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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]

Electric propulsion, which is an advanced form of thrust production, is widely used in small satellites. Measuring the thrust of electric propulsion in an orbit is difficult because it is very small (μN∼mN). However, the thrust of electric propulsion can be determined by evaluating the change in the speed of a momentum wheel. The contributions of this paper mainly include two aspects: 1) a robust PID controller and 2) a thrust inversion method. The torque of the electric propulsion is regarded as the disturbance torque in this paper. To achieve finite-time attitude stabilization of a satellite during electric propulsion, a robust PID controller is designed. Furthermore, the thrust of the electric propulsion can be determined in the stable attitude of the satellite by only measuring the speed derivative of the momentum wheel. Compared with the traditional thrust inversion method, the designed thrust inversion method is simple and effective and does not use a filtering algorithm. In the HIL experiment, the error rate between the inverted thrust (31.287 μN) and the designed thrust (30 μN) is 4.29%. Most importantly, the thrust inversion method is verified in orbit based on the small satellite APSCO SSS-1, which is consistent with the HIL experimental results and ground test results. • The torque of electric propulsion is regarded as the disturbance torque. • The controller can fulfill finite-time attitude stabilization of the satellite. • Thrust can be calculated by evaluating the speed change of the momentum wheel. • Thrust inversion method is verified in orbit based on the small satellite SSS-1. [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]

The agility of satellites equipped with large and flexible structures, such as solar panels or antennas, poses a significant challenge due to the influence of elastic dynamics on attitude motion. This results in reduced pointing accuracy and dimensional stability of the payload. To address this issue, an Active Collocated Attitude Control (ACAC) system has been developed, utilizing actuators for both the elastic dynamics (distributed piezoelectric devices) and attitude control (ON-OFF thrusters). In the initial phase of the project, the algorithm underwent extensive verification using a high-fidelity numerical model of a spacecraft with two very-low frequency flexible solar panels and a payload consisting of two long booms and a transversal antenna's reflector. The paper focuses on the second phase of the project, which revolves around the experimental verification of the control architecture concept. To achieve this, a scaling procedure employing the dynamic analogy approach was implemented to determine the design parameters of a free-floating platform equipped with elastic appendages. This ensured that the resulting elastic multibody system displayed similar dynamic behavior to the full-scale satellite. Based on the scaling parameters, a test rig was manufactured, and actuators and sensors were optimally positioned. The performance of the ACAC system was then compared to that of traditional benchmark controllers. The results show that the ACAC system demonstrates superior capabilities in performing fast slew maneuvers without developing unstable interactions with the elastic dynamics. • Agile maneuvers of large satellites could excite rigid-flexible interactions. • A robust control of attitude and elastic dynamics is implemented to improve agility. • Performance obtained with numerical models is verified with scaled-down test-rig. • Simplified dynamic equations are developed to determine scaling conditions. • Experimental results of the robust control are consistent with simulation. [ABSTRACT FROM AUTHOR]

This paper concentrates on robust attitude control issues for spacecraft formation in gravitational-wave detection missions. Notably, with the large-scale communication delays brought by the million kilometers-scale inter-satellite distance, the exponential convergence of the closed-loop delayed system is firstly achieved via a new delay-dependent nominal attitude coordinated control scheme. Then, a finite-time disturbance observer is deployed to reconstruct the unknown lumped disturbance from internal uncertainties and the external environment. Meanwhile, an integral sliding manifold is embedded in the delay-dependent attitude controller for superior systems' robustness. Finally, two simulation examples illustrate the feasibility and effectiveness of the proposed attitude control strategy for in-orbit gravitational-wave detection spacecraft systems. • Spacecraft formation attitude control in gravitational-wave detection is studied. • Large-scale communication delays (LCD) among formation members are firstly modeled. • An attitude coordinated controller is put forward to handle the LCD. • An integral sliding manifold with disturbance observer is embedded for superior robustness. [ABSTRACT FROM AUTHOR]

Flexible vibration has a great impact on flexible spacecraft attitude control, which will reduce the pointing accuracy of the spacecraft platform and bring severe challenges to the design of a high-precision attitude control system for flexible spacecraft. Aiming at the problem of attitude and vibration coupling control of flexible spacecraft, an anti-unwinding attitude coupling controller using the positive position feedback (PPF) control is proposed in this paper. The model compensation control term is constructed based on the prior information to improve the tracking performance. The proposed attitude coupling control method can effectively suppress the flexible vibration while achieving more accurate attitude control. The obtained closed-loop system under the proposed control scheme is asymptotically stable, which is proved by the Lyapunov stability theory. The numerical simulation illustrates that the proposed method can effectively improve the attitude control accuracy of the flexible spacecraft and is capable of large-angle attitude maneuver control because of its prominent performance in attitude tracking and its unwinding-free performance. • Coupling control for flexible spacecraft. • Active vibration control using positive position feedback. • Model compensation control term is constructed to improve the dynamic performance. • The obtained closed-loop system achieves the asymptotic stability and anti-unwinding performance. [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]

There is an increasing demand in the space industry for methods allowing for gyroless spacecraft attitude determination to either provide low cost sensor fusion or to recover from gyro failures. The gyroless configuration used in this paper is a combination of magnetometers and Sun sensors. Sun sensors are subject to measurement discontinuities during solar eclipses. Gyroless attitude sensing is known to increase sensitivity to measurement noise and model uncertainty. This paper presents a new Laplace Particle Filter with a Singular Value Decomposition to enhance attitude determination performance under these circumstances, while also preventing filter degeneracy. An orthogonal Procrustes problem is formulated to minimise the Wahba cost function and determine the attitude matrix from the Sun and magnetic field vectors and a third virtual predicted measurement vector, which is also available during Sun occlusion. The proposed filter is compared to a regularised particle filter approach, which handles particle filter degeneracy. Both filters are applied to the attitude determination of a small Earth observation satellite using three Sun sensors available during the sunlit phase and three orthogonal magnetometers. A Monte Carlo numerical simulation analysis demonstrates that the proposed particle filter significantly outperforms the regularised particle filter in terms of accuracy and robustness to the Sun sensor measurement discontinuities during solar eclipse and is also more accurate when both types of sensor measurements are available. • Gyroless attitude determination was obtained with Sun sensors and magnetometers. • A Laplace particle filter uses a Procrustes algorithm to solve Wahba's problem. • A Procrustes problem is formulated to approximate the maximum a posteriori attitude. • Estimation accuracy and robustness is significantly enhanced during solar eclipse. [ABSTRACT FROM AUTHOR]

Attitude control methods of highly flexible spacecraft have seen increased interest over the last decades thanks to the technological development of flexible solar panels and deployables, which improves the capabilities of satellites. However, a high-fidelity model of the flexible dynamics is hard to obtain on-ground testing because not all modes can be observed, complicating the controller design. This paper proposes a method to develop a high-fidelity model of a spacecraft with a flexible appendage subject to large deformations by modeling it as a finite series of rigid links connected by torsional springs and dampers. To overcome the uncertainties or unknowns in the flexible dynamics, an onboard estimation through an adaptive controller is performed for these unknowns while the spacecraft is maneuvered. The controller uses integral concurrent learning (ICL), an adaptive scheme that records inputs and outputs provided by sensors mounted on the flexible body. The novelty of this investigation is the development of self-adapting control gains for the learning matrix obtained from ICL. This approach achieves the objective of tracking a desired trajectory while accurately learning the unknown physical parameters of the flexible appendage by only using the recorded measurements of the appendage displacements and the spacecraft attitude and velocities. It was observed that for a finer discretization of the flexible appendage and therefore a higher fidelity model of the flexible dynamics, the estimation algorithm is able to observe all the frequencies necessaries to learn the unknown mechanical properties of the flexible body. • Modeling a flexible deployable as finite chain of rigid links connected by unknown torsional joints to consider large deformations and reduce onboard computational effort. • Accurately estimated unknown flexible dynamics using integral concurrent learning (ICL). • Introduced a self-tuning algorithm for eigenvalue placement of the learning matrix to guarantee the convergence of any number of estimated unknowns to their true value. [ABSTRACT FROM AUTHOR]

With the ability to detect low-frequency gravitational waves (GWs), space-borne detectors will play an important role in exploring the universe in the future. The TianQin project proposed in China participates in this challenge by aiming to detect millihertz GWs. The TianQin GW detector consists of three spacecraft forming a triangular constellation. Each spacecraft carries a pair of test masses (TM) which, inside shielding cages, are left to free fall along the local geodesic. In this way, a pair of TMs on two separate spacecraft can be each other aligned to become the gravitational references of the GW detector, made by an inter-satellite laser interferometer. Each free-falling TM is limited by parasitic forces to be appropriately bounded. The relevant acceleration bound has been fixed to 10−15 m/s2/ Hz within the TianQin measurement bandwidth (MBW) ranging from 0.1 mHz to 0.1 Hz. In turn, TM to cage fluctuations must be kept below 4nm/ Hz to limit the stiffness coupling with the spacecraft displacement caused by non-gravitational disturbances, such as solar radiations and the thruster noise. Suppression of such disturbances calls for a challenging drag-free control technology, since the center of mass (CoM) of a single spacecraft cannot track the separated CoMs of two TMs simultaneously, and consequently tracking must be limited to three degrees of freedom (DoF): two non-orthogonal sensitive axes (one for each TM) and the perpendicular direction to their plane. Control design and simulated tests of this paper will be restricted to the drag-free control along the sensitive axes. The remaining nine DoFs of the two TMs (position and attitude) are controlled by electrostatic suspensions, not to be treated here, but accounted for in the simulated trials. The design of the two-DoF drag-free control relies on a model-based control methodology, the Embedded Model Control (EMC), capable of predicting and suppressing unknown disturbances within the required bandwidth and of decoupling the TM dynamics along the non-orthogonal sensitive axes. The paper starts with the nonlinear model of the TM to cage dynamics, followed by the relevant EMC design, restricted to the sensitive axes. Numerical simulations are employed to validate closed-loop performance and robust stability. Simulated results show that the residual TM to cage fluctuations can be kept below 3nm/ Hz , which leaves a margin within the required spectral bound. The EMC methodology discussed in the paper can provide a reference for future developments and implementations. • Test mass dynamics can be decoupled by 'thruster coupling' or 'sensor coupling'. • Embedded Model Control predicts and compensates unknow disturbances. • Reference generator smooths test mass centering process. • Test mass to cage displacement has satisfied TianQin requirement. • Monte Carlo trails have validated the closed-loop robust stability. [ABSTRACT FROM AUTHOR]

50,000+ satellites are scheduled for launch by 2030, bringing both massive opportunities and equally massive challenges. Key among these challenges is the ability to sustainably operate Low Earth Orbit (LEO) spacecraft in an increasingly congested environment where collision avoidance manoeuvres are routine and fuel is limited. This paper explores the application of ionospheric drag as an alternative, propellantless collision avoidance mechanism for satellites with altitudes ranging between 350 and 500 km. This is achieved through the use of an in-house propagation suite developed to support UNSW Canberra Space's in-orbit Ionospheric Aerodynamic Experiment (IEX) on the M2 spacecraft (NORAD ID: 47967 & 45727). The relative impact of neutral (thermospheric) to charged (ionospheric) drag on along-track separation for a range of spherical bodies with surface potentials ranging between 0 V to -100 V with respect to a quasi-neutral freestream plasma was investigated. This provides a first-order indication of the potential effectiveness of ionospheric drag for collision avoidance based on prior numerical and ground-based investigations into ionospheric drag. The Debris Risk Assessment and Mitigation Analysis software from the European Space Agency was also harnessed to discern the relative performance of both the charged and flared attitude satellites against a common control satellite. The ESA stipulates that a Collision Probability Level value greater than the accepted value of 1 E − 04 between two satellites requires a risk avoidance manoeuvre, with a conservative approach requiring manoeuvres for values greater than 1 E − 05 . For all test regimes at 350 k m altitude, at least a 60% reduction in time to achieve the NASA approved risk reduction factor of 31.6 at the time of closest approach for a pre-manoeuvre Collision Probability Level of 2. 8 E − 05 was achieved when using the flared neutral method in comparison to the feathered charged method. Based on these results, this paper finds that ionospheric drag may represent a useful alternative propellantless collision avoidance mechanism for satellites within the altitude range of 350 to 500 km. Future work is needed to refine ionospheric drag models and consider the absolute effectiveness of using ionospheric drag at higher altitudes and in tandem applications to the flared arrangement. • Ionospheric drag can successfully manoeuvre satellites to avoid collisions. • Ionospheric drag is inferior to industry standard avoidance manoeuvres. • High Low Earth Orbit altitudes show ionospheric drag becomes more competitive. • The adopted python approach provides flexibility for future satellite simulations. [ABSTRACT FROM AUTHOR]

Magnetic control has been used successfully for momentum management and attitude control when pointing precision requirements are loose. Typically, such systems are constructed using configurations of magnetic coils oriented in a way to provide command authority by varying the magnetic moment about all three body axes. Although magnetic control actuators and related systems are typically robust and not as prone to failure when compared to reaction wheels or thrusters, it is of interest to see if coarse 3-axis attitude pointing can be achieved with the loss of control about one of the 3 axes of magnetic moment actuation. These types of cases could occur during the normal life of the spacecraft or could also occur intentionally on very small satellites where space is limited, so any control modes that could prolong the life of a damaged satellite or intentionally underactuated satellite would be useful for some missions. This paper demonstrates the feasibility of 3-axis pointing of spacecraft using only two axes of magnetic moment actuation using linear time varying analysis to provide sufficient conditions for controllability. The paper then develops an LQR-based controller for these two torquer cases. Finally, the paper concludes with simulations that indicate it is feasible to control with only two axes of magnetic moment, although with the expected increased time to achieve targeted attitudes and reduced disturbance rejection. • Identifies sufficient conditions for controllability in systems with only two magnetic coils. • Implementation of a LQR controller using two magnetic coils. • Testing of controller with disturbance torques relevant to Low Earth Orbit. • Monte-Carlo analysis to test robustness of this approach to parameter variations. [ABSTRACT FROM AUTHOR]

Experimental verification of numerical models is a key aspect in all scientific and technological disciplines, yet space systems can be particularly hard to test on ground: in fact, a terrestrial mock-up replicating the orbital dynamics should be in a microgravity environment. One of the most common approaches to (at least partially) realize this condition, consists in free-floating systems, in which the friction between an experimental platform and the working surface is almost completely removed. The vertical axis is still subject to gravity, and only 3° of freedom (two horizontal translations and one rotation about the vertical axis) are available to replicate the orbital behavior. Our research group had been developing since 2012 a free-floating platform named PINOCCHIO (Platform Integrating Navigation and Orbital Control Capabilities Hosting Intelligent Onboard). The need to investigate with higher accuracy the attitude control problem for flexible satellites called for the design, realization and characterization of a new platform with enhanced performance. This paper describes the main components of the system, including the pressure air system, the navigation and communication system, the Guidance, Navigation and Control (GNC) architecture and relevant hardware. The attention is specifically on the characterization of the subsystems' performance. The physical (mass, moment of inertia), control (thrusters' torque and force) and Inertial Measurement Unit (IMU) characteristics are experimentally measured with the associated uncertainty level. The accuracy of the external metrology system in static and dynamic conditions is also characterized, obtaining a 100 μm-level resolution for external measurement of rigid and elastic displacements; this information is fundamental for performing tests of fine pointing control algorithms for flexible and agile satellites. The proposed characterization methodology and associated results could be of interest for university and industry laboratories intended to develop new free-floating platforms. • The characterization procedure of a free-floating platforms is reported. • A stepwise procedure requires to first characterize the supplied thrust accurately. • Inertia properties can be next determined. • Accurate optical navigation is presented and characterized. • Performance of the platform for flexible satellite attitude test is introduced. [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]

The robust non-fragile attitude stabilization controller for flexible spacecraft with model parameters uncertainty is difficult to be synthesized with negative imaginary system which is widely applicable for the vibration control of flexible structures. To address this issue, this paper proposes a negative imaginary output feedback H ∞ synthesis strategy, in which the model parameters uncertainty, controller gain perturbations, unknown inertia and actuator faults are considered simultaneously. This strategy can guarantee that the closed-loop attitude control system has asymptotical stability with H ∞ performance and approximates to negative imaginary. The iterative algorithm under linear matrix inequality constraints gives the conditions for the existence of such controller. In addition, external disturbance and input saturation are also taken into account. Compared with the general non-fragile H ∞ controller, the numerical simulations demonstrate the superiority of the proposed controller in attitude stabilization and vibration suppression of flexible spacecraft. • The spacecraft attitude and vibration are stabilized with controller perturbations. • The NI controller can well suppress the elastic vibration of flexible appendages. • The two-stage framework is of effective controller performance synthesis. [ABSTRACT FROM AUTHOR]

This paper proposed and verified an optimal estimation method of satellite attitude deviation in orbit based on the frequency domain features of microwave remote sensing data from the Gaofen-3B (GF-3B) satellite. The Synthetic Aperture Radar (SAR) imaging results of the Amazon forest's strip pattern were obtained at different imaging incident angles of the GF-3B satellite. Then, the inversions of the Doppler center deviation of the SAR signal obtained the corresponding beam pointing errors at different imaging incident angles. Next, a three-degree-of-freedom nonlinear least square fitting model was established between the Doppler center deviation of the SAR signal and satellite attitude deviation. Combined with the measured SAR signals, satellite attitude deviations in different dimensions (pitch, yaw, and roll) were separated and estimated. On this basis, the GF-3B satellite proceeded with an on-orbit attitude correction test. The maximum Doppler center deviation at all imaging incident angles decreased from 400Hz to 46Hz, and the residual beam pointing errors were less than 0.01°. The proposed optimal estimation method can guarantee the on-orbit attitude deviation suppression and image quality improvement of the SAR satellites. • An optimal attitude deviations estimation method of SAR satellites is proposed based on the echo frequency domain features. • The fitting model of attitude deviations is established by combining microwave remote sensing imagery and orbit dynamics. • The proposed method has been used for the GF-3B satellite and verified by in-orbit imaging experiments. [ABSTRACT FROM AUTHOR]

This paper deals with the attitude motion of LEO satellites with deployable side panels designed for passive aerodynamic stabilization in a rarefied atmosphere. The influence of the aerodynamic and gravitational torques on the planar attitude motion near the unstable and stable equilibrium positions is studied. The presence of the unstable equilibrium position and small perturbations such as the oscillations of the flexible panels is the cause of chaos. A critical altitude is found above which the chaos is possible. The equations of planar attitude motion of the satellite with deployed flexible panels are obtained. The chaotic behavior of the system is demonstrated through numerical simulations of the attitude motion of a 3U CubeSat. The results of this paper can be used to analyze the applicability of passive aerodynamic stabilization for LEO satellites. • Planar attitude motion of a LEO satellite with deployable side panels is studied. • The satellite as a rigid body with the flexible appendages is considered. • For a wide range of altitudes, stable and unstable equilibrium positions are found. • The chaotic behavior of the satellite with flexible panel is proven. • It's been shown that flexible panels do not always lead to stabilization of motion. [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 model-free formation control problem is addressed for multiple satellites, whose dynamics involves rotation-translation coupling and nonlinearities. A two-step adaptive control strategy is proposed based on the reinforcement learning theory to iteratively solve the optimal formation control problem without knowledge of each satellite dynamics. The resulted controller is composed of an attitude controller to achieve the desired attitude, and a relative position controller to construct the satellite formation. The proposed global closed-loop control system is proven to be asymptotically stable. Simulation results on rotation-translation coupled satellite system confirm the effectiveness of the proposed adaptive optimal formation control strategy. • The model-free formation control problem is addressed for multiple satellites. • The dynamics of each satellite involves rotation-translation coupling and nonlinearities. • Optimal formation control for the real uncertain coupled multi-satellite system can be achieved. • A two-step learning scheme is proposed to address the rotation-translation coupled dynamics. • The accurate dynamical model parameters are not required to be known by the learning method. [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]

This paper investigates the attitude control problem of a two-body combined spacecraft by considering the unilateral contact constraint on the active spacecraft's end-effector. Previous work has shown that this constraint may cause a relative sliding motion between the end-effector and the target, which has a large impact on the combined spacecraft dynamics and postcapture control process. In this paper, an extra degree-of-freedom approach is adopted and a new dynamical model is developed to efficiently describe the movement of the combined spacecraft in both rigid connecting and relative motion modes. A new control law is also developed to reduce the relative sliding motion between the end-effector and the target while controlling the attitude of the combined system. Simulation results show that the extra degree-of-freedom model proposed in this paper has the correct dynamic properties and that the newly developed controller is a significant improvement over traditional PD controllers in maintaining the geometric configuration of the combined spacecraft. • The post-capture operation with unilateral contact constraint is studied. • A new dynamic model is developed based on extra-degree-of-freedom approach. • The modified reactionless controller is proposed to control the combined spacecraft. [ABSTRACT FROM AUTHOR]

Disturbance torques, such as atmospheric drag torque and gravity gradient torque, have always been an issue in attitude control for satellites in low Earth orbit. Although attitude control systems using disturbance torques has been studied with various methods, most previous research has focused on attitude control using aerodynamic stability to remove disturbances. We propose a new attitude control system that actively uses disturbance torque. This system controls the attitude using disturbance torques by changing the equilibrium point of the attitude motion by changing the shape of the satellite. We formulate the attitude dynamics of a variable-shape satellite as a preliminary study. This paper shows that the satellite attitude can be adjusted by varying the equilibrium point of attitude motion by changing the satellite shape. Also, we investigate the stability around the equilibrium point for each shape, and the results show the distribution of the stability index and the equilibrium point for each shape. Finally, we show that the attitude asymptotically converges to an equilibrium point by using the variable shape function and magnetic control. • A new attitude control system that uses variable-shape function and disturbance torques is proposed. • Equilibrium point can be changed by changing shape of satellite. • Attitude dynamics of variable-shape satellite is formulated. • Relationship among the shapes, equilibrium points, and stability is investigated. [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]

Spacecraft are required to achieve fast attitude reorientation in many space missions. However, the attitude maneuver will be limited by the bounded and pointing constraints. The presence of these complex multiaxis constraints will largely reduce the feasible attitude space and greatly degrade the solving efficiency of the planning algorithm. For the time-optimal spacecraft reorientation problem under multiaxis constraints, a new heuristic decomposition planning on the virtual domain (HDPV) method is proposed in this paper. The parameterized attitude path is first determined on the virtual domain. Then the virtual-domain-based variable-step approach is presented to rapidly check the pointing constraints and generate the sets of candidate rotational-path decomposition nodes. The composite-position heuristics is developed to select the best node during the decomposition optimization, which considers the maneuver time and connection smoothness of paths. The total continuous parameterized path under keep-out and keep-in constraints is yielded through recursive planning. Finally, the time-optimal path parameterization approach is adopted to minimize the maneuver time along the total path and generate a fast attitude trajectory under bounded constraints on the time domain. Simulation results demonstrate the high computational efficiency and good optimization effect of the proposed method. • Propose the heuristic decomposition planning on the virtual domain method. • New virtual-domain-based variable-step approach and composite-position heuristics. • Address the time-optimal problem under multiaxis constraints. • The HDPV method has a high computational efficiency. • Generate a fast spacecraft reorientation trajectory. [ABSTRACT FROM AUTHOR]

This paper considers the nonlinear interaction of the rigid body dynamics of a spacecraft with the sloshing dynamics of the liquid propellant in its partially filled spherical tank during an orbital transfer. Fuel slosh dynamics are described by a three-dimensional two-pendulum model characterizing the first two sloshing modes, and the coupled equations of motion of the vehicle and the fuel masses are derived by means of quasi-Lagrangian equations. In addition, this paper studies active nonlinear control methods to control the attitude of the spacecraft while suppressing the sloshing of fuel. Simulations are conducted to illustrate the effectiveness of these control methods. • Equations of motion for a spacecraft with fuel slosh are obtained. • Fuel slosh dynamics are described by a two-spherical-pendulum model characterizing the first two sloshing modes. • Sliding mode, fuzzy logic, and Lyapunov-based feedback control laws are designed. • Simulation results are included to illustrate the effectiveness of the controllers. [ABSTRACT FROM AUTHOR]

