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The development of telescopes with large apertures is driven by the increasing demand for higher imaging resolution and coverage. Large membrane diffraction space telescopes are particularly attractive due to their lightweight nature, ease of folding and unfolding, and high-precision imaging capabilities. However, their substantial size and flexibility make them susceptible to long-duration, low-frequency vibrations during attitude maneuvers, which degrade attitude and imaging accuracy and risk instrument damage. Consequently, an urgent need exists for a high-precision integrated control scheme. In this paper, the integrated control of vibration and attitude in a large space telescope is studied using cable actuators and control moment gyroscopes (CMGs). Since cables can only withstand limited tension, the unilateral and constrained characteristics of the control input are taken into account during the control design process. Firstly, a rigid-flexible coupling dynamic model of the space telescope is established using the velocity variation principle with hybrid coordinates. Additionally, a two-body astrodynamic model is developed to assess the impact of gravity gradient torque. Next, combining the computational torque method and H∞ control theory, an integrated control scheme is proposed, which achieves vibration and attitude control during maneuvers. Then, this paper optimizes the cable actuator positions via the discrete particle swarm optimization (PSO) algorithm and plans various attitude maneuver paths. Finally, numerical simulations are adopted to demonstrate the effectiveness of the control scheme. The results show that this integrated control scheme swiftly suppresses structural vibrations during attitude maneuvers, enabling high-precision attitude control of the space telescope. • A vibration-attitude integrated control for a novel large membrane diffraction space telescope is studied • 3 types of attitude maneuver paths are planned for the two requirements of time optimal and minimum vibration excitation. [ABSTRACT FROM AUTHOR]

Surface movement exploration techniques are an important approach to further the understanding of small celestial bodies. Hopping movement is employed to take advantage of the weak gravitational field of small celestial bodies, facilitating long-distance mobility of the rover. The rover exhibits a strong attitude-orbit coupling characteristic during the hopping movement. Different contact attitudes contribute to the challenge of determining the rebound trajectory, rendering it difficult for the rover to maintain continuous movement towards the target position along the nominal trajectory. In this paper, a dual-stage control method for continuous surface hopping movement is developed, wherein the control is applied separately during the flight stage and the collision stage. During the flight stage, applying attitude control enables the rover to establish contact with the surface in a specific attitude, ensuring accuracy in movement direction and creating favorable conditions for the control process in the collision stage. The flywheel is controlled in the collision stage to correct the state change and energy loss due to the collision, enabling the rover to attain the desired hopping state. Considering the energy consumption of the rover, the minimum hopping velocity sequence and angular velocity sequence are designed based on the hopping angle and heading angle. The intricacies and uncertainties inherent in the surface environment of small celestial bodies pose challenges for the rover during hopping and flying. In response to the issue, a finite-time adaptive sliding-mode control law is devised. The control law facilitates the rapid adjustment of the rover to the target attitude in flight, demonstrating strong robustness. In addition, by integrating the changes in contact point velocity with the impulse contact model, the accuracy of the rover's rebound state updates is enhanced. Finally, a set of simulations is performed, and simulation results indicate that the dual-stage control method can achieve a final landing position error of less than 0.2 m for a continuously hopping. The finite-time adaptive sliding-mode control law can stabilize the rover's attitude within 10 s. Conducting 500 Monte Carlo simulations of the continuously hopping rover in three small celestial body surface simulation environments shows that the position endpoint error is within 0.3 m. • An innovative control method for continuous hopping on the surface of small celestial bodies with low gravity is proposed. • The attitude-orbit coupling challenge is addressed through a staged control. • A sequence of hopping states with minimum velocities is devised. • A finite-time adaptive sliding-mode control law is designed to effectively address unknown disturbances. • A corrected sliding-mode control law is developed for precise hopping state updates. [ABSTRACT FROM AUTHOR]

Laser ablation technology enables contactless debris removal, providing the advantages of safety and cost efficiency. To deorbit space debris using laser ablation, the attitude motion of the debris must be controlled in advance. Although a previous study derived an attitude control procedure, it empirically determines control gains and is easily affected by assumed preconditions. To solve this problem, this paper proposes an improved attitude control law using laser ablation. Previous methods use feedback control based on the angular velocity error and error quaternions to obtain a reference torque. By analogy with a magnetic attitude control law, this study designs a reference torque based on the angular momentum errors and presents theoretical condition for designing the control gains. Numerical examples using the proposed control law are conducted to control the target attitude from a random initial rotation of 1 rpm to an arbitrary attitude, which verify the effectiveness of the proposed method. Furthermore, the position, direction, and magnitude uncertainties of laser irradiation are introduced to numerically examine their effects on the control accuracy under environmental disturbances. • Conventional attitude control method by laser ablation is improved. • The effective control gains can be set theoretically by applying the concept of magnetic attitude controller. • The designed control gains depend on the rate of target rotation. • The method is effective under laser uncertainties and environmental disturbances. [ABSTRACT FROM AUTHOR]

This paper proposes an integrated architecture for navigation and attitude control of low-cost suborbital launch vehicles, propelled by a solid motor. Single-nozzle, two Degrees-of-Freedom (DoF) Thrust Vector Control (TVC) actuation is adopted. For architecture design purposes, a non-linear, unstable, 6 DoF model for the generic thrust-vector-controlled launcher dynamics and kinematics is deduced, and a linear state-space representation is proposed. The navigation system provides full-state estimates resorting to novel complementary kinematic filters, whose design allows to establish an explicit relation with steady-state Kalman filtering. A globally stable estimation solution is obtained, apart from the singularities arising from the use of Euler angles. The attitude control law is derived from the Linear Quadratic Regulator (LQR) using the state-space models for each linearization point of the reference trajectory, with an integral action (LQI) added to improve robustness and to provide null steady state attitude tracking error. A correction method is proposed to allow for pitch and yaw control in the presence of spinning motion, precluding the need of a supplementary roll control system. The control system is implemented through gain scheduling, resorting to an altitude-based linear parametric varying method. The architecture is implemented in a realistic simulation environment, composed by the 6 DoF non-linear model, the Navigation and Control solutions, and the environmental disturbances, to assess its performance through Monte Carlo simulations. The navigation system is able to provide accurate estimates of the state vector, while the control system satisfies attitude tracking performance and robustness to both external disturbances and model parametric uncertainties. • An integrated architecture for navigation and attitude control of suborbital launch vehicles is proposed. • An original time-varying state-space representation is deduced. • Novel complementary kinematic filters are designed to provide accurate full-state estimation. • LQI-based control allows for accurate pitch and yaw reference tracking via thrust vectoring. • The architecture performance is assessed through extensive Monte Carlo simulations. [ABSTRACT FROM AUTHOR]

The paper proposes algorithm for angular motion control of a dynamically elongated spacecraft. The algorithm is based on the direct Lyapunov method using matrix control coefficients. The calculated mechanical torque is implemented using magnetorquers. Control parameters are selected using Floquet theory to ensure convergence to the required motion. • An algorithm for attitude control of a dynamically elongated spacecraft is proposed. • The algorithm is based on the direct Lyapunov method using matrix control gains. • A methodology for matrix gain selection is based on Floquet theory. • The algorithm performance for 3U CubeSat three-axis stabilization is demonstrated. [ABSTRACT FROM AUTHOR]

Satellites in very low Earth orbits can leverage the aerodynamics forces for control purposes. Three-axis attitude control can be achieved by adjusting the orientation of aerodynamic surfaces, benefiting from both drag and lift forces. However, complete control over the orbital motion of a satellite is unattainable using aerodynamic forces alone, as positive thrust cannot be generated. Nonetheless, this limitation does not apply when focusing on the control of a satellite formation. By properly modulating the orientation of panels, positive and negative relative forces can be generated. This paper shows that even in a formation of simple microsatellites, it is possible to achieve formation and attitude simultaneous control by using no other actuators than the aerodynamics surfaces. • Very Low Earth orbits are appealing but the large atmospheric density can be a problem. • The atmospheric forces can be used for control purposes. • Formation control and attitude control can be individually realized by using aerodynamic forces. • A GNC architecture is implemented to realized both goals with the same aerodynamic actuators. • Performance in keeping or acquiring a desired state (attitude and relative position) is promising. [ABSTRACT FROM AUTHOR]

Liquid crystal diffractive sails can generate controllable tangential radiation pressures, which have advantages in the Solar Polar Imager mission (SPI). This paper proposes a composite diffractive sail design consisting of an ideal passive diffraction grating sail and liquid crystal polarization gratings (LCPGs). The main diffractive grating sail generates thrust for orbital transfer with a sun-facing attitude. An optimal 3.8-year transfer trajectory is obtained by an improved direct method. Small-area LCPGs distributed around the edge are used for three-axis attitude control. The LCPGs can continuously modulate the force and torque by controlling the applied voltage based on the electro-optical characteristic of liquid crystal materials. A control strategy, including an adaptive attitude-tracking PD controller and a control allocation method, is designed to track the attitude profile generated by the transfer trajectory. Finally, the attitude-orbit coupling dynamics are used to simulate the whole process, and the results show that orbital insertion errors are minor. • A combined diffractive solar sail model using liquid crystal polarization gratings. • Independent and controllable three-axis attitude control of diffractive solar sails. • Optimizing transfer trajectory for solar polar mission based on the proposed sail. • An adaptive attitude-tracking PD controller and a control allocation method. [ABSTRACT FROM AUTHOR]

This paper addresses the fault-tolerant attitude tracking control problem for spacecraft with limited communication capability and input saturation. The design of a hysteresis quantizer aims to alleviate communication burden between the controller module and the actuator module. Then, the attitude tracking problem for spacecraft with input quantization, input saturation, actuator faults and external disturbances is transformed into an attitude control problem with uncertain input coefficients and bounded disturbances. Thereafter a dynamic loop gain function-based approach and a hyperbolic tangent function term with a time-varying boundary layer are introduced to address the uncertainties of input coefficients and the bounded disturbances, respectively. To facilitate the boundness analysis of the signals in the closed-loop system, a pertinent lemma about the dynamic loop gain function is proved. Finally, numerical simulations are employed to validate the effectiveness of the proposed scheme. • The spacecraft attitude tracking problem with disturbances and uncertainties. • The quantized control law alleviates communication burdens of spacecraft. • A dynamic loop gain function method is used to address the impact of uncertain. • A hyperbolic tangent term is devised to counteract bounded disturbances. [ABSTRACT FROM AUTHOR]

Modern space missions for Earth Observation (EO) purposes often rely on satellites equipped with very large flexible appendages, such as antennas and solar panels, which are demanded to perform agile slew manoeuvres. In most cases, the elasticity of such systems cannot be neglected in the design of the attitude controller, as excessive elastic displacements of the structural elements may compromise their stability and pointing performance. In this scenario, this paper aims at proposing an automated nested optimization framework to simultaneously optimize spacecraft structural and control dynamics, to be applicable to a wide range of flexible spacecraft. The objective of such a co-design architecture is to modify design parameters, at both structural and control levels, to minimize the mass of the spacecraft while maximizing its agility and satisfying imposed requirements. Moreover, as robust multivariable techniques have become more and more applied to ensure satisfactory robust performance margins, this paper's goal is to pose a multi-channel structured H ∞ control architecture in the co-design problem. To guarantee the generality of implementation, a structural design tool (MSC Nastran) is interfaced with a coding environment (Matlab/Simulink) to set-up an autonomous exchange of information between structural and control domains. Starting from an initial definition of the spacecraft material, geometry and control requirements (in terms of loop-shaping transfer functions), relevant parameters are extracted from the structural tool and a linearized dynamic model assembled. Then, a controller is synthetized based on the provided requirements, followed by a verification on the nonlinear plant of the satellite. The procedure is repeated until the stop criteria (based on tolerance and max iterations) is satisfied. Finally, the output of the proposed architecture is obtained as an optimized structural model and robust controller tailored for the satellite dynamics. • The problem of structural/control co-design of flexible spacecraft is addressed. • A general framework interfacing a commercial finite element code and a coding environment is proposed. • The architecture is posed as a nested robust optimization. • A case study composed of a satellite with two flexible solar panels is proposed. • The framework is tested on the satellite non linear model. [ABSTRACT FROM AUTHOR]

This paper investigates the effects of model fidelity and parameter uncertainty on the performance of a hybrid model-based feedback-feedforward control scheme for attitude tracking of a satellite with flexible appendages. The feedforward component is an inverse model-based term produced through a computational approach known as inverse simulation (InvSim), which works by iteratively solving a discretised reference trajectory. The hybrid controller's feedback is proportional-derivative (PD) based, using body attitude and rate feedback to provide stability and robustness. Furthermore, to ensure that the flexible modes do not trigger instability, the PD control gains are tuned to give a closed-loop response that is significantly slower than the flexible modes. Additionally, excitation of the flexible modes is reduced by minimising jerk through polynomial rest-to-rest manoeuvres, following the shortest quaternion path using spherical–linear-interpolation (SLERP). The effects of the appendage flexing on attitude tracking are then compensated through the feedforward element of the hybrid controller, with performance being compared to a traditional PD tracking law. The effect of the model fidelity on the performance of the hybrid controller is investigated through the use of both rigid body and multiple-fidelity finite-element mathematical models. Additionally, the effect of uncertainties in the model parameters is investigated to determine the accuracy of the model required to obtain significant improvement in attitude tracking. It is found that in the absence of any model parameter uncertainty, the hybrid controller outperforms the PD tracking control law by at least one order of magnitude when the finite-element model is used. Increasing the number of finite elements was found to provide no significant improvement in performance, with one element being sufficient and favourable with its lower computational overhead. It was also found that to ensure good performance compared to the PD tracking controller, the uncertainty in the inertia tensor should be < 1%. Similarly, uncertainty in the first flexible modal frequency should be < 0.5 rad/s. • Inverse simulation (InvSim) used to produce model-based feedforward control. • Combined with PD feedback to improve attitude tracking. • Effects of model uncertainty and fidelity on model-based control investigated. [ABSTRACT FROM AUTHOR]

An active debris removal method using a laser is a promising technology for its advantage in contactless operations. This paper deals with the attitude control of an uncooperative target by a laser, which is an important phase before deorbiting. The difficulty of attitude control by the laser stems from the torque directional constraint because the laser generates thrust along the normal vector of the irradiated face irrespective of the irradiating direction. Thus, the control torque along the normal vector cannot be generated, which makes the attitude control with the laser torque challenging. To tackle this problem, this paper first designs a reference controller that assumes arbitrary control torques are available. Then, a method for determining the irradiating point is proposed so that the difference between the reference torque and the actual one is minimized. Although the proposed controller does not guarantee theoretical convergence to the desired attitude, the effectiveness of the proposed controller is numerically verified for a box-type object. Furthermore, the robustness to the uncertainties of thrust magnitude and direction is also examined by Monte Carlo simulations. • An attitude control method by laser ablation is proposed. • The method determines the irradiating point with respect to reference torque. • The method is effective under thrust magnitude uncertainty. • The method is effective under thrust directional uncertainty. [ABSTRACT FROM AUTHOR]

OrigamiSats, a new concept in solar sailing, are origami spacecraft with reflective panels that, when flat, operate as a conventional solar sail. Shape reconfiguration, i.e. "folding" of the origami design, allows the OrigamiSat to change operational modes, performing different functions as per mission requirements. For example, a flat OrigamiSat could be reconfigured into the shape of a parabolic reflector, before returning to the flat configuration when required to operate again as a solar sail, providing propellant-free propulsion. The attitude dynamics and shape reconfiguration of OrigamiSats are known to be highly coupled, thus presenting a challenge from a control perspective. This paper investigates the problem of integrating attitude and shape control of a Miura-fold pattern OrigamiSat through the use of variable reflectivity, allowing differences in solar radiation pressure to be used to enact shape reconfiguration and attitude manoeuvres. A closed-loop feedback controller is presented which combines and balances the attitude and shape control requirements, and gain-scheduling is implemented to address some specific features of the system dynamics. Numerical simulations of the multibody dynamics of the system are used to test the proposed controller and simulations of some example manoeuvres are performed which demonstrate the system's performance. • OrigamiSats are a new concept in solar sailing. • Solar radiation pressure can be used to "fold" an origami solar sail. • An integrated attitude and shape control law is developed. • Gain-scheduling is used to address some nonlinearities of the system. • Attitude and shape control is demonstrated through numerical simulations. [ABSTRACT FROM AUTHOR]

Attitude control is an essential operation for spacecraft to perform space missions, and control strategies play a decisive role. In this paper, a second-order quaternion system of rigid spacecraft attitude dynamics is derived based on the fully actuated system approach. The quaternion system has two constraints that lead it to be a fully actuated system. Two closed-loop control strategies are proposed to achieve attitude tracking and stabilization for the fully actuated system. The stabilization strategy can control the scalar quaternion to reach the desired value and be used to deal with the unwinding phenomenon. Numerical simulations demonstrate the second-order quaternion system's validity and the control strategies' effectiveness. • A second-order quaternion system is derived from an intuitive perspective. • The control laws are much easier to design than the first-order system. • Arbitrary three quaternions can be employed as the controlled variables. • The unwinding phenomenon is handled in a simple way. [ABSTRACT FROM AUTHOR]

In 2017, Technische Universität Berlin (TU Berlin) has launched the first space-rated fluid-dynamic actuator (FDA) on the TechnoSat mission. This type of actuator is unique with respect to its dynamic properties, as it offers a very high torque but limited angular momentum storage capabilities in comparison to reaction wheels (RWs). In combination with the tetrahedron RW assembly featured on the spacecraft, a unique combination of attitude control actuators has been formed. The spacecraft axis parallel to the FDA is over-actuated, featuring control actuators of contrasting nature which offers novel options in terms of spacecraft operations. However, with the camera axis on TechnoSat being aligned parallel to the actuator's axis of rotation, there is limited operational use to this setup. In July 2019, TU Berlin launched Berlin Experimental and Educational Satellite 9 (BEESAT-9) – the first CubeSat featuring a set of three orthogonal RWs and a single picosatellite fluid-dynamic actuator (pFDA) for attitude control. The camera axis of BEESAT-9 is perpendicular to the actuator axis, and therefore allows the demonstration of the proposed novel modes of operation. Among the proposed modes is an artificial increase of the swath and the acquisition of stereo images of multiple, successive targets on the ground. Crucial to these operation scenarios is the control allocation between the RW assembly and the single pFDA. This paper first introduces the spacecraft and describes the proposed novel modes of operation that are enabled by the actuator combination realized on BEESAT-9, followed by a review of the state of the art of control allocation for over-actuated spacecraft. The main part of the paper comprises a description of the implementation of the dynamics models of the attitude control actuators and spacecraft, followed by a description of the implemented high-level control laws and control allocation methods. Finally, simulation results for the artificial swath increase and the single spacecraft stereo imaging mode are appended and discussed. • Picosatellite fluid-dynamic actuators enable novel modes of satellite operation. • Control allocation enables combination of novel actuator and reaction wheels. • First CubeSat with fluid-dynamic actuator and reaction wheels launched in 2019. • Attitude maneuvers in the orbit plane enable stereo imaging from single spacecraft. • Attitude maneuvers out off the orbit plane enable artificial swath increase. [ABSTRACT FROM AUTHOR]

In this paper, a comparison between optimal adaptive and fuzzy control methods is performed for the problem of satellite attitude control. The optimal adaptive controller is designed based on Markov parameters, which can be utilized to follow a given control command. The fuzzy controller is shown to achieve faster response than the optimal adaptive controller. Moreover, the problem of satellite attitude control using the adaptive controller with on–off thrusters is studied. Simulation results affirm that the adaptive control method with on–off thrusters is robust against the variations in the inertia matrix and that the optimal adaptive control method is robust against constant disturbances. • Optimal adaptive and fuzzy control methods are compared for the satellite attitude control problem. • The fuzzy controller is shown to achieve faster response than the optimal adaptive controller. • The optimal adaptive control method with on–off thrusters is shown to be robust against constant disturbances and variations in the inertia matrix. [ABSTRACT FROM AUTHOR]

In this paper, the attitude control of spacecraft using a pyramid-type variable-speed control moment gyro (VSCMG) system is considered. For successive attitude maneuvers, it is desirable that the wheel angular momentums are regulated to nominal values after an attitude maneuver. In addition, avoidance of reference aligned singularity (RAS) is also important. RAS is a singularity point where the singular direction of the constant-speed control moment gyro (CSCMG) system is aligned to a reference torque. Regarding the motivations, this paper proposes a steering law which achieves attitude control and regulates the wheel angular momentums while successfully avoiding RAS. Both numerical simulations and experimental verification are performed to demonstrate not only the validity of the proposed method, but also a certain robustness against unmodeled dynamics and the gravity torque. • Spacecraft attitude is controlled by variable speed control moment gyros. • Reference aligned singularities of CMGs are particularly focused in the control. • Wheel angular momentums are usually regulated to nominal values. • Effectiveness of the control law is validated in all maneuver directions. • Ground experiments are conducted in order to examine the proposed control. [ABSTRACT FROM AUTHOR]

This paper presents an integrated control method for an extremely large spacecraft, such as the space interferometry telescope, to achieve control task with high accuracy and high stability. The flexibility of the spacecraft bus would be more obvious when their size gets larger, so the vibration reduction for this kind of spacecraft would be more challenging. This paper presented a dynamic model of the spacecraft with flexible central body, and with flexible appendages and CMG (Control Moment Gyroscope) with disturbances at the same time. Then the integrated control method is presented to make the relative attitude of the two payloads on the spacecraft bus track with each other. The vibration isolation techniques are adopted to reduce the high frequency vibrations, and the nonlinear feedback control is presented to achieve the relative attitude convergence to 0 with high accuracy and high stability. Finally numerical simulations are carried out to verify the performance of the integrated control method. • Dynamic model which include high and low frequency disturbances is presented. • An integrated control method for space interferometry telescope is presented. • Passive and active vibration isolation platforms are used. • Nonlinear controllers are presented to suit the vibration isolation platforms. [ABSTRACT FROM AUTHOR]

