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Modern spacecraft are usually equipped with various flexible appendages, which bring great challenges to the design of high-performance control systems. This paper proposes a novel approach to provide active damping to spacecraft flexible appendages. Vibrations of flexible appendages are actively attenuated using a base-mounted angular displacement mechanism and tip-mounted non-collocated sensors. Dynamic models are presented and the controller design techniques provided. Effectiveness of the approach is demonstrated through an experimental apparatus, in which the damping ratio of the target beam is increased to be better than 40%. Furthermore, an active damping-control scheme using no extra sensors beyond attitude sensors onboard the spacecraft bus is proposed and demonstrated through numerical simulations. [ABSTRACT FROM AUTHOR]

The problem of finite-time attitude consensus for multiple rigid spacecraft with a leader–follower architecture is investigated in this paper. To achieve the finite-time attitude consensus, at the first step, a distributed finite-time convergent observer is proposed for each follower to estimate the leader's attitude in a finite time. Then based on the terminal sliding mode control method, a new finite-time attitude tracking controller is designed such that the leader's attitude can be tracked in a finite time. Finally, a finite-time observer-based distributed control strategy is proposed. It is shown that the attitude consensus can be achieved in a finite time under the proposed controller. Simulation results are given to show the effectiveness of the proposed method. [ABSTRACT FROM AUTHOR]

The attitude optimal control problem (OCP) of a two-rigid-body spacecraft with two rigid bodies coupled by a ball-in-socket joint is considered. Based on conservation of angular momentum of the system without the external torque, a dynamic equation of three-dimensional attitude motion of the system is formulated. The attitude motion planning problem of the coupled-rigid-body spacecraft can be converted to a discrete nonlinear programming (NLP) problem using the Chebyshev-Gauss pseudospectral method (CGPM). Solutions of the NLP problem can be obtained using the sequential quadratic programming (SQP) algorithm. Since the collocation points of the CGPM are Chebyshev-Gauss (CG) points, the integration of cost function can be approximated by the Clenshaw-Curtis quadrature, and the corresponding quadrature weights can be calculated efficiently using the fast Fourier transform (FFT). To improve computational efficiency and numerical stability, the barycentric Lagrange interpolation is presented to substitute for the classic Lagrange interpolation in the approximation of state and control variables. Furthermore, numerical float errors of the state differential matrix and barycentric weights can be alleviated using trigonometric identity especially when the number of CG points is large. A simple yet efficient method is used to avoid sensitivity to the initial values for the SQP algorithm using a layered optimization strategy from a feasible solution to an optimal solution. Effectiveness of the proposed algorithm is perfect for attitude motion planning of a two-rigid-body spacecraft coupled by a ball-in-socket joint through numerical simulation. [ABSTRACT FROM AUTHOR]

The robust adaptive learning control for spacecraft autonomous proximity is investigated in this paper. The relative motion dynamics is firstly proposed and transformed into the state space representation. An adaptive learning controller that allows performing autonomous proximity maneuvers in the presence of uncertain orbital parameters is then developed. The unknown disturbances are considered, and a modified adaptive learning controller is therefore designed in the presence of uncertain parameters and unknown disturbances. The asymptotic stability of the proposed controllers for the closed-loop systems is proven based on the Lyapunov theory. Different numerical results are finally provided and compared to illustrate the performance of the proposed adaptive learning controllers. [ABSTRACT FROM AUTHOR]

This article discusses the design-ballistic optimization for spacecraft, which move under the influence of the gravitational fields of complex configuration and electric propulsion. Such conditions exist near asteroids, for example in the systems of two gravitating bodies, in the vicinity of the libration points. In this paper, we propose to use a method of iterative mission optimization for electric propulsion spacecraft, using sequences of the mathematically specified models of movement and the spacecraft design. The previously obtained analytical results, which describes the planar movement of spacecraft with solar electric propulsion, allows the initial approach in the iterative scheme of optimization to be constructed. A method of modeling and optimization of interplanetary missions ballistic schemes is developed, based on a combination of Pontryagin's maximum principle formalism conditions of transversality and methods of mathematical programming, allowing restriction characteristics for real interplanetary missions to be considered. Recommendations for the choice of the design-ballistic parameters of the interplanetary mission's spacecraft are obtained taking into account the features of a nuclear and solar power plant for purposes of the delivery of a payload to the required heliocentric orbit, to the Moon and near asteroid. [ABSTRACT FROM AUTHOR]

The problem of the optimal control of spacecraft (SC) with an electric ramjet (ER) to change the orbital parameters is considered. ER uses an air as a fuel. Optimal control is determined by solving the problem on the conditional maximum of the local (at the current point of the orbit) impact of the perturbing acceleration on the functional, assuming the perturbations’ smallness compared to gravity. An approximate synthesis of the optimal thrust vector is obtained in view of the power limitation. Parametric investigations of the apogee's changes, depending on the SC with ER parameters, were carried out. The efficiency of the energy and/or fuel accumulation for the implementation of the given orbital flight of SC with ER is shown. [ABSTRACT FROM AUTHOR]

Models of spacecraft guided motion in the gravitational fields of different configurations were developed. Parameterization of the optimal control laws for all the trajectory segments of the spacecraft moving in the N-body gravitational field was realized. The dependence patterns of control laws on mission purposes were determined. The simulation of the number of missions in the Solar system was performed: low Earth orbit – L1 and L2 of the Earth-Moon and the Sun-Earth systems, L1 – L2 and L2 – L1 of the Earth-Moon and the Sun-Earth systems, low Earth orbit – asteroid Eros, low Earth orbit – Venus, low Earth orbit – Mars. Moreover, the Venus gravity assist maneuver was considered. [ABSTRACT FROM AUTHOR]

The problem of cooperative optimization of trajectory and main design parameters of spacecraft, equipped with propulsion systems with energy accumulation, during low Earth orbit formation and correction in non-central gravitational field is considered. The working principle of these propulsion systems is a periodic cycle of energy accumulation from the energy source at the passive stages of orbital transfer, and its discharge to supply power to the engine during the active stages of orbital transfer. The special characteristics of propulsion systems with energy accumulation is limited operation time per one switching, which determines by the energy, stored in an accumulator. For design-ballistic analysis, the mass model of the spacecraft with an uncontrollable power-limited engine with energy accumulator is used. The general problem is separated into dynamic and parametric parts. Using averaging mathematical models of motion and Pontriagin's maximum principle, the calculation method of the ballistic characteristics of approximately-optimal orbital transfer is obtained. The method of design-ballistic optimization is obtained as an iterative procedure. It was found, that inclusion of an energy accumulator in a propulsion system provides an increasing of orbital transfer efficiency. [ABSTRACT FROM AUTHOR]

In the mid-90s of the past century Arnon Spitzer offered a flight scheme comprising an elliptical synchronous equatorial orbit for the SC injection into geostationary orbit (GEO). The advantages of such scheme were related to the SC control simplicity during its transfer from the elliptical synchronous orbit into GEO. The main advantage of Spitzer's scheme is the possibility of observing spacecraft from Earth using a limited number of ground stations along the entire trajectory of the SC flight from an intermediate orbit into GEO. Some flight schemes, which preserve the main advantage of the Spitzer's scheme, are analyzed. The inclination and the eccentricity of the intermediate synchronous orbit are optimized. Several variants of the yaw angle control are analyzed. Numerical analysis was carried out for a space transportation system based on the launch vehicle “Angara-A5”, chemical upper stage “KVTK”, and electric propulsion system comprising four thrusters SPD-140 with the specific impulse of 1700 s. [ABSTRACT FROM AUTHOR]

The problem of controlled spacecraft motion in vicinity of asteroids is an important flight dynamics task. Solution of the problem consists of gravitational field simulations and control trajectory scheme defininition. We propose to use the model of single gravity points for motion simulation of a spacecraft in an irregular gravitational field of the Eros 433 asteroid. The equations of spacecraft motion are the corresponding equations of the n-body problem. The motion trajectory can be split into a number of from-point-to-point motion parts. Each part can be considered as a two-point Boundary task. The Authors developed a concept of Boundary problem solution based on the Pontryagin's principle and modified Newton's step-by-step approximation method. The Authors developed a special software for the Boundary problem solutions. A boundary task of the controlled spacecraft's transfer between circular orbits from 200 km to 100 km is presented in the paper. [ABSTRACT FROM AUTHOR]

The ballistic optimization of the L1-L2 and L2-L1 missions in the Earth-Moon system using electric propulsion is considered. The motion control is studied using the Fedorenko's successful linearization method to estimate the derivatives and the gradient method to optimize the control laws. The transfers are designed in frames of the restricted circular three-body problem. The mathematical model of the missions is described within the barycentric system of coordinates. The perturbation from the Earth, the Moon and the Sun are taking into account. The impact factor of the shaded areas, produced by the Earth and the Moon, is also accounted for. As for the results, the optimization of the optimal control laws and corresponding trajectories were obtained. [ABSTRACT FROM AUTHOR]

Spacecraft performance in an orbital maneuver relies on guidance and control systems which manage the thrust direction within orbit transfer. In this article, the guidance and control approach for spacecraft having a 3D orbit transfer mission is proposed. To derive the optimal variation of steering angles with initial and terminal constraints on the space orbits, a mathematics polynomial function of the guidance command with unknown coefficients is introduced, one of which is determined to achieve the transfer accuracy requirement between space orbits. Genetic Algorithm is employed in finding optimal variation of guidance command and the optimal initial states within the transfer. The attitude control system is also modeled to evaluate the spacecraft response with respect to generated commands by the guidance system. Gas thrusters are considered as attitude actuators for space mission and linear controller with pulse-width pulse-frequency modulator and unconstrained control allocation is employed for controlling steering angles. Results indicate that the presented approach for guidance and control system fairly satisfies the mission requirement. [ABSTRACT FROM AUTHOR]

This brief investigates the finite-time control problem associated with attitude stabilization of a rigid spacecraft subject to external disturbance, actuator faults, and input saturation. More specifically, a novel fixed-time sliding mode surface is developed, and the settling time of the defined surface is shown to be independent of the initial conditions of the system. Then, a finite-time controller is derived to guarantee that the closed-loop system is stable in the sense of the fixed-time concept. The actuator-magnitude constraints are rigorously enforced and the attitude of the rigid spacecraft converges to the equilibrium in a finite time even in the presence of external disturbances and actuator faults. Numerical simulations illustrate the spacecraft performance obtained using the proposed controller. [ABSTRACT FROM AUTHOR]

The issue of attitude tracking control with finite-time convergence is investigated for spacecraft in the presence of external disturbance and uncertain actuator configuration as well. More specifically, in view of backstepping technique, the attitude quaternion can firstly track the demanded command in finite time under the virtual control law by introducing a novel terminal sliding manifold. Then, the finite-time convergence to the sliding mode of the angular velocity error will be achieved under the proposed novel finite-time control and a parameter updating law, even under uncertain actuator misalignment. In addition, the finite-time stabilization/convergence of the sliding surface and the spacecraft attitude tracking error have been proved and analyzed using the Lyapunov method theoretically. Finally, numerical simulation results are presented to illustrate the performance of the proposed scheme. [ABSTRACT FROM AUTHOR]

Purpose – The purpose of this paper is to present a variable structure multiple model adaptive estimator (VSMMAE) for liaison navigation system. Liaison navigation is an autonomous navigation method where inter-satellite range measurements are used to estimate the orbits of all participating spacecrafts simultaneously. Design/methodology/approach – To overcome the problem caused by an inaccurate initial state, a navigation algorithm is designed based on the multiple model adaptive estimation technique. The multiple models are constructed by different initial error covariance matrices. To reduce the computational cost, the likely-model set (LMS) algorithm is adopted to eliminate the unlikely models. Findings – It is specified that the performance of the liaison navigation based on the extended Kalman filter (EKF) is sensitive to the initial error. Simulation results show that the VSMMAE outperforms the EKF in the presence of a large initial error. Practical implications – The presented algorithm is applicable to spacecraft autonomous navigation. Originality/value – A novel navigation algorithm based on the VSMMAE is developed. It is an effective method for the liaison navigation system. [ABSTRACT FROM AUTHOR]

A dynamic control allocation approach is presented to address the attitude stabilization problem of a rigid spacecraft. The approach is developed by using a least-square support vector machine. Actuator uncertainty including misalignment and magnitude deviation is explicitly addressed. A dynamic inverse control law is firstly designed. A least-square support vector machine-based adaptive compensator is then designed to handle actuators uncertainties, external disturbances and unknown moment of inertia. Lyapunov stability analysis shows that the closed-loop attitude system is asymptotically stable. More specifically, constrained quadratic programming-based robust dynamic control allocation is implemented to manage the redundancy actuators. The goal of minimizing the assumption of total energy is achieved. A numerical example is provided to demonstrate the effectiveness of the proposed scheme. [ABSTRACT FROM AUTHOR]

SUMMARY The problem of fault tolerant attitude stabilization with finite-time convergence is investigated for spacecraft with redundant actuators. On the basis of the sliding mode control technique, a robust controller is derived with uncertain inertia parameters, actuator faults, and external disturbances explicitly addressed. It is shown that finite-time reachability into the small neighborhood of sliding surface, and faster time convergence of attitude orientation are achieved. To address actuator input constraints, an adaptive fault tolerant controller is further proposed. One feature of the proposed strategy is that the design of the fault tolerant control does not require any fault detection and isolation mechanism to detect, separate, and identify actuator faults. The attitude stabilization performance using the controller is evaluated through a numerical example. Copyright © 2012 John Wiley & Sons, Ltd. [ABSTRACT FROM AUTHOR]

Abstract: Adaptive backstepping control combined with a novel dynamic control allocation scheme is developed for attitude tracking control of a rigid spacecraft, in which multiple constraints are simultaneously considered, such as actuator uncertainties including torque magnitude deviation and misalignment, actuator constraints like saturation, external disturbances and even constrained energy consumption. More specifically, an adaptive backstepping control algorithm is firstly developed for the relative attitude tracking problem, in which the external disturbances and actuator uncertainties are explicitly considered, and Lyapunov stability analysis shows that asymptotical stability of the overall system is guaranteed. Moreover, to overcome the singularity or parameter drift caused by the parameter updating law, a switching parameter modification strategy is involved. In view of the above, a dynamic control allocation is further investigated to distribute the designed total control torques among the individual actuators. This method extends the conventional quadratic programming control allocation by penalizing the previous step before sampling intervals at a lower cost, and also the actuator magnitude and rate limitation are considered simultaneously by an auxiliary saturation function. Finally, a numerical simulation example for a rigid spacecraft is included to illustrate the feasibility and effectiveness of using the proposed scheme. [Copyright &y& Elsevier]

This paper proposes a novel observer-based fault diagnosis method, incorporating an online control allocation scheme, for an orbiting spacecraft in the presence of actuator failures/faults, unexpected disturbances and input saturation. The proposed scheme solves a difficult problem in spacecraft fault tolerant control design by compensating for the loss in effectiveness and time-varying faults using redundant actuators, so that the overall system is stable, despite external disturbances and input saturation. This is accomplished by developing an observer-based fault diagnosis mechanism to reconstruct or estimate the actuator faults/failures. An online control allocation scheme is then used to redistribute the control signals to the healthy actuators in the case of faults/failures, without reconfiguring the controller, so that the control signal distribution is based on the reconstructed actuator effectiveness level. Simulation results using a rigid spacecraft model show good performance in the presence of faults, including total actuator failure scenarios, external disturbances and actuator input saturation, which validates the effectiveness and feasibility of the proposed scheme. [ABSTRACT FROM AUTHOR]

Abstract: In this paper, a 6DOF spacecraft formation control problem is investigated via a composite control method, which consists of a feedback control law based on a finite time control technique and a feedforward compensator based on a nonlinear disturbance observer technique. The composite controller can guarantee that the closed-loop error system is globally asymptotically stable with respect to a set of two equilibria, that is, the 6DOF spacecraft can reach the desired formation. Simulation results are provided to demonstrate the effectiveness of the proposed method. The results show that the control performance of the proposed method in terms of both stabilization and disturbance rejection are considerably better than the traditional method. [Copyright &y& Elsevier]

Control-moment gyroscopes (CMGs) are power. efficient attitude-control actuators that produce high torques for agile spacecraft. We propose the use of CMGs in actuating joint degrees of freedom in a spacecraft-mounted agile imaging payload, whose tasks include acquiring and tracking a high-speed target. High agility, on the order of several radians per second, characterizes the system of interest here; however, such capabilities are achieved with traditional actuators only at the expense of excessively high electrical power. The proposed design provides agile slewing of an imaging payload for a very small fraction of the power required by fixed rotors in a reaction wheel assembly. This study provides a dynamics analysis for a general system, but we focus on an example in which the inertia dyadics are spherical. We demonstrate by simulation that a CMG-driven system offers the same agility with less than 1% of the power of a telescope actuated by reaction wheels. [ABSTRACT FROM AUTHOR]

Robotic on-orbit servicing is the key to the commercial development of a space robotics infrastructure. A major technical difficulty has been the problem of control of the robotic manipulators mounted onto a freeflying platform in space. The problem of control is directly related to the fact that manipulator motions exert reaction effects on the mounting spacecraft. A solution to this problem is outlined—one in which no fuel is expended and that demands no excessive computational resources that would otherwise preclude real-time performance. [ABSTRACT FROM AUTHOR]

The objective of this paper is to present formation control laws for maintaining attitude alignment among a group of spacecraft in either deep space or earth orbit. The paper presents two control strategies based on emergent behavior approaches. Each control strategy considers the desired formation behaviors of convergence to the final formation goal, formation keeping, and the desire to rotate the spacecraft about fixed axes. The first approach uses velocity feedback and the second approach used passivity-based damping. In addition, we prove analytically that our approach guarantees formation keeping throughout the maneuver. Simulation results demonstrate the effectiveness of our approach. [Copyright &y& Elsevier]