Traditionally, controlled spacecraft re-entries have been conducted using propulsive de-orbit burns which are risky, expensive, and may not be possible for all vehicles. Recently, the miniaturization of technology has ushered in a new class of small satellites (such as CubeSats) that are too small to host thrusters but may require a controlled de-orbit if they contain materials capable of surviving re-entry. For all space vehicles requiring a controlled re-entry, the ability to harness the naturally occurring aerodynamic drag force for orbit control provides a cheaper and more reliable alternative to chemical propulsion. This paper discusses a comprehensive method for drag-controlled re-entry that is applicable to any vehicle capable of modulating its ballistic coefficient. First, a novel guidance generation algorithm efficient enough to run onboard a CubeSat outputs a desired ballistic coefficient profile and corresponding numerically propagated trajectory that if followed, will lead the spacecraft to a desired de-orbit location. This guidance generation algorithm is based on an analytical solution that provides convergence guarantees, ensures rapid performance, and facilitates a controllability analysis. Next, the guidance tracking algorithm utilizes an extended Kalman filter and GPS measurements to estimate the position and velocity of the satellite relative to the guidance. A full state feedback linear-quadratic-regulator (LQR) control strategy is then used to drive the relative position and velocity to zero using solely aerodynamic drag. This paper also discusses a novel retractable drag de-orbit device (D3) that can be attached to existing CubeSat structures and can easily be scaled up for larger satellites. The D3 provides passive three-axis attitude stabilization using aerodynamic and gravity gradient forces and can be repeatedly modulated to perform aerodynamically-based orbital maneuvering and controlled re-entry. The design of the planned 2U CubeSat to test the D3 and control algorithms in flight is also discussed. The re-entry point targeting algorithms were validated through extensive Monte Carlo simulations which included realistic GPS measurement errors and drag force uncertainties. The algorithms were able to guide the satellite to a desired de-orbit location with an average error below 25 km and in all cases, the targeting error was low enough for debris mitigation purposes. The accuracy and reliability of these algorithms coupled with the D3 device that has successfully undergone thermal vacuum, vibration, and fatigue testing provide a cheap, reliable, and comprehensive attitude, orbit, and de-orbit control solution that can be used on large and small space vehicles, possibly replacing conventional propulsion and attitude control systems and making space more accessible to everyone. • Drag De-Orbit Device (D3) increases the host CubeSat's surface area by.5 m2 • D3 is fully retractable to both expedite de-orbit and provide orbital maneuvering capabilities. • Targeting algorithm specifies D3 deployment levels to guide CubeSat to desired de-orbit location. • CubeSat mission planned to test D3 and control algorithms in space. [ABSTRACT FROM AUTHOR]

Abstract This paper focuses on investigating and analyzing hybrid control strategies of onboard magnetic torque bars (MTs) and reaction wheels (RWs) for attitude stabilization on three-axis stabilized small satellites. The two reaction wheels are mounted on the spacecraft in three different configurations. The proposed attitude control algorithms in this study are based on Lyapunov design approach that almost globally and asymptotically stabilize the spacecraft attitude and this method is not restricted to zero angular momentum assumption that most existing two-wheeled control techniques exploit. Simulation and air bearing testing results presented in the paper demonstrate that the attitude control system can provide successful pointing and tracking in the presence of external disturbances and actuator constraints for a specified class of two-wheeled small satellite. Highlights • Three-Axis attitude control for a specified class of two-wheeled small satellite is presented. • Lyapunov design approach almost globally and asymptotically stabilize the spacecraft attitude. • The active control approaches are evaluated and proven on a three-axis simulator that duplicated spacecraft dynamics. • This method is not restricted to zero angular momentum assumption that most existing two-wheeled control techniques exploit. [ABSTRACT FROM AUTHOR]

The success of the solution to the task of controlling the spacecraft's angular motion around its centre of mass largely depends on the chosen model of the spacecraft's angular motion. In this paper, the spacecraft's angular motion dynamic models, described with second-order quaternionic differential equations, are proposed. The proposed models use only quaternionic calculus. The spacecraft attitude control algorithms' synthesis using these models are considered. Using the example of solving the reorientation problem, it is shown that the use of the proposed models allows one to obtain analytical solutions to problems of synthesis of control algorithms for spacecraft. In this case, the control algorithms are much easier to implement than the algorithms obtained using the traditional model. • Spacecraft state vector consists of Rodrigues–Hamilton parameters and their derivatives. • Proposed models simplify the synthesis of spacecraft attitude control algorithms. • Algorithm synthesis problems have an analytical solution when proposed models are used. • The proposed approach had been experimentally tested in the Egyptsat-1 satellite. [ABSTRACT FROM AUTHOR]

The increase of satellites and rocket upper stages in low earth orbit (LEO) has also increased substantially the danger of collisions in space. Studies have shown that the problem will continue to grow unless a number of debris are removed every year. A typical active debris removal (ADR) mission scenario includes launching an active spacecraft (chaser) which will rendezvous with the inactive target (debris), capture the debris and eventually deorbit both satellites. Many concepts for the capture of the debris while keeping a connection via a tether, between the target and chaser have been investigated, including harpoons, nets, grapples and robotic arms. The paper provides an analysis on the attitude control behaviour for a tethered de-orbiting mission based on the ESA e.Deorbit reference mission, where Envisat is the debris target to be captured by a chaser using a net which is connected to the chaser with a tether. The paper provides novel insight on the feasibility of tethered de-orbiting for the various mission phases such as stabilization after capture, de-orbit burn (plus stabilization), stabilization during atmospheric pass, highlighting the importance of various critical mission parameters such as the tether material. It is shown that the selection of the appropriate tether material while using simple controllers can reduce the effort needed for tethered deorbiting and can safely control the attitude of the debris/chaser connected with a tether, without the danger of a collision. [ABSTRACT FROM AUTHOR]

Retired spacecraft attitude takeover control using two types of cellular space modules, namely orientation cellular space module (OCSM) and rate cellular space module (RCSM), is considered in this paper. Both the OCSM and the RCSM are endowed with identical mechanical actuators for providing 3-axis torques for attitude regulation. The difference is: the OCSM is equipped with orientation sensors which provide unit-length vector measurements to determine attitude, while the RCSM is mounted with rate sensors which measure angular velocities. Correspondingly, the attitude takeover controller is composed of a pre-wrap proportional control part and an output (angular velocity) feedback control part, which are executed by the OCSM and the RCSM respectively. A gain-scheduled control framework is adopted in order to formulate the rate feedback control part as subcontrollers for RCSMs. Two kinds of output feedback subcontrollers are proposed: a static output feedback control scheme with constant gains which are yielded from the linear–differential-game-based output feedback control; and a dynamic output feedback control scheme which is developed using the Kalman–Yakubovich–Popov Lemma and the linear–differential-game-based state feedback control. Both output feedback control schemes possess robustness to modeling errors, since the stability of the closed-loop systems hinge on the passivity theorem rather than the inertia properties of the retired spacecraft. Numerical simulation results show the effectiveness of the proposed controllers for attitude takeover control of a retired spacecraft fixed with the OCSM and the RCSM. A comparison between the proposed methods and an existing method is carried out to show the benefits of the proposed controllers. • Retired spacecraft with inoperable control systems is reused by installing multiple space cellular modules to provide control torques. • The control torques which need to be provided by each space cellular module are calculated from the linear–differential-game-based output feedback control theory and the gain-scheduled passive control framework. • The proposed control schemes possess robustness to modeling errors since the stability of the closed-loop systems hinge on the passivity theorem rather than the inertia properties of the retired spacecraft. [ABSTRACT FROM AUTHOR]

A direct visual-servoing algorithm for control of a space-based two-arm manipulator is proposed in this paper. The algorithm can be utilized in a two-arm manipulators configuration, where one of the arms performs the manipulation and the second arm is dedicated to the observation of the target zone of manipulation. The algorithm utilizes both visual features extracted from onboard cameras and force and torque measured at the manipulator's end-effector to control the movements of the manipulator during on-orbit servicing operations. The algorithm takes into account the relative dynamics of the bodies involved, it relies on images taken independently from de-localized cameras, e.g. at the end-effector of a second manipulator, and it integrates an impedance control for the compensation of eventual contact reactions when the end effector touches and operates the target body. The analytical derivations demonstrate the stability of the algorithm and incorporate an impedance compliance strategy into an optimal framework formulation. Simulations results in two different scenarios have been presented to show the adequate behavior of the presented approach in on-orbit-servicing operations. • We developed a visual servoing controller for a dual arm space manipulator. • The manipulator is driven by images taken by a camera mounted on the other arm. • The controller considers measurements of torques and forces by a sensor on manipulator end effector. • Visual features tracking and tool insertion tasks can be performed by using this controller. [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]

A modified unscented Kalman filter is presented to estimate the quaternion parameters as well as the angular velocities of a rigid gyro-less satellite under faulty sensor conditions. The task is carried out using the Sun sensor and magnetometers as attitude sensors with bounded noise and unknown fault(s). Following the presentation of the satellite attitude dynamics and filtering formulations, a new fault detection and isolation algorithm is proposed. The latter is based on a modified unscented Kalman filter structure for improved fault detection, sensor isolation, and attitude control (AC). A Backtracking Search Algorithm (BSA) is also used to design and optimize the PID controller gains, as well as to optimize sensor installation orientation for residual signal decoupling. In this regard, the paper presents and proves a theory that describes how the proposed optimization process for sensor installation can analytically decouple the faulty orientation signal from the healthy ones. A comparison of attitude and angular velocity estimation approaches using the standard Kalman, extended Kalman, and a proposed modified unscented Kalman filter is also provided. The results show that the proposed filter performs significantly better in closed-loop AC systems and copes adequately with sensor faults. • Gyro-less satellite attitude estimation and control with faulty sensors is studied. • A modified unscented Kalman filter is proposed and compared with KF and EKF. • A novel fault detection and isolation (FDI) algorithm is developed. • FDI is optimized with respect to sensor installation (SI) for improved accuracy. • A theorem is proven that demonstrates the decupling of the residual signals. [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]

This paper studies a robust attitude coordinated control method for spacecraft formation flying system in the presence of external disturbances based on event-triggered control and sliding mode control. To decrease the communication frequency of the spacecraft formation flying system, a sliding mode coordinated controller without Zeno phenomenon is developed by using a novel triggering condition. For the first time, the proposed method addresses the problem that the existence of the closed-loop event-triggered control system solution in the classical sense may not be guaranteed due to the sliding mode surface with discrete variables fundamentally. The theoretical analysis and numerical simulation results demonstrate the effectiveness of the proposed control approach. • A novel method is developed for the theoretical analysis of the complex system based on ETC. • The existence of the closed-loop ETC system solution in the classical sense is guaranteed. • Both attitude consensus and angular velocity consensus are achieved without continuous communication among spacecraft. • A novel triggering condition is established based on exponential function and hyperbolic tangent function. [ABSTRACT FROM AUTHOR]

This paper presents the in-orbit performance of the ESTCube-1 attitude control system that used electromagnetic actuators to achieve a high angular velocity. ESTCube-1 is a one-unit CubeSat that aimed to perform the first electric solar wind sail experiment. The attitude control system was designed to provide enough centrifugal force by spinning up the satellite to deploy a 10 m long tether. The required spin rate was a minimum of one rotation per second. The actuators used were three electromagnetic coils, each able to produce a magnetic moment of up to 0.1 A m 2 . In this paper, we describe the design of the attitude control system, implementation of the spin controller and the in-orbit performance of the system. In addition we describe the effect that a residual magnetic moment had on the attitude control of the satellite and the measures taken to overcome this issue. During testing of the satellite, ESTCube-1 achieved the highest known spin rate of 841°/s for small scale satellites. The satellite ended its operations on the 19th of May, 2015 after 2 years in orbit. [ABSTRACT FROM AUTHOR]

In this paper, a novel algorithm is developed for Nano satellite attitude determination with delayed observations.The particularity of delayed measurements in attitude determination system is analyzed, and then employing a set of Bernoulli distributed random variables to build delayed sensor model. Based on the model, a modified EKF for nonlinear system is deduced by reorganizing the state update equation when data is time stamped or not, which does not increase the dimension of the state or measurement, and avoids the computation of singular matrix inversion. In addition, an adaptive estimation algorithm of measurement noise matrix is combined to reduce the impact of modeling uncertainty. Finally, simulation results demonstrate that the proposed algorithms lead to significantly improved performance compared with the traditional EKF and some delayed filters in the context of networked control systems. • Sensor delay model is established by using a group of Bernoulli random variables. • Filter developed is a modification of EKF, and can apply to delayed data with/without timestamp. • Attitude determination provides up to a twentyfold improvement relative to EKF without timestamp. [ABSTRACT FROM AUTHOR]

This paper investigates attitude dynamic tracking control for spacecraft formation in the presence of unmeasurable velocity information with time-varying delays and switching topology. Based on an extended state observer, a nonlinear attitude tracking control approach is developed for spacecraft attitude model formulated by Euler–Lagrangian equations. The attitude tracking controller allows for external disturbances and absence of angular velocity information. Both auto-stable region techniques and a Lyapunov function approach are developed to prove ultimately bounded tracking. Simulation results demonstrate effectiveness of the nonlinear control techniques proposed in this paper. [ABSTRACT FROM AUTHOR]

This paper studies the coordinative coupled attitude and orbit control system (AOCS) of a satellite formation composed of one microsatellite and two 3U CubeSats. It is described as a highly nonlinear and coupled system subject to actuator saturation and multiple uncertainties including parametric perturbations and external disturbances. To address this issue, the centralized architecture and leader-following method are adopted in the formation firstly. The closed-loop system of each satellite member is divided into orbit and attitude control loops respectively. The coupling relationship between orbit and attitude is introduced as a constraint of the control law. Then a robust controller consisting of a nominal controller, a robust compensator and an advanced proportional–derivative (APD) controller is proposed to achieve high-accuracy formation control. The robust compensator is constructed from an auxiliary system and a finite-time observer to compensate the influence of an equivalent lumped disturbance. The APD controller is developed to achieve desired performances under actuator saturation. Stability analysis indicates that both relative position errors and attitude tracking errors can converge asymptotically. Finally, numerical simulations are carried out to verify the effectiveness of proposed control law in both relatively-stationary and fly-around scenarios. • The coordinative attitude–orbit control system of a satellite formation is formulated. • The coupling relationship is transformed into a constraint of the controller design. • A compensator is constructed from an auxiliary system and a nonlinear observer. • An advanced proportional–derivative controller is developed under actuator saturation. • The proposed control law is easy to implement and suitable for most satellites. [ABSTRACT FROM AUTHOR]

This paper investigates the control issue of the noncooperative spacecraft proximity with prescribed performance by exploiting only relative motion information. To accommodate the direct use of relative measurement, the 6-degree-of-freedom (6DOF) relative dynamics is formulated in the body frame of the pursuer. Based on the smooth Gauss error function, a new asymmetric saturation function is developed to satisfy the practical control limitations. To maintain the relative attitude and position within specified constrained ranges, a novel performance function in possession of the preassigned settling time and convergence set of steady state error is proposed. Based on the backstepping framework, the relative attitude and position controllers are surveyed in sequence. In particular, a state constrained virtual control is first developed for each relative dynamics by introducing a nonlinear transformation. Then a neuroadaptive saturated control law is designed by implementing a Nussbaum function based adaptive control coefficient. Stability analysis indicates that the closed-loop 6DOF relative dynamics are asymptotically stable with prescribed performance. Numerical simulations further verify the proposed control strategy. • A control strategy is developed by using only relative motion information. • A novel performance function with accurate and intuitive settling time is proposed. • An asymmetric saturation function is designed to satisfy the control constraints. • A neuroadaptive algorithm is introduced to ensure the high control accuracy. [ABSTRACT FROM AUTHOR]

In this paper, the problem of controlling the attitude of a CubeSat in low Earth orbit using only the environmental torques is considered. The CubeSat is equipped with a Drag Maneuvering Device (DMD) that enables the spacecraft to modulate its experienced aerodynamic and gravity gradient torques. An adaptive controller is designed to achieve attitude tracking of the spacecraft in the presence of uncertain parameters such as the atmospheric density, drag and lift coefficients, and the time-varying location of the Center of Mass (CoM). The proposed controller also accounts for modeling inaccuracy of the inertia matrix of the spacecraft. A Lyapunov-based analysis is used to prove that the quaternion-based attitude trajectory tracking error is uniformly ultimately bounded. The designed controller is also examined through numerical simulations for a spacecraft with time-varying uncertain drag, lift coefficients and CoM location parameters and the NRLMSISE-00 model for the atmospheric density. • Novel drag and gravity based estimation of unknown parameters of a spacecraft. • Attitude maneuvers without using propellant or momentum exchange devices. • Example based on a real CubeSat mission. [ABSTRACT FROM AUTHOR]

This paper investigates the finite-time attitude tracking control problem for rigid spacecraft exposed to the external disturbance and unknown inertial parameters. Initially, the basic controller is developed under the assumption that the inertial parameters and the upper bound of external disturbances are available to the designer. However, this assumption is inapplicable in practical applications. Considering this limitation, adaptive laws and Minimum-learning- parameter (MLP) algorithm are employed in the second controller to estimate the unknown system dynamics. Finite-time stability for tracking errors is achievable under the proposed methods. Though theoretical analysis and numerical simulations, effectiveness of the proposed controllers has been validated. • Two finite-time attitude tracking controllers are proposed. • The proposed controller is inertia-free for implementation. • MLP Algorithm has been employed to reduce the computational burden. [ABSTRACT FROM AUTHOR]

The orientation of reaction wheels and their individual maximum torque capacity limits the agility of a rigid body. However, maximizing the isotropic torque envelope can underestimate the true capability of the system because the inertia properties are not fully considered. This is true even when performing eigenaxis maneuvers with isotropic acceleration and rate limits. Improvement in agility is possible when compared to conventional principal axis orientation of three orthogonal reaction wheels. Reorienting the reaction wheels while maintaining orthogonality can increase agility available for slew maneuvers. This paper uses the agility envelope as a tool in developing analytical equations for the optimal reaction wheel orientation to maximize agility, which can be up to 36% more than conventional methods. The equations and corresponding design curves are also presented for any system based on the normalized Inertia Ratios. • Conventional methods underestimate actual spacecraft agility performance. • Reorientation of the reaction wheels increase agility up to 36%. • Eigenaxis maneuvering maintained with increased performance. • Design curves are presented to aid in agility design trades. [ABSTRACT FROM AUTHOR]

The airborne lunar spectral irradiance mission is an inter-agency partnership between the US National Aeronautics and Space Administration and the US National Institute of Standards and Technology that aims to make SI-traceable measurements of lunar spectral irradiance at visible to near-infrared wavelengths with unprecedented accuracy. This information is vital to using the Moon as a calibration source for Earth observing satellites. To minimize uncertainty, the lunar measurements are made above 90% of the Earth's atmosphere from an Earth Resources 2 aircraft, a civilian descendant of the U-2 spy plane. Situated in a large wing pod, a custom-designed telescope automatically tracks the Moon and the measurements are fed into a spectrometer. This information is being used to develop an extremely accurate model that can be used to calibrate satellites. An Engineering Flight Campaign was completed in August 2018 and a Demonstration Flight Campaign in November 2019, which demonstrated autonomous lunar acquisition and tracking as well as measurements of the Moon's spectral irradiance from an altitude of approximately 21 km. This article presents the simplified double gimbal control system design that was used to manipulate the telescope, and was capable of targeting the Moon with a root mean squared tracking error of about 0. 1 ∘ . • This paper presents the results of a robotic system used for NASA science flights. • Robotic system autonomously tracks the Moon to obtain lunar spectral measurements. • Information is used by NASA and NIST to calibrate Earth-observing satellites. • Operates in extremely low pressures and temperatures at an altitude of 21 km. • Capable of targeting the Moon with a root mean squared error under 0.11 degrees. [ABSTRACT FROM AUTHOR]

Target capture is a fundamental and essential task of space robots for on-orbit servicing in the future. Accordingly, it is important to study the method for space robot stabilization after impact during target capture. In this paper, the stabilization control problem of a combined spacecraft after multiple impacts during target capture is studied. A dual-integral sliding mode (DISM) control method based on the reconfiguration of the reaction wheels is proposed. The stability control of the combined spacecraft is achieved by controlling the reaction wheels of the space robot to absorb the angular momentum arising from the impact. First, given the structure of the gripper, the contact detection algorithm and contact dynamics model are established, and then multiple impacts are simulated during the target capturing operation. Second, an improved sliding mode control method is proposed to overcome the influence of multiple impacts, and the stability of the controller is confirmed by the Lyapunov method. Third, through an optimization method, the configuration of the reaction wheels of the space robot is reconstructed to avoid control torque saturation. The validity and availability are eventually verified by simulations and experiments. The results show that the proposed method can achieve the stability of the combined spacecraft after the target capture, and the algorithm is robust and has engineering applicability. • The contact detection algorithm and contact dynamics model are established. • A dual-integral sliding mode control method is proposed to overcome the influence of multiple impacts. • The configuration of the reaction wheels of the space robot is reconstructed to avoid control torque saturation. [ABSTRACT FROM AUTHOR]

This paper investigates the attitude maneuver problem of a spacecraft containing internal movable masses with zero angular momentum. The spacecraft dynamics in one dimension is first studied using the proposed moving mass control system with predefined mass motion. Next, a small-angle attitude maneuver method based on the sliding mode control technique is designed to accomplish the reorientation when the required maneuver angle is within the control capability of single-direction movable mass motion. Furthermore, a novel mass-shifting procedure is proposed for achieving full-domain attitude reorientation. The procedure includes a rest-to-rest maneuver method together with the trajectory optimization. Which, the rest-to-rest maneuver takes advantage of the attitude difference caused by sequential mass motion and the trajectory optimization is carried out based on time sub-optimal motion planning methods. The procedure is next combined with the small-angle maneuver method to achieve a high precision reorientation. Finally, numerical simulation results show that high precision full-domain attitude reorientation can be realized using the proposed control system with three internal movable masses. • Feasibility of full-domain attitude reorientation using internal movable masses is investigated for the first time. • A novel control procedure using internal movable masses to achieve full-domain attitude reorientation is proposed. • A high precision reorientation method in a combination of sequence rotation with small angle maneuver is presented. [ABSTRACT FROM AUTHOR]

This paper presents an adaptive unscented Kalman filter (AUKF) to recover the satellite attitude in a fault detection and diagnosis (FDD) subsystem of microsatellites. The FDD subsystem includes a filter and an estimator with residual generators, hypothesis tests for fault detections and a reference logic table for fault isolations and fault recovery. The recovery process is based on the monitoring of mean and variance values of each attitude sensor behaviors from residual vectors. In the case of normal work, the residual vectors should be in the form of Gaussian white noise with zero mean and fixed variance. When the hypothesis tests for the residual vectors detect something unusual by comparing the mean and variance values with dynamic thresholds, the AUKF with real-time updated measurement noise covariance matrix will be used to recover the sensor faults. The scheme developed in this paper resolves the problem of the heavy and complex calculations during residual generations and therefore the delay in the isolation process is reduced. The numerical simulations for TSUBAME, a demonstration microsatellite of Tokyo Institute of Technology, are conducted and analyzed to demonstrate the working of the AUKF and FDD subsystem. [ABSTRACT FROM AUTHOR]

This paper addresses the problem of adaptive sliding mode attitude tracking control for a class of nonlinear flexible spacecraft systems with a redundant four reaction wheels' setting and unknown actuator dead-zone, where configuration misalignment and external disturbance are considered simultaneously. By using the Takagi-Sugeno (T-S) fuzzy modelling method, the overall nonlinear spacecraft attitude dynamics system is first reconstructed into several local linear systems. Then, an adaptive integral-type sliding mode control scheme is introduced based on a T-S fuzzy model to stabilize the considered attitude tracking control system of flexible spacecraft. In this design, the external disturbance with unknown bound is counteracted in real-time by the adaptive estimate technique. Additionally, the integral sliding surface we introduced is insured to be asymptotically stable under the condition of given matrix inequality. The proposed adaptive sliding mode controller guarantees the finite-time reachability of the specific sliding surface. A practical example with simulation results is given to verify the validity of the attitude tracking control strategy we put forward. • The flexible spacecraft attitude tracking control problem is investigated. • Configuration misalignment and input dead-zone are considered simultaneously. • A T-S fuzzy adaptive sliding mode control strategy is proposed. • The global asymptotical stability of attitude tracking errors is guaranteed. [ABSTRACT FROM AUTHOR]