Formations of multiple satellites observe target areas on the Earth's surface from different perspectives, enabling construction of 3D surface maps by photogrammetric methods. The main objective of TOM (T elematics Earth O bservation M ission) is to monitor ash clouds during Volcanic Eruptions with a satellite formation consisting of three CubeSats. TOM is Bavaria's contribution to the multi-national T elematics I nternational M ission (TIM), composed of 7 satellites by partners from 4 continents. To achieve sufficient overlap of the images for the usage of photogrammetric methods, very precise and coordinated attitude control of the satellites is necessary. This paper presents a vision-based attitude control approach using image data to calculate the control input, also referred to as visual servoing. Thus, the cameras and data processing capacity onboard the CubeSats will be used for both image acquisition and attitude determination, eliminating the need for additional precise attitude sensors like star trackers. To perform visual servoing, highly distinctive features are identified and tracked in the images. The control input for reaction wheels is calculated from the movement of the features between consecutive images. To ensure that the images captured by the different satellites overlap sufficiently for photogrammetric processing, the satellites track the same target area based on feature data exchange directly between the satellites. Simulations of realistic scenarios based on the TOM project show the applicability of the presented attitude control approach. In a first step, a software-based Earth observation simulator generates the acquired images, demonstrating feasibility to compute and track features in satellite images to compute the AOCS (Attitude and Orbit Control System) input. In the second step, hardware-in-the-loop (HiL) simulations allow to realistically characterize the performance of this approach. This HiL environment utilizes two high-precision/high-dynamics turntables as 3-axis attitude simulators, where the satellites with their cameras are mounted, while the observation target is mounted on a mobile robot to simulate the relative orbit motion. • TOM is a CubeSat formation for the joint observation of volcanic ash clouds. • Observation from different perspectives enables the generation of 3D surface maps. • Visual servoing provides the required precise and coordinated attitude control. • The features to be tracked cooperatively are exchanged via inter-satellite link. • Software and hardware-in-the-loop simulations show the feasibility of the approach. [ABSTRACT FROM AUTHOR]

The commented paper discusses the attitude dynamics representation using quaternions. The main idea of the paper is to eliminate angular velocity from the dynamics equations. The obtained equations are interesting. However, all applications presented in the discussed paper have plain errors in the equations derivations. This affects the main results and conclusions of the paper, which require significant corrections. • Errors in commented paper are outlined. • Quaternion dynamics equations application for the attitude control is discussed. • Control singularity is shown. [ABSTRACT FROM AUTHOR]

To meet demanding mission objectives, most current satellites for Earth and Universe observation are equipped with large flexible appendages. However, due to the coupling between spacecraft rigid and elastic dynamics, unwanted elastic vibrations of such flexible elements may compromise the achievement of pointing and stability requirements. Therefore, control solutions aiming at avoiding instabilities and, at worst, the failure of the mission, are currently needed. In this scenario, growing interest has been recently devoted to Active Vibration Control (AVC) strategies relying on the use of smart materials. Among them, piezoelectric patches are the most studied and tested devices for both actuating and sensing purposes. This paper aims to contribute to this line of research by investigating a different type of actuator, namely an Offset Piezoelectric Stack Actuator (OPSA), to assess its performance in damping out elastic vibrations of large space structures. Moreover, special attention is devoted to compare OPSA performance with standard patch actuators behaviour. In order to develop the AVC system, an equivalent electro-mechanical coupled Finite Element (FE) formulation is implemented, integrating both sensors and piezo-stack elements on the passive hosting structure. The final structural model including both electrical inputs/outputs, as well as modified mass and stiffness due to the additional piezo devices, is then obtained. A parametric analysis is carried out to optimize the maximum control action exerted by the OPSA device. Finally, the OPSA numerical model is experimentally validated by mounting it on a cantilever plate, representative of a scaled solar panel. Experimental data are used not only to verify the device expected functioning, but also to tune the parameters in the OPSA FE model, which can be finally integrated in a planar satellite dynamics simulator to evaluate the AVC system efficiency when performing attitude manoeuvres. • Active Vibration Control of large flexible space structures. • Use of offset piezoelectric stack actuators. • Finite Element formulation implemented, integrating both sensors and piezo-stack elements. • OPSA numerical model validation on a cantilever plate, representative of a scaled solar panel. • Finite element tuning using experimental data. [ABSTRACT FROM AUTHOR]

Particle swarm optimization (PSO) is a non-gradient biologically inspired global optimization method. It utilizes direct calculation of the optimized cost function during the numerical simulation of the process in question. Significantly complex and implicit functions that prohibit clear gradient expression are effectively handled by PSO. This feature is utilized in the paper to construct specific angular trajectory of a magnetically actuated satellite. A cost function is suggested that confines the control torque corresponding to this trajectory to the plane orthogonal to the geomagnetic induction vector. Optimal trajectory coefficients are found by PSO. The control structure is based on the Lyapunov feedback approach and explicitly accommodates the gravitational and aerodynamic torques influence. PSO procedure provides the optimal control gains on this stage. Stability of the resulting motion is demonstrated. Numerical simulation example with different disturbance factors illustrates the expected stabilization accuracy of the order of few degrees. • Attitude trajectory is constructed for a magnetically actuated satellite. • Particle swarm optimization is used to find the trajectory and minimize torque along the induction vector. • Feedback control structure is adopted with gravitational and aerodynamic disturbances mitigation. • Stability of the trajectory maintenance is demonstrated. • Simulation example illustrates stabilization accuracy of few degrees. [ABSTRACT FROM AUTHOR]

An electric solar wind sail (electric sail or E-sail) is a new type of propulsion system that does not require propellant. This paper proposes a nonlinear model predictive control law based on deep learning to control the attitude of a barbell electric sail. A barbell E-sail is composed of two tip satellites connected via long conductive tethers to a central insulated confluence point. The two tethers are insulated from each other at the confluence point such that the voltages of the two tethers can be independently controlled. The attitude dynamics of the system is modelled using a nonsingular description of tether orientation. To reduce online computation loads, the proposed control scheme has a two-stage design, namely, an offline stage and an online stage. In the offline stage, a large amount of data is generated using the nonlinear model predictive control law and then taken as a dataset to train a deep neural network. In the online stage, feedback control of the system attitude is achieved with extremely low computational cost by performing real-time mapping from the system state to the control output using the trained deep neural network. Finally, the efficacy and performance of the proposed deep learning-based control law are demonstrated via numerical case studies. • Attitude dynamics of an E-sail system is modelled by a nonsingular formulation. • Nonlinear dynamics and inherent constraints of the system are accounted for. • Extremely low computational cost of feedback control is achieved using a DNN. [ABSTRACT FROM AUTHOR]

Abstract To meet the spacecraft attitude control requirements with high accuracy and stability, all vibrations in the spacecraft should be reduced in appropriate ways. This paper presents an integrated control method for attitude and the vibrations in both high frequency and low frequency in the spacecraft. The integrated control method includes a vibration isolation platform and a modified adaptive attitude control method. The paper presents a vibration isolation platform with magnetic suspension to reduce high frequency vibrations and a parameter design method for the platform. An adaptive control method is presented to reduce low frequency vibrations while accounting for the bandwidth constraint due to the vibration isolation platform. Firstly, a parameter design method is proposed for the vibration isolation platform, and an entire 6 × 12 dimensional transformation matrix is derived for the case that the inertia of the payload is of the same order of magnitude as that of spacecraft bus. Then, an adaptive attitude controller is presented that accounts for the coupling characteristics of the spacecraft, the vibration isolation platform and the uncertain or unknown payload inertia parameters. To ensure the robustness of the attitude control system and the performance of the vibration isolation system, a method of estimating the initial value of the payload inertia is presented using classical control theory. Finally, numerical simulations demonstrate that the integrated control method presented in this paper can achieve the attitude control task for spacecraft with high accuracy and stability. Highlights • An integrated jitter control method for both high and low frequencies is presented. • The entire 6 × 12 dimensional transformation matrix of the platform is derived. • An adaptive attitude controller with inertia estimation is presented. • Parameter design method is proposed for adaptive attitude controller. [ABSTRACT FROM AUTHOR]

Solar sails can realize propellant-free attitude control if they have the ability to change their shape by moving reflective surfaces to harness the force of incident solar radiation pressure (SRP). In this paper, we formulate a stability analysis of the attitude motion of a sail spacecraft with large deformation under SRP, and propose stable attitude control methods using variable-shape mechanisms. The basic principle is that the attitude is maintained at a stable equilibrium point that suppresses the disturbance torque due to SRP and prevents disturbance angular momentum from accumulating in the reaction wheels. The equilibrium point can be shifted by moving reflective surfaces, and the attitude follows the new equilibrium point by reaction wheel control. The stability of the equilibrium points can be quickly evaluated by a stability analysis based on linearized equations of motion, and one can choose any equilibrium point by setting the reflective surfaces to a corresponding configuration. This control scheme enables the spacecraft body to be pointed towards desired directions, and even if the attitude maneuver and maintenance accumulate angular momentum in the reaction wheels, this momentum can also be unloaded without propellant by using SRP and shape variation. The dynamics formulation aims at a zero-momentum system that is generally applicable to arbitrarily shaped spacecraft, and the formulation of the stability analysis and control methods were verified by numerical simulations. • Variable-shape mechanisms realize propellant-free attitude control of solar sails. • Stability analysis is formulated for attitude of solar sails with arbitrary shape. • Relationship among the shapes, equilibrium points, and stability is investigated. • Strategies of attitude maneuver, maintenance, and unloading are proposed. [ABSTRACT FROM AUTHOR]

Modern small satellite constellations are demanding quick spacecraft commissioning and high duty cycle for their payload operations. In a conventional small satellite mission architecture, operations time is often split between different modes that require different spacecraft attitudes to achieve their goal. For example, drag-based orbit maneuvering can be used for creating relative separation between constellation members, and with spacecraft consisting of single rigid bodies, dedicated attitude manipulation is required to perform these operations. The attitudes that take advantage of the body asymmetries to modify the drag area are generally incompatible with the attitude maneuvering required for payload operations. At the constellation scale, these downtimes in the payload operation schedule can greatly reduce the overall capability of the system. By including deployable, articulating solar arrays in the design of small spacecraft, array pointing can be decoupled from the main payload pointing operations. With these pieces decoupled, payload operations can proceed uninterrupted while the articulating arrays enable power generation or drag-based orbit maneuvering. The focus of this paper is the demonstration of the articulating arrays as a means of enabling drag-based orbit phasing, while agile attitude maneuvering for payload operations proceeds concurrently. The dynamic equations of the multibody system are derived, and guidance, navigation, and control considerations for achieving decoupled attitude and articulation objectives are presented. Results are shown for attitude performance as well as for the orbit maneuvering performance achieved through array articulation. • Drag-based orbit phasing is achieved with decoupled attitude maneuvering. • Concurrent operations streamline the commissioning of satellite constellations. • Orbit phasing of 30 degrees is achieved in 120 days in a sample mission. • Payload-driven attitude maneuvering is minimally impacted. [ABSTRACT FROM AUTHOR]

Satellite-on-a-printed circuit board (PCBSat) is a kind of femto-satellite that integrates all the electronic devices onto a PCB. Constrained by the energy/volume capacity, traditional attitude actuators no longer suit the PCBSat. In the paper, employing the flexible PCB material and micro-steppers, a new design that enables the PCBSat to fold and unfold is proposed, and a prototype is manufactured. The foldable PCBSat could achieve attitude maneuver by the spatial multibody dynamical effect. Meanwhile, the copper coils embedded in the boards can form a 3D magnetorquer when the PCBSat is at the orthogonal-fold status, making magnetic de-tumbling possible. Related theories combined with the detailed design of the PCBSat are introduced, and simulations are conducted to demonstrate the effectiveness of the attitude control mechanism. • A foldable PCBSat enabling attitude control is designed and a prototype is manufactured. • Embedded coils in the boards could form a 3D magnetorquer when the PCBSat folds properly. • Attitude maneuver could be achieved by alternative turns of the boards. • Dynamic simulations are conducted to demonstrate the effectiveness of the attitude control mechanism. [ABSTRACT FROM AUTHOR]

This paper addresses the attitude reorientation problem under the pointing-constraint zone, angular velocity limits, and actuator constraints. The ERG-based control scheme is designed such that the attitude is stabilized, while the state and control constraints stay in the allowable regions defined by the Lyapunov invariant subset. The disturbance observer-based control law is developed in the inner loop of ERG that enables attitude stabilization in the presence of external disturbance. The navigation layer is utilized to manipulate the reference state to enforcing constraints satisfaction. The effectiveness of the proposed constrained attitude control algorithm is then verified through numerical simulations. • Address the constrained attitude control problem via explicit reference governor. • Design a disturbance observer-based inner loop controller. • Design the Lyapunov thresholds of state and actuator constraints. [ABSTRACT FROM AUTHOR]

The paper focuses on the attitude dynamics and control of a barbell electric sail (E-sail). The barbell E-sail is composed of two tip satellites connected via long conductive tethers to a central insulated confluence point. The two tethers are insulated from each other at the confluence point such that the voltages of the two tethers can be independently controlled. A nonsingular formulation based on the dumbbell assumption is proposed to describe the attitude dynamics of the barbell E-sail, and exploited to develop a nonlinear model predictive controller for attitude adjustment and maintenance of the barbell E-sail by regulating the tether voltages. The nonlinear model predictive control law explicitly accounts for the nonlinear attitude dynamics and mission-related constraints of the barbell E-sail. Finally, the efficiency and performance of the proposed control law are demonstrated via numerical case studies. • A nonsingular formulation of attitude dynamics is proposed to model a barbell E-sail. • The controller design of the barbell E-sail is based on a nonlinear model predictive control formulation. • Attitude adjustment and maintenance are achieved through voltage regulation. [ABSTRACT FROM AUTHOR]

Before an active debris removal process is carried out, the attitude stabilization of the debris needs to be executed to ensure security. Fortunately, for the tethered system, a full despun is not always necessary; a stabilization of the motion of the tether attachment point is usually sufficient. This paper discusses the way to stabilize the motion of the tether attachment point of a captured noncooperative space debris without any auxiliary mechanisms. The debris possesses unknown inertia information and states after capture, while the tether tension is the only input to manipulate its attitude. The direction and value of the tension is regulated based on the continuous dissipation of the rotational energy of the debris. Moreover, the range of the tether direction is restricted to a small conical region, in order to reduce the tether libration motion and the orbit maneuver of the main satellite. The motion of the tether attachment point is desired to be observed instead of measuring the attitude motion of the debris in the feedback loop, which might be easier to attain in practical applications. Numerical simulations are carried out to validate the effectiveness of the control method and to check the performance of the controller with different parameters.] • The motion of the tether attachment point is stabilized by the tether tension. • No inertia information concerning the debris or the measure of its attitude motion is needed. • The range of the tension direction and convergence rate can be adjusted by choosing proper control parameters. • Different equilibrium states are discussed by theoretical analyses. [ABSTRACT FROM AUTHOR]

This paper presents the modeling and robust H ∞ control of a novel ultra-quiet spacecraft which employs the non-contact Stewart platform to actively control its support module (SM) and payload module (PM). The SM and PM are mechanically separated such that the disturbances and vibrations from the SM can be perfectly suppressed. However, this novel spacecraft has internal instability. To overcome the instability, this paper derives the dynamics model of the non-contact Stewart spacecraft using Newton–Euler approach at first. Then, the mixed sensitivity robust H ∞ control is developed to guarantee both the attitude stability and precision pointing. Finally, simulation studies are provided to validate the effectiveness of the proposed approach. [ABSTRACT FROM AUTHOR]

In this paper, a robust control method is proposed for rigid spacecraft attitude stabilization under inertia matrix uncertainties and external disturbances. The attitude stabilization problem is firstly reformulated into an optimal control problem, then a state-dependent indirect Chebyshev pseudospectral (SDICP) method is designed to solve the resultant optimal control problem. It is proved that, by incorporating the uncertainties into the performance index and adaptively estimating the parameters, the stabilization problem can be converted into an infinite horizon optimal control problem. Based on successive state-dependent coefficient parameterization, time domain transformation and Chebyshev spectral discretization, the proposed SDICP method turns the resultant optimal control problem into a sequence of systems of linear equations, which can be efficiently solved by basic matrix operations. Simulation results also verify the effectiveness of the proposed robust adaptive attitude stabilization method. [ABSTRACT FROM AUTHOR]

Attitude and vibration control of a general form of flexible satellites is addressed in this paper. Partial differential dynamic equations are derived considering new details such as multi sectioned solar panels and elastic connections between main hub and solar panels. Boundary control approach is adopted to eliminate simplification errors of discrete models, using just one actuator in the hub. Asymptotic stability of attitude dynamics is proved for a group of boundary controllers and necessary conditions for asymptotic stability of vibrations are discussed. Being independent of modeling accuracy and using easily measurable feedbacks are among advantages of the proposed class of controllers. Through simulations using FEM technique and a controller with the least number of boundary feedback parameters, good performance of the introduced method is illustrated. • A class of boundary controllers is proposed for attitude and vibration control of flexible satellites. • The least number of easily measurable feedback signals from the boundary can be used. • Asymptotic stability conditions are discussed and illustrated via simulations. • Modeling accuracy does not matter. • Multi-sectioned panels and elastic/viscous connections between panels and the hub are considered. [ABSTRACT FROM AUTHOR]

This paper introduces an Optimal Fuzzy Sliding Mode (OFSM) controller for the control of fuel sloshing in a spacecraft during orbital maneuvers. In order to optimize controller's gains for minimizing control inputs, a Particle Swarm Optimization (PSO) algorithm is employed. A novel gain scheduling of the OFSM controller is developed for the coupled spacecraft-fuel system. Fuzzy gains are produced on-line based-on rules that utilize dynamic system variables and the sign of saturation function of sliding surfaces. The fuel slosh is modeled as a fixed spherical mass and two spherical pendulum masses moving in a spherical tank. Computer simulations illustrate the effectiveness of the proposed controller in maneuvering spacecraft while suppressing fuel sloshing. The performance of the OFSM controller is compared to that of the optimum State-Dependent Riccati Equation (SDRE) controller. • An Optimal Fuzzy Sliding Mode (OFSM) controller is designed for the control of fuel sloshing in 3D spacecraft maneuvers. • A Particle Swarm Optimization (PSO) algorithm is employed to optimize control gains. • The performance of the OFSM controller is compared to that of the optimum State-Dependent Riccati Equation (SDRE) controller. [ABSTRACT FROM AUTHOR]

This paper is devoted to developing collocated attitude and vibrations controllers for a square solar sail spacecraft containing four pre-tensioned triangular sails supported by flexible diagonal booms and four reflective control vanes at boom tips. Since the control torques for attitude control are created by the solar radiation pressure forces on the tip vanes, the attitude control problem is effectively non-collocated. The present work identifies collocated sensing for the tip vane forces which furnishes a passive input-output model. Various controllers are developed which furnish attitude and vibrations control. A finite-element-based linear structural model is used to evaluate the developed controllers and examine the controller-structure interactions. • Dynamics and control of solar sails with 2-DOF control tip vanes are considered. • New approach to solve non-collocated attitude control without spillover is addressed. • Various collocated controllers are developed for attitude and vibrations control. • A finite-element-based linear model is used to evaluate the developed controllers. • It's shown that simple controllers add negative damping and destabilize Solar Sails. [ABSTRACT FROM AUTHOR]

Abstract: The design of large space structures (LSS) requires the use of design and analysis tools that include different disciplines. For such a kind of spacecrafts it is in fact mandatory that mechanical design and guidance navigation and control (GNC) design are developed within a common framework. One of the key-points in the development of LSS is related to the dynamic phenomena. These phenomena usually lead to two different interpretations. The former one is related to the overall motion of the spacecraft, i.e., the motion of the centre of gravity and motion around the centre of gravity. The latter one is related to the local motion of the elastic elements that leads to oscillations. These oscillations have in turn a disturbing effect on the motion of the spacecraft. From an engineering perspective, the structural model of flexible spacecrafts is generally obtained via FEM involving thousands of degrees of freedom (DOFs). Many of them are not significant from the attitude control point of view. One of the procedures to reduce the structural DOFs is tied to the modal decomposition technique. In the present paper a technique to develop a control-oriented structural model will be proposed. Starting from a detailed FE model of the spacecraft and using a special modal condensation approach, a continuous model is defined. With this transformation the number of DOFs necessary to study the coupled elastic/rigid dynamic is reduced. The final dynamic model will be suitable for the control design implementation. In order to properly design a satellite controller, it is important to recall that the characteristic parameters of the satellite are uncertain. The effect that uncertainties have on control performance must be investigated. A possible solution is that, after the attitude controller is designed on the nominal model, a Verification and Validation (V&V) process is performed to guarantee a correct functionality under a large number of scenarios. The V&V process can be very lengthy and expensive: difficulty and cost do increase because of the overall system dimension that depends on the number of uncertainties. Uncertain parameters have to be parametrically investigated to determine robust performance of the control laws via gridding approaches. In particular in this paper we propose to consider two methods: (i) a conventional Monte Carlo analysis, and (ii) a worst-case analysis, i.e., an optimization process to find an estimation of the true worst-case behaviour. Both techniques allow to verify that the design is robust enough to meet the system performance specification in case of uncertainties. [Copyright &y& Elsevier]

Abstract: Navigation, control and guidance of the propulsive phase of planetary landing, e.g. on Mars (or the Moon), with a soft landing being the only target, are driven by Inertial Measurement Units and a radar altimeter/velocimeter. Their measurements are affected by bias and scale errors. The latter ones are aggravated by the attitude navigation error as it accumulates during the ballistic (and aerodynamic) flight after orbiter separation and couples for most of the descent trajectory with the vehicle axis inclination from the local vertical direction. By complementing the center-of-mass dynamics with appropriate disturbance state equations driven by noise vectors and estimating the noise from the model error (plant measurements minus model output), scale errors and bias can be retrieved in real time in the form of disturbance state variables. Although a similar complement is adopted in the standard navigation algorithms, it takes the form of an output disturbance, which may lead to unobservability. In this paper instead, the disturbance complement is designed to be fully observable, which may require that the derivatives of smooth systematic errors be pushed up to the command channel (a form of back-stepping). It is then viable, unlike standard navigation, to eliminate them from position and velocity tracking errors through disturbance rejection, under appropriate convergence conditions and sensor layout. It will, however, be demonstrated in this paper that the same result cannot be achieved under pure feedback control. Since constant errors (bias) become zero through back-stepping, a well known fact derives: bias can only be eliminated by disposing of supplementary sensors. To further enlighten and solve the question of bias rejection, a further case study is treated. The attitude control of drag-free satellites is considered, where fine accelerometers allow for the rejection of wide-band aerodynamic torques (think of low-Earth orbit spacecrafts) at the price of attitude divergence because of accelerometer bias and drift. The spacecraft attitude can be made bounded and accurate, if bias and drift are modeled as angular accelerations, affecting the attitude. They are estimated by attitude sensors like star trackers and then are rejected by the attitude control. The results in the soft landing and drag free case studies are illustrated by simulated runs and Monte Carlo trials. [Copyright &y& Elsevier]

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 J 2 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. • A two-stage mission design has been proposed. • Spacecraft attitude, tether flexibility and extensibility are accounted for. • Attitude stabilization and tether deployment are simultaneously achieved. [ABSTRACT FROM AUTHOR]

Abstract: Nano-satellites provide space access to broader range of satellite developers and attract interests as an application of the space developments. These days several new nano-satellite missions are proposed with sophisticated objectives such as remote-sensing and observation of astronomical objects. In these advanced missions, some nano-satellites must meet strict attitude requirements for obtaining scientific data or images. For LEO nano-satellite, a magnetic attitude disturbance dominates over other environmental disturbances as a result of small moment of inertia, and this effect should be cancelled for a precise attitude control. This research focuses on how to cancel the magnetic disturbance in orbit. This paper presents a unique method to estimate and compensate the residual magnetic moment, which interacts with the geomagnetic field and causes the magnetic disturbance. An extended Kalman filter is used to estimate the magnetic disturbance. For more practical considerations of the magnetic disturbance compensation, this method has been examined in the PRISM (Pico-satellite for Remote-sensing and Innovative Space Missions). This method will be also used for a nano-astrometry satellite mission. This paper concludes that use of the magnetic disturbance estimation and compensation are useful for nano-satellites missions which require a high accurate attitude control. [Copyright &y& Elsevier]

Abstract The problem of stabilizing the time delayed control of a flexible space structure is analyzed in this paper. A free floating platform is used to investigate the space multibody dynamics and control. A first necessary step to develop stabilizing techniques is considered the availability of a set of measurements as complete as possible: in particular measurements of the elastic vibrations are necessary in addition to classic attitude measurements. At the scope, a net of PZT sensors have been designed and manufactured on a composite material panel, purposely built to resemble a space structure. A combined use of the PZT/optical sensor is proposed, where the role of the camera is to estimate the PZT parameters that can be changed after the manufacturing or for environmental aging. When this calibration process is performed, PZT can be used as standalone sensors for measuring also the elastic displacement of the structure. Once these measurements of attitude and elastic displacement are obtained, two stabilizing techniques have been developed, the Finite Spectrum Analysis, already known in literature, and the newly developed High Fidelity Filter approach, based on the design of a Kalman filter with large confidence on the process dynamics. It is shown that both techniques manage to increase the delay margin of the system, thus obtaining a stable maneuver, but the second approach reach this goal with very low residual vibrations and a remarkable fuel saving. Highlights • The problem of controlling a flexible spacecraft is faced. • A net of sensors including PZT patches and a camera is developed. • The performance of different algorithms for vibration reduction is analyzed. • The results suggest that the use of properly designed Kalman filters is a meaningful solution. • Experimental attitude maneuvers validate the conclusions. [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]

Spacecraft attitude control solutions typically are torque-level algorithms that simultaneously control both the attitude and angular velocity tracking errors. In contrast, robotic control solutions are kinematic steering commands where rates are treated as the control variable, and a servo-tracking control subsystem is present to achieve the desired control rates. In this paper kinematic attitude steering controls are developed where an outer control loop establishes a desired angular response history to a tracking error, and an inner control loop tracks the commanded body angular rates. The overall stability relies on the separation principle of the inner and outer control loops which must have sufficiently different response time scales. The benefit is that the outer steering law response can be readily shaped to a desired behavior, such as limiting the approach angular velocity when a large tracking error is corrected. A Modified Rodrigues Parameters implementation is presented that smoothly saturates the speed response. A robust nonlinear body rate servo loop is developed which includes integral feedback. This approach provides a convenient modular framework that makes it simple to interchange outer and inner control loops to readily setup new control implementations. Numerical simulations illustrate the expected performance for an aggressive reorientation maneuver subject to an unknown external torque. [ABSTRACT FROM AUTHOR]

Based on Nodal Coordinate Formulation (NCF) and Absolute Nodal Coordinate Formulation (ANCF), this paper establishes rigid-flexible coupling dynamic model of the spacecraft with large deployable solar arrays and multiple clearance joints to analyze and control the satellite attitude under deployment disturbance. Considering torque spring, close cable loop (CCL) configuration and latch mechanisms, a typical spacecraft composed of a rigid main-body described by NCF and two flexible panels described by ANCF is used as a demonstration case. Nonlinear contact force model and modified Coulomb friction model are selected to establish normal contact force and tangential friction model, respectively. Generalized elastic force are derived and all generalized forces are defined in the NCF-ANCF frame. The Newmark- β method is used to solve system equations of motion. The availability and superiority of the proposed model is verified through comparing with numerical co-simulations of Patran and ADAMS software. The numerical results reveal the effects of panel flexibility, joint clearance and their coupling on satellite attitude. The effects of clearance number, clearance size and clearance stiffness on satellite attitude are investigated. Furthermore, a proportional-differential (PD) attitude controller of spacecraft is designed to discuss the effect of attitude control on the dynamic responses of the whole system. [ABSTRACT FROM AUTHOR]

In space the manipulators working space is characterized by the microgravity environment. In this environment the spacecraft floats and its rotational/translational motion may be excited by any internal and external disturbances. The complete system, i.e., the spacecraft and the associated robotic manipulator, floats and is sensitive to any reaction force and torque related to the manipulator's operation. In this sense the effort done by the robot may result in torque about the system center of mass and also in forces changing its translational motion. This paper analyzes the impact of the robot manipulator dynamics on the attitude motion and the associated control effort to keep the attitude stable during the manipulator's operation. The dynamics analysis is performed in the close proximity phase of rendezvous docking/berthing operation. In such scenario the linear system equations for the translation and attitude relative motions are appropriate. The computer simulations are implemented for the relative translational and rotational motion. The equations of motion have been simulated through computer by using the MatLab software. The LQR and the PID control laws are used for linear and nonlinear control, respectively, aiming to keep the attitude stable while the robot is in and out of service. The gravity-gradient and the residual magnetic torque are considered as external disturbances. The control efforts are analyzed for the manipulator in and out of service. The control laws allow the system stabilization and good performance when the manipulator is in service. [ABSTRACT FROM AUTHOR]

Up to April 6, 2016, there are 17,385 large debris in orbit around the Earth, which poses a serious hazard to near-Earth space activities. As a promising on-orbit debris capture strategy, tethered space robots (TSRs) have wide applications in future on-orbit service owing to its flexibility and great workspace. However, lots of problems may arise in the Tethered Space Robots (TSRs) system from the approaching, capturing, postcapturing and towing phases. The postcapture combination attitude takeover control by the TSR is studied in this paper. Taking control constraints, tether oscillations and external disturbances into consideration, a fast terminal sliding mode control (FTSMC) methodology with dual closed loops for the flexible combination attitude takeover control is designed. The unknown upper bounds of the uncertainties, external disturbances are estimated through adaptive techniques. Stability of the dual closed loop control system and finite time convergence of system states are proved via Lyapunov stability theory. Besides, null space intersection control allocation was adopted to distribute the required control moment over TSR's redundant thrusters. Simulation studies have been conducted to demonstrate the effectiveness of the proposed controller with the conventional sliding mode control(SMC). [ABSTRACT FROM AUTHOR]

The dynamic characteristics and performance evaluation for the part strut failure of the vibration isolation platform are presented in this paper. The first step provides and mathematically describes two types of strut failure: fractured and stuck. Secondly, the dynamic model of the vibration isolation platform, which considers the part strut failure, is established using the Newton-Euler method and a constraint equation to evaluate its dynamic characteristics and performance. Then, with reasonable assumptions, the dynamic model of the satellite, which has a vibration isolation platform and vibration sources (such as control moment gyros) of three working situations (without and with two types of strut failure), is simplified to analyse the frequency domain characteristic and coupling characteristic with the attitude control system. Finally, a numerical simulation is used to study the effect of the vibration isolation platform with part strut failure on the attitude control and stabilization, and the attitude control performance is evaluated. [ABSTRACT FROM AUTHOR]

This paper investigates the combined attitude and sun-tracking control problem in the presence of external disturbances and internal disturbances, caused by flexible appendages. A new method based on Pythagorean trigonometric identity is proposed to drive the solar arrays. Using the control input and attitude output, a disturbance observer is developed to estimate the lumped disturbances consisting of the external and internal disturbances, and then compensated by the disturbance observer-based controller via a feed-forward control. The stability analysis demonstrates that the desired attitude trajectories are followed even in the presence of external disturbance and internal flexible modes. The main features of the proposed control scheme are that it can be designed separately and incorporated into the baseline controller to form the observer-based control system, and the combined attitude and sun-tracking control is achieved without the conventional attitude actuators. The attitude and sun-tracking performance using the proposed strategy is evaluated and validated through numerical simulations. The proposed control solution can serve as a fail-safe measure in case of failure of the conventional attitude actuator, which triggered by automatic reconfiguration of the attitude control components. [ABSTRACT FROM AUTHOR]

The paper describes the development and simulation results of a full ADCS subsystem for the deOrbitSail drag sail mission. The deOrbitSail satellite was developed as part of an European FP7 collaboration research project. The satellite was launched and commissioning started on 10th July 2015. Various new actuators and sensors designed for this mission will be presented. The deOrbitSail satellite is a 3U CubeSat to deploy a 4 by 4 m drag sail from an initial 650 km circular polar low earth orbit. With an active attitude control system it will be shown that by maximising the drag force, the expected de-orbiting period from the initial altitude will be less than 50 days. A future application of this technology will be the use of small drag sails as low-cost devices to de-orbit LEO satellites, when they have reached their end of life, without having to use expensive propulsion systems. Simulation and Hardware-in-Loop experiments proved the feasibility of the proposed attitude control system. A magnetic-only control approach using a Y-Thomson spin, is used to detumble the 3U Cubesat with stowed sail and subsequently to 3-axis stabilise the satellite to be ready for the final deployment phase. Minituarised torquer rods, a nano-sized momentum wheel, attitude sensor hardware (magnetometer, sun, earth) developed for this phase will be presented. The final phase will be to deploy and 3-axis stabilise the drag sail normal to the satellite’s velocity vector, using a combined Y-momentum wheel and magnetic controller. The design and performance improvements when using a 2-axis translation stage to adjust the sail centre-of-pressure to satellite centre-of-mass offset, will also be discussed, although for launch risk reasons this stage was not included in the final flight configuration. To accurately determine the drag sail’s attitude during the sunlit part of the orbit, an accurate wide field of view dual sensor to measure both the sun and nadir vector direction was developed for this mission. The calibration results for this new Cubesat sensor ( CubeSense ), will also be presented. [ABSTRACT FROM AUTHOR]

This paper investigates the time efficient maneuver of rigid satellites with inertia uncertainty and bounded external disturbance. A redundant cluster of four reaction wheels is used to control the spacecraft. To make full use of the controllability and avoid frequent unload for reaction wheels, a maximum output torque and maximum angular momentum constrained torque distribution method is developed. Based on this distribution approach, the maximum allowable acceleration and velocity of the satellite are optimized during the maneuvering. A novel braking curve is designed on the basis of the optimization strategy of the control torque distribution. A quaternion-based sliding mode control law is proposed to render the state to track the braking curve strictly. The designed controller provides smooth control torque, time efficiency and high control precision. Finally, practical numerical examples are illustrated to show the effectiveness of the developed torque distribution strategy and control methodology. [ABSTRACT FROM AUTHOR]

Gyroscopic actuators for satellite control have attracted significant research interest over the years, but their viability for the control of small satellites has only recently started to become clear. Research on variable speed gyroscopic actuators has long been focused on single gimbal actuators; double gimbal actuators typically operate at constant wheel spin rate and allow tilt angle ranges far larger than the ranges needed to operate most satellite missions. This research examines a tilted wheel, a newly proposed type of inertial actuator that can generate torques in all three principal axes of a rigid satellite using a spinning wheel and a double tilt mechanism. The tilt mechanism tilts the angular momentum vector about two axes providing two degree of freedom control, while variation of the wheel speed provides the third. The equations of motion of the system lead to a singularity-free system during nominal operation avoiding the need for complex steering logic. This paper describes the hardware design of the tilted wheel and the experimental setup behind both standalone and spherical air-bearing tables used to test it. Experimental results from the air bearing table are provided with the results depicting the high performance capabilities of the proposed actuator in torque generation. [ABSTRACT FROM AUTHOR]

This paper examines attitude synchronization and tracking problems with model uncertainties, external disturbances, actuator failures and control torque saturation. Two decentralized sliding mode control laws are proposed and analyzed based on algebraic graph theory. Using Barbalat׳s Lemma, it is shown that the control laws guarantee each spacecraft approaches the desired time-varying attitude and angular velocity while maintaining attitude synchronization among the other spacecraft in the formation. The first controller is designed in the presence of model uncertainties, external disturbances, and actuator failures. The results are extended to the case with control input saturation in the second controller. Both control laws do not require online identification of failures. Numerical simulations are presented to show the effectiveness of the proposed attitude synchronization and tracking approaches. [ABSTRACT FROM AUTHOR]