybrid control system to deal with the dynamic coupling, underactuation, unknown external disturbances, and model approximation errors simultaneously. The parameter adaptation laws are derived via the Lyapunov theory to guarantee the stability of the closed-loop system. Optimal distribution laws of the hybrid control inputs are then analytically solved. Finally, numerical examples are simulated in a -perturbed environment to verify the feasibility of the proposed hybrid control system and the validity of the closed-loop control scheme.

Расчет аэродинамического момента в задачах математического моделирования вращательного движения транспортных грузовых кораблей Прогресс

We describe the method of representation the aerodynamic torque in problems of mathematical modeling uncontrolled attitude motion of spacecraft Progress. The components of the torque are calculated in the nodes of the grid on the two-dimensional sphere. The calculations were based on the detailed geometric model of the spacecraft external surface. The found arrays of torque data were approximated by the segments of the Fourier series by spherical functions up to the order (8,8) inclusive. These approximations are used in the equations of attitude motion of the spacecraft. We describe the use of such equations in reconstruction of the motion of spacecraft Progress MS-07, MS-08 in the gravitational orientation mode by measurements of their angular rates and in the prognosis of electric charge from the solar arrays of the spacecraft in the mode of its single-axis solar orientation.

Deep learning for spacecraft guidance, navigation, and control

The advances in deep learning have revolutionized the field of artificial intelligence, demonstrating the ability to create autonomous systems with a high level of understanding of the environments where they operate. These advances, as well as new tasks and requirements in space exploration, have led to an increased interest in these deep learning methods among space scientists and practitioners. The goal of this review article is to analyze the latest advances in deep learning for navigation, guidance, and control problems in space. The problems of controlling the attitude and relative motion of spacecraft are considered for both traditional and new missions, such as orbital service. The results obtained using these methods for landing and hovering operations considering missions to the Moon, Mars, and asteroids are also analyzed. Both supervised and reinforcement learning are used to solve such problems based on various architectures of artificial neural networks, including convolutional and recurrent ones. The possibility of using deep learning together with methods of control theory is analyzed to solve the considered problems more efficiently. The difficulties that limit the application of the reviewed methods for space applications are highlighted. The necessary research directions for solving these problems are indicated.

Multifarious Chaotic Attractors and Its Control in Rigid Body Attitude Dynamical System

The Euler dynamical equation which describes the attitude motion of a rigid body will exhibit very complex dynamic behaviors under the action of different external torques. Many special types of new chaotic attractors are presented, including hidden attractors, double-body-double-core chaotic attractors, and single-body-three-core-tree-wing chaotic attractors. The position of equilibrium points in several typical cases of the Euler dynamic equation is solved, and the stability of linearized equation at each equilibrium point and its influence on the formation of the chaotic attractor are analyzed. An improved nonlinear relay control law based on Euler angle feedback is developed to stabilize a new chaotic spacecraft attitude motion to an appointed equilibrium point or a periodic orbit.

UNCONSTRAINED MODAL ANALYSIS OF DYNAMIC MODEL OF FLEXIBLE SPACECRAFT$^{\bf 1)}$

The influence coefficient of flexible coupling in dynamic modeling of flexible spacecraft is an important mechanical concept in dynamic modeling, which reflects the elastic vibration effect of spacecraft attitude and orbit motion and flexible accessories. The equivalent relationship between the influence coefficients of the flexible coupling, i.e. the inertial completeness criterion, is an important basis for the reduction of the order and the mode truncation of the dynamic model of the flexible spacecraft. Taking the center rigid body spacecraft with flexible appendages as the research object, the constrained mode and unconstrained mode are used to describe the structural deformation of flexible appendages, and the dynamic model of flexible spacecraft is established by using Euler Lagrange equation. Based on the research results of Hughes, the unconstrained modal identity of flexible spacecraft and the inertial completeness criterion for dynamic model reduction are proved and applied. The relationship between the inertia of two dynamic models is discussed, and the inertial completeness criterion of unconstrained mode is derived by using the inertial completeness criterion of constrained mode. Finally, the numerical simulation of the flexible spacecraft model composed of the central rigid body with two side solar panels and one side solar panels is carried out to find out the unconstrained mode translational coupling coefficient of the flexible appendages. The change of the unconstrained mode eigenvalue and translational coupling coefficient with the rigid flexible mass ratio is analyzed, and the mass characteristic identity of the unconstrained mode inertial completeness criterion is used to test the model The flexible spacecraft model is tested.

Исследование эволюции вращательного движения спутника Фотон-12

This preprint presents the results of the repeated reconstruction of the uncontrolled attitude motion of the spacecraft Foton-12 (was in orbit 09.09 – 24.09.1999) by measurements of onboard magnetometers. This problem was already solved successfully a few years ago. The repeated reconstruction was made by using the more simple mathematical model of spacecraft attitude motion than the model used in processing just after the flight. Simplifications were made in order the new model corresponded to the models in Beletsky's theory of spacecraft attitude motion closed to Euler's regular precession of an axially symmetric rigid body. The results of processing measurements on 15 short (3.5 hours) time intervals are directly compared with the averaged equations arising in Beletsky’s theory. This allowed us to interpret the evolution of the rotational motion of Foton-12 and to explain the occurrence of the motion regime established at the end of the flight.

Control of nonlinear spacecraft attitude motion via state augmentation, Lyapunov-Floquet transformation and normal forms

This article analyzes and controls the quasi-periodic attitude motion of a gravity-gradient stabilized spacecraft in eccentric orbit by way of system states augmentation, Lyapunov-Floquet transformation and normal forms simplification. Perturbing torques in the ambient space environment can be shown to engender spacecraft attitude motion represented by nonlinear dynamics coupled in the roll-yaw axes; and, uncoupled planar dynamics in the pitch axis. The non-planar dynamics equations are homogeneous and analytically solvable. However, the pitch attitude motion is nonlinear, possesses parameter-varying coefficients and is subjected to external periodic excitations. Consequently, we transform the unwieldy attitude dynamics into relatively more amenable schemes for analysis and control law synthesis. Subsequently, we demonstrate the implementation of linear and nonlinear control laws (i.e. bifurcation and sliding mode control laws) on the relatively acquiescent transformed attitude dynamics. By employing a two-pronged approach, the quasi-periodic planar motion is independently shown to be stabilizable via the nonlinear control approaches.

Robust stabilization of spacecraft attitude motion under magnetic control through time‐varying integral sliding mode

SummaryMagnetic spacecraft attitude control is suitable for missions with moderate pointing and stabilization requirements. It would be especially preferable if an implemented three‐axis magnetic control law guarantees global and moreover robust stabilization. Motivated by this consideration, in this work, the integral sliding mode control method is successfully applied to the purely magnetic attitude control problem for the first time in literature. The resulting magnetic integral sliding mode control system, which is designed via Lyapunov's direct method by taking the four major environmental disturbance torques and the full inertia matrix uncertainty into account, has the expected insensitivity against environmental disturbances in the plane of continuous actuation. Along the instantaneous direction of underactuation, which continuously varies in control space as the satellite moves along the orbit, there is no counteraction to perturbations. However, as seen in the results of the high‐fidelity simulation, state trajectories are ultimately bounded in the vicinity of the reference state independently from the initial state thanks to the proven robustness of the global stability. Performance robustness of the obtained control system against disturbances and model uncertainty is evaluated to be superior through a couple of comparisons with existing control laws even though it is partial due to the intrinsic underactuation in the system.

