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The paper aims to address the combined attitude control and Sun tracking problem in a flexible spacecraft in the presence of external and internal disturbances. The attitude stabilization of a flexible satellite is generally a challenging control problem, because of the facts that satellite kinematic and dynamic equations are inherently nonlinear, the rigid–flexible coupling dynamical effect, as well as the uncertainty that arises from the effect of actuator anomalies.

To deal with these issues in the combined attitude and Sun tracking system, a novel control scheme is proposed based on the adaptive fuzzy Jacobian approach. The augmented spacecraft model is then analyzed and the Lyapunov-based backstepping method is applied to develop a nonlinear three-axis attitude pointing control law and the adaptation law.

Numerical results show the effectiveness of the proposed adaptive control scheme in simultaneously tracking the desired attitude and the Sun.

Reaction wheels are commonly used in many spacecraft systems for the three-axis attitude control by delivering precise torques. If a reaction wheel suffers from an irreversible mechanical breakdown, then it is likely going to interrupt the mission, or even leading to a catastrophic loss. The pitch-axis mounted solar array drive assemblies (SADAs) can be exploited to anticipate such situation to generate a differential torque. As the solar panels are rotated by the SADAs to be orientated relative to the Sun, the pitch-axis wheel control torque demand can be compensated by the differential torque.

The proposed Jacobian control scheme is inspired by the knowledge of Jacobian matrix in the trajectory tracking of robotic manipulators.

The satellite pointing accuracy plays a crucial role in ensuring a successful satellite mission itself. Therefore, this paper aims to enhance the attitude pointing accuracy of the combined energy and attitude control system (CEACS) in a satellite in the presence of external disturbance torques through a robust controller, which can produce high pointing accuracies with smaller control torques.

To improve the CEACS attitude pointing accuracy, a maiden fuzzy proportional derivative (PD)-based CEACS architecture is proposed. The mathematical models along with its numerical treatments of the fuzzy PD-based CEACS attitude control architecture are presented. In addition, a comparison between the PD and fuzzy PD controllers in terms of the CEACS pointing accuracies and control torques is provided.

Numerical results show that the fuzzy PD controller produces a considerable CEACS pointing accuracy improvement for a lower control torque compartment.

CEACS has gained a renew interest because of significant increase in the projected onboard power requirements for future space missions. Therefore, it is of paramount importance to improve the CEACS pointing accuracy itself with a minimum control torque compartment. In fact, this proposed fuzzy PD controller is shown to be a potential CEACS attitude controller.

The fuzzy PD-based CEACS architecture not only provides a better attitude pointing accuracy but also ensures a lower control torque compartment, which corresponds to a lower onboard power consumption.

This paper aims to investigate the attitude control pointing improvement for a small satellite with control moment gyroscopes (CMGs) using the active force control (AFC) method.

The AFC method is developed with its governing equations and integrated into the conventional proportional-derivative (PD) controller of a closed-loop satellite attitude control system. Two numerical simulations of an identical attitude control mission namely the PD controller and the PD+AFC controller were carried out using the MATLAB^®-SimulinkTM software and their attitude control performances were demonstrated accordingly.

Having the PD+AFC controller, the attitude maneuver can be completed within the desired slew rate, which is about 2.14 degree/s and the attitude pointing accuracies for the roll, pitch and yaw angles have improved significantly by more than 85% in comparison with the PD controller alone. Moreover, the implementation of the AFC into the conventional PD controller does not cause significant difference on the physical structure of the four single gimbal CMGs (4-SGCMGs).

To achieve a precise attitude pointing mission, the AFC method can be applied directly to the existing conventional PD attitude control system of a CMG-based satellite. In this case, the AFC is indeed the backbone for the satellite attitude performance improvement.

The present study demonstrates that the attitude pointing of a small satellite with CMGs is improved through the implementation of the AFC scheme into the PD controller.

This paper aims to describe a design enhancement for the satellite attitude control system using reaction wheels, and the wheel momentum unloading using magnetorquers.

The proportional – integral–derivative-controller and active force control (AFC) schemes are developed together with their governing equations for closed loop system of attitude control. Four numerical simulations were carried out using the Matlab – Simulink™ software and results were compared.

From the results, it is evident that the attitude accuracies for roll–pitch–yaw axes have improved significantly through the proportional – derivative (PD) – AFC controller for the attitude control and the wheel momentum can be well maintained during the momentum unloading scheme. The results show that the AFC has a high potential to be implemented in the satellite attitude control system.

Using AFC, the actual disturbance torque is considered totally rejected by the system without having to have any direct prior knowledge on the actual disturbance itself.

The results demonstrate the satellite attitude control using reaction wheel is enhanced by PD–AFC attitude controller.

The purpose of this paper is to develop a theoretical design for the alternative attitude control of the rotation about the pitch axis for the nadir-pointing spacecraft in the event of inertial actuator faults.

This paper presents a novel and viable solution to that problem using the combined attitude and sun tracking system (CASTS) that was conceived from an engineering problem-solving toolkit called TRIZ. Linear and fuzzy controllers are used to test the spacecraft CASTS architecture. All the relevant governing equations of the control system and disturbance rejection methods are developed.

The performance of the proposed CASTS control strategy is tested through numerical simulations. The results strongly suggest that the novel proposed control scheme is effective and promising for controlling the satellite attitude and sun tracking simultaneously in the presence of disturbance torques.

This work is mainly focused on the rigid body of the spacecraft hub that contains all attitude control hardware and payload instrumentation, and does not deal with the vibrations evolving from the propellant sloshing and large flexible appendages such as the deployable solar panels and synthetic aperture radar antennas.

The results from this work reveal several practical applications worthy of reducing the weight, size of the spacecraft and, therefore, cost of missions while increasing the instrumentation capabilities.

The proposed CASTS solution is a result of looking much wider than one system from a new combination of attitude control and sun tracking, as well as innovative ways of using it.

An architecture is proposed based on the torque mode operation for the combined energy and attitude control system (CEACS). The CEACS energy storage and attitude control performances are demonstrated.

All the relevant system equations are established. The torque mode‐based architecture is designed. The architecture is evaluated through numerical treatments.

The proposed CEACS architecture can simultaneously manage the energy storage and attitude control tasks.

This research work is exclusively for small satellites in the LEO missions. However, the proposed CEACS architecture is applicable for other missions.

The torque mode operation will allow an ironless motor/generator design for CEACS.

The results demonstrate the CEACS operation in the torque mode.

The purpose of this paper is to develop analytic solutions for a tethered satellite system (TSS) subjected to internal tether tension moment and external aerodynamic torque for spin-up and spin-down manoeuvres.

Analytic solutions for TSS based on the approximation of Euler’s equations of motion via Fresnel integrals and sine and cosine integrals. Test simulