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The robust finite-time control for a flexible spacecraft attitude maneuver in the absence and presence of inertia uncertainties and external disturbances is investigated in this paper. The finite-time controllers, based on the Lagrange-like model and nonsingular terminal sliding mode technique, which guarantee the convergence of attitude maneuver errors in finite time rather than in the asymptotic sense, are proposed to perform an attitude maneuver. By constructing a particular Lyapunov function, the convergences of the proposed controllers for the closed-loop systems are proven theoretically. The robustness problem associated with inertia uncertainties and bounded disturbances is addressed. Numerical simulations are finally provided to illustrate the performance of the proposed controller by comparing it with the conventional sliding mode control. The finite-time controller demonstrates superior performance, such as fast convergence performance, ideal robustness, and effectiveness in suppressing vibration. [ABSTRACT FROM AUTHOR]

Abstract: This paper investigates the spacecraft attitude tracking control problem. Two robust sliding mode controllers based on the quaternion and Lagrange-like model are proposed to solve this problem both in the absence of model uncertainties and external disturbances as well as in the presence of these. The controllers can guarantee the convergence of attitude tracking errors in finite time rather than in the asymptotic sense, where time tends to infinity. By constructing a particular Lyapunov function, the convergences of the proposed controllers for the closed-loop systems are proven theoretically. To alleviate the chattering phenomenon while at the same time guaranteeing the finite convergence during the process of attitude tracking, a new function is introduced into the controller. Numerical simulations are finally provided to illustrate the performance of the proposed controllers. [Copyright &y& Elsevier]

A new type of space debris in near geosynchronous orbit (GEO) was recently discovered and later identified as exhibiting unique characteristics associated with high area-to-mass ratio (HAMR) objects, such as high rotation rates and high reflection properties. Observations have shown that this debris type is very sensitive to environmental disturbances, particularly solar radiation pressure, due to the fact that its motion depends on the actual effective area, orientation of that effective area, reflection properties and the area-to-mass ratio of the object is not stable over time. Previous investigations have modelled this type of debris as rigid bodies (constant area-to-mass ratios) or discrete deformed body; however, these simplifications will lead to inaccurate long term orbital predictions. This paper proposes a simple yet reliable model of a thin, deformable membrane based on multibody dynamics. The membrane is modelled as a series of flat plates, connected through joints, representing the flexibility of the membrane itself. The mass of the membrane, albeit low, is taken into account through lump masses at the joints. The attitude and orbital motion of this flexible membrane model is then propagated near GEO to predict its orbital evolution under the perturbations of solar radiation pressure, Earth’s gravity field (J 2 ), third body gravitational fields (the Sun and Moon) and self-shadowing. These results are then compared to those obtained for two rigid body models (cannonball and flat rigid plate). In addition, Monte Carlo simulations of the flexible model by varying initial attitude and deformation angle (different shape) are investigated and compared with the two rigid models (cannonball and flat rigid plate) over a period of 100 days. The numerical results demonstrate that cannonball and rigid flat plate are not appropriate to capture the true dynamical evolution of these objects, at the cost of increased computational time. [ABSTRACT FROM AUTHOR]

Multilayer insulation (MLI) is a recently-discovered type of debris originating from delamination of aging spacecraft; it is mostly detected near the geosynchronous orbit (GEO). Observation data indicates that these objects are characterised by high reflectivity, high area-to-mass ratio (HAMR), fast rotation, high sensitivity to perturbations (especially solar radiation pressure) and change of area-to-mass ratio (AMR) over time. As a result, traditional models (e.g. cannonball) are unsuitable to represent and predict this debris' orbital evolution. Previous work by the authors effectively modelled the flexible debris by means of multibody dynamics to improve the prediction accuracy. The orbit evolution with the flexible model resulted significantly different from using the rigid model. This paper aims to present a methodology to determine the dynamic properties of thin membranes with the purpose to validate the deformation characteristics of the flexible model. A high-vacuum chamber (10 −4 mbar) to significantly decrease air friction, inside which a thin membrane is hinged at one end but free at the other provides the experimental setup. A free motion test is used to determine the damping characteristics and natural frequency of the thin membrane via logarithmic decrement and frequency response. The membrane can swing freely in the chamber and the motion is tracked by a static, optical camera, and a Kalman filter technique is implemented in the tracking algorithm to reduce noise and increase the tracking accuracy of the oscillating motion. Then, the effect of solar radiation pressure on the thin membrane is investigated: a high power spotlight (500–2000 W) is used to illuminate the sample and any displacement of the membrane is measured by means of a high-resolution laser sensor. Analytic methods from the natural frequency response and Finite Element Analysis (FEA) including multibody simulations of both experimental setups are used for the validation of the flexible model by comparing the experimental results of amplitude decay, natural frequencies and deformation. The experimental results show good agreement with both analytical results and finite element methods. [ABSTRACT FROM AUTHOR]

With the development of satellite mobile communications, large antennas are now widely used. The precise pointing of the antenna’s optical axis is essential for many space missions. This paper addresses the challenging problem of high-precision autonomous pointing control of a large satellite antenna. The pointing dynamics are firstly proposed. The proportional–derivative feedback and structural filter to perform pointing maneuvers and suppress antenna vibrations are then presented. An adaptive controller to estimate actual system frequencies in the presence of modal parameters uncertainty is proposed. In order to reduce periodic errors, the modified controllers, which include the proposed adaptive controller and an active disturbance rejection filter, are then developed. The system stability and robustness are analyzed and discussed in the frequency domain. Numerical results are finally provided, and the results have demonstrated that the proposed controllers have good autonomy and robustness. [ABSTRACT FROM AUTHOR]

A Riccati-based tracking controller with collision avoidance capabilities is presented for proximity operations of spacecraft formation flying near elliptic reference orbits. The proposed dynamical model incorporates nonlinear accelerations from an artificial potential field, in order to perform evasive maneuvers during proximity operations. In order to validate the design of the controller, test cases based on the physical and orbital features of the Prototype Research Instruments and Space Mission Technology Advancement (PRISMA) will be implemented, extending it to scenarios with multiple spacecraft performing reconfigurations and on-orbit position switching. The results show that the tracking controller is effective, even when nonlinear repelling accelerations are present in the dynamics to avoid collisions, and that the potential-based collision avoidance scheme is convenient for reducing collision threat. [ABSTRACT FROM AUTHOR]

This paper investigates the attitude control problem in the presence of wheel low-speed friction and external disturbances. The critical issues surrounding wheel friction are firstly discussed, and the spacecraft attitude model and wheel friction model are introduced. A friction observer that can effectively estimate the friction torque is developed. An observer-based sliding mode controller, to perform attitude maneuver, is then proposed. By constructing a particular Lyapunov function, the stability and performance of the proposed controller for the closed-loop systems is discussed theoretically. Numerical simulations are finally provided, and the results have demonstrated that the friction is compensated by the observer and the proposed controller has better transient and steady-state performance. [ABSTRACT FROM AUTHOR]

Abstract: This paper investigates the adaptive control problem for spacecraft relative translation. The critical issues surrounding relative translation are firstly discussed, and then the relative motion model implemented in this paper is introduced. An adaptive controller that allows performing relative translation in the presence of uncertain orbital parameters is developed. The escape acceleration of the target is further considered, and a modified adaptive controller is then proposed in the presence of uncertain orbital parameters and escape trajectory. By constructing a particular Lyapunov function, the stability of the proposed controllers for the closed-loop systems is proven theoretically. Numerical simulations are finally provided to illustrate the performance of the proposed controllers. [Copyright &y& Elsevier]

Purpose - One of the primary problems in the field of on-orbit service and space conflict is related to the approach to the target. The development of guidance algorithms is one of the main research areas in this field. The objective of this paper is to address the guidance problem for autonomous proximity manoeuvres of a chase-spacecraft approaching a target spacecraft. Design/methodology/approach - The process of autonomous proximity is divided into three phases: proximity manoeuvre, fly-around manoeuvre, and final approach. The characteristics of the three phases are analyzed. Considering the time factor of autonomous proximity, different orbits for the three phases are planned. Different guidance algorithms, which are based on multi-pulse manoeuvres, are then devised. Findings - This paper proposes three phases of autonomous proximity and then designs a guidance method, which hinges on a multi-pulse algorithm and different orbits for the three phases; in addition, a method of impulse selection is devised. Practical implications - An easy methodology for the analysis and design of autonomous proximity manoeuvres is proposed, which could also be considered for other space applications such as formation flying deployment and reconfiguration. Originality/value - Based on this guidance method, the manoeuvre-flight period of the chase-spacecraft can be set in accordance with the mission requirements; the constraints on fuel mass and manoeuvre time are both considered and satisfied. Consequently, this proposed guidance method can effectively deal with the problem of proximity approach to a target spacecraft. [ABSTRACT FROM AUTHOR]

Abstract An equal-collision-probability-curve (ECPC) method is developed in this paper to address the problem of safe spacecraft proximity maneuvers. Considering the uncertainties' influence, the ECPC, which represents the curve of equal-collision-probability-points in the space around the target spacecraft, is firstly established. It is optimal to maneuver along the gradient direction of the ECPC, which is the fastest change in the ECPC. To calculate this direction, a novel auxiliary function, which has the same gradient direction as the collision probability function, is proposed. Compared to traditional collision probability functions, the proposed function does not contain transcendental elements and hence the computational burden can be greatly decreased while maintaining the necessary accuracy. Then, the safe close-range proximity maneuver generated by ECPC method can be implemented along the estimated gradient direction. Analytical validation is performed to assess the use of such collision avoidance scheme for safety critical operations. Furthermore, an improved Linear Quadratic Regulator (LQR) is designed to track the reference trajectory and a Lyapunov-based analysis verifies the stability of the overall closed-loop system. Numerical simulations show that the novel ECPC method is more computationally efficient than traditional methods while maintaining the same accuracy. Moreover, the novel scheme can be easily validated to guarantee the safety of the mission. [ABSTRACT FROM AUTHOR]

Abstract This paper investigates the multi-objective attitude tracking problem of a flexible spacecraft in the presence of disturbances, parameter uncertainties and imprecise collocation of sensors and actuators. An integrated robust H∞ controller, including an output feedback component and a feedforward component, is proposed, and its gains are calculated by solving Linear Matrix Inequalities. The output feedback component stabilizes the integrated control system while the feedforward component can drive the attitude motion to track the desired angles. The system robustness against disturbances, parameter uncertainties and imprecise collocation is addressed by the H∞ approach and convex optimization. Numerical simulations are finally provided to assess the performance of the proposed controller. Highlights • The attitude tracking control of flexible spacecraft is a multi-objective problem. • An integrated robust H∞ controller is proposed to perform attitude tracking. • The proposed controller is robust to imprecise collocation of sensors and actuators. • The system robustness is addressed by the H∞ approach and convex optimization. [ABSTRACT FROM AUTHOR]

The robust optimal sun-pointing control strategy for a large geostationary solar power satellite (SPS) is addressed in this paper. The SPS is considered as a huge rigid body, and the sun-pointing dynamics are firstly proposed in the state space representation. The perturbation effects caused by gravity gradient, solar radiation pressure and microwave reaction are investigated. To perform sun-pointing maneuvers, a periodically time-varying robust optimal LQR controller is designed to assess the pointing accuracy and the control inputs. It should be noted that, to reduce the pointing errors, the disturbance rejection technique is combined into the proposed LQR controller. A recursive algorithm is then proposed to solve the optimal LQR control gain. Simulation results are finally provided to illustrate the performance of the proposed closed-loop system. [ABSTRACT FROM AUTHOR]

Abstract: This paper proposes a way to model uncertainties and to introduce them explicitly in the design process of a preliminary space mission. Traditionally, a system margin approach is used in order to take them into account. In this paper, Evidence Theory is proposed to crystallise the inherent uncertainties. The design process is then formulated as an optimisation under uncertainties (OUU). Three techniques are proposed to solve the OUU problem: (a) an evolutionary multi-objective approach, (b) a step technique consisting of maximising the belief for different levels of performance, and (c) a clustering method that firstly identifies feasible regions. The three methods are applied to the BepiColombo mission and their effectiveness at solving the OUU problem are compared. [Copyright &y& Elsevier]

The distributed adaptive vibration control for solar power satellite (SPS) during on-orbit assembly is investigated in this paper. Focusing on the platform configuration SPS, the different control units (CUs) and relationship matrices, that describe the topology of whole SPS structure, are firstly defined, and the dynamic model of the CU is proposed through the relationship between the CU and the whole SPS structure. Using the linear quadratic optimal control technique, the adaptive controller of the CU is designed to suppress vibrations, and a compensatory component is included in the controller to deal with the interactions with the whole structure. The closed-loop distributed control system is then developed in the presence of the impact at the specific assembling location, which needs to adaptively update along with the varying dynamic model of the SPS structure during on-orbit assembly. The asymptotic stability of the closed-loop distributed system is investigated, and two representative numerical cases are finally provided. The results demonstrate that the proposed distributed adaptive controllers can significantly suppress the vibration caused by the impact among SPS modules during on-orbit assembly. [ABSTRACT FROM AUTHOR]