Deployment of flexible space tether system with satellite attitude stabilization

This paper focuses on the complex dynamics and control of deploying a payload using a long flexible tether from a spacecraft. Two different system models are involved, one for detailed dynamics study and the other for control design. Both the flexibility and extensibility of tether and the attitude motion of spacecraft have been incorporated into the system dynamics. In addition, the orbital dynamics of the tethered system are modeled using a set of modified equinoctial elements and considering the J2 perturbation forces. A two-stage mission design has been proposed to divide the deployment into two parts. The first is an ejection separation stage for the payload to be ejected away from the spacecraft to a predetermined position, and the second is a deployment control stage for active tether deployment by regulating tether tension. The attitude motion control of the spacecraft is performed during deployment based on an optimal fuel control method. A closed-loop feedback controller is developed to achieve the deployment by tracking an open-loop optimal trajectory of tether libration states and control input. Finally, numerical simulations have been performed to demonstrate the feasibility of conceptual mission design and effectiveness of the synthetic control scheme.

Spacecraft Attitude Control Compensating Internal Payload Motion Using Disturbance Observer Technique

Spacecraft attitude motion compensation approach to deal with internal slewing payload motion is addressed in this study. The internal payload motion is generated by a steering mechanism such as steerable gimbals for image acquisition as well as error correction due to orbital and attitude errors. Disturbance caused by payload motion in general cannot be measured using sensing devices. The disturbance effect can be estimated by using the so-called disturbance observer technique. The estimated disturbance can be compensated by a feedforward control law. The new control law can accommodate internal disturbance by a combined conventional feedback and feedforward control for estimated disturbance.

A parallel distributed compensation approach to fuzzy control of spacecraft combined attitude and sun tracking

A spacecraft combined attitude and sun tracking system (CASTS) is a synergized system in which solar array drive assemblies are used as sun trackers and simultaneously as attitude control actuators. This paper, a continuous research on CASTS, addresses its attitude control problem. The kinematics and dynamics equations of a rigid spacecraft attitude motion is inherently nonlinear. In the attitude regulation problem, the attitude motion can be treated as a simple linear system for a constrained range of operating conditions, but it has impacts on the accuracy of the linear model when a model-based controller is implemented. Naturally, this is a compromise between simplicity and accuracy that all design engineers have to face. In this paper, we present a systematic approach to improve the accuracy while preserving the model as linear as possible, by deriving a quasi-linear approximation of a nonlinear spacecraft attitude motion. The quasi-linear approximation is based on the framework of the Takagi–Sugeno (T–S) fuzzy model. If a spacecraft can be modeled in the form of a rule-based T-S fuzzy system that acts as an interpolator between linear state-space systems, an approach called parallel distributed control (PDC) can be used to stabilize the attitude motion. The design philosophy of PDC is to create a simple fuzzy controller, where each rule’s consequent is a control law designed to stabilize the linear system in the corresponding consequent of the spacecraft T-S fuzzy system. Numerical results validate that the attitude and sun-tracking performances are achievable using the proposed PDC strategy.Â

Adaptive Singularity-Free Control Moment Gyroscopes

Design considerations of agile, precise, and reliable attitude control for a class of small spacecraft and satellites using adaptive singularity-free control moment gyroscopes (ASCMG) are presented. An ASCMG differs from that of a conventional control moment gyroscopes (CMG) because it is intrinsically free from kinematic singularities and so does not require a separate singularity avoidance control scheme. Furthermore, ASCMGs are adaptive to the asymmetries in the structural members (gimbal and rotor) as well as misalignments between the center of mass of the gimbal and rotors. Moreover, ASCMG clusters are highly redundant to failure and can function as variable and constant-speed CMGs without encountering singularities. A generalized multibody dynamics model of the spacecraft–ASCMG system is derived using the variational principles of mechanics, relaxing the standard set of simplifying assumptions made in prior literature on CMG. The dynamics model so obtained shows the complex nonlinear coupling between the internal degrees of freedom associated with an ASCMG and the spacecraft attitude motion. The general dynamics model is then extended to include the effects of multiple ASCMG, called the ASCMG cluster, and the sufficient conditions for nonsingular cluster configurations are obtained. The adverse effects of the simplifying assumptions that lead to the intricate design of the conventional CMG, and how they lead to singularities, become apparent in this development. A bare-minimum hardware prototype of an ASCMG using low-cost commercial off-the-shelf components is developed to show the design simplicity and scalability. A geometric variational integration scheme is obtained for this multibody spacecraft–ASCMG system for numerical and embedded implementation. Attitude pointing control of a CubeSat with three ASCMGs in the absence of external torques is numerically simulated to demonstrate the singularity-free characteristics and redundancy of the ASCMG cluster.

Определение вращательного движения космического аппарата по измерениям звездных датчиков и датчика угловой скорости методом наименьших квадратов

Определение вращательного движения космического аппарата по измерениям звездных датчиков и датчика угловой скорости методом наименьших квадратов

К задаче оценивания ошибок измерений в системах управления при неполной информации

The article deals with the problem of increasing the accuracy of the state estimation of linear dynamical systems under conditions of incomplete information. This problem is linked to the feature of guaranteed estimation methods, which is that the allowable sets of values of disturbances acting on the system and measurement errors in the information-measuring channel can be only rough upper bound on a short observation interval. In particular, for a single short measurement implementation, the probability of the realization of disturbances and measurement errors is worst-case less than one. The approach to adaptive algorithm development of guaranteed estimation is proposed. The approach is based on processing the innovation sequence values of the Kalman filter. The Kalman filter implementation is performed for measurement data preprocessing. The innovation sequence is considered as a time series for processing of which statistical and guaranteed estimation methods are used. The results of processing the innovation sequence values are used to construct bounded estimates of the measurement errors used in the equations of the guaranteed estimation algorithm. With this approach an informational definition of unknown measurements errors is carried out. The main feature of the problem studied in this article is a small number of available measurements the results of which are used to find the best estimate of the state vector. Therefore, the implementation of measurements is considered as a short time-series. To reduce the computational cost of implementing the guaranteed real time estimation algorithm, methods for decomposition of the estimation problem are proposed. The effectiveness of the proposed approach is demonstrated by the example of a model describing the spacecraft attitude motion. The results of simulation and a comparative analysis of the accuracy of the obtained estimates of the state vector are presented. The results of simulation and a comparative analysis of the accuracy of the obtained state estimates are given.

Dynamics Modeling and Analysis of Spacecraft with Large Deployable Hoop-Truss Antenna

Large deployable hoop-truss antennas are widely used in satellite communication systems for their high antenna gain and small packed volume. Knowledge of the deployment dynamics of the large deployable hoop-truss antenna and its coupling with spacecraft attitude motion are essential for the design of deployment mechanisms and attitude control systems. In this paper, a spacecraft with two flexible solar panels, which are reduced by Craig–Bampton component modal synthesis, is modeled by natural coordinate formulation. A large deployable hoop-truss antenna is modeled with the absolute nodal coordinate formulation. Both lead to a constant-mass matrix without the centrifugal and Coriolis forces. The whole dynamics model is used to solve the attitude and the reaction moment during the deployment process. Nonsynchronous deployment phenomena due to the structural flexibility and the decay of the driving force are investigated, and results show that these phenomena can increase the stresses suffered by the hoop truss, which should be considered in the design of a large deployable hoop-truss antenna.

Output feedback second order sliding mode control for spacecraft attitude and translation motion

This paper studies an output feedback control problem for spacecraft position and attitude control when uncertainties related to system parameters and external disturbances are present. Firstly, a new finite-time control law is designed using second order sliding mode concepts. In the presence of external disturbances and inertia uncertainties, the new control law provides finite-time convergence and high tracking precision. Secondly, a new sliding-mode-based filter is developed to estimate the first time derivatives of attitude and position in finite time. Instead of the translational and angular velocity variables, the estimated derivative values are used for the controller design. The proposed controller with this filter is an output feedback controller since translational and angular velocity measurements are not required. The closed-loop system under this controller is non-homogeneous and the stability is proven by using concepts of a strong Lyapunov function and Lyapunov stability theory. The trajectories of the closed-loop system can be controlled to converge to a ball centered at the origin that can be made as small as desired. Numerical simulations of position and attitude control of spacecraft are given to demonstrate the performance of the proposed controller and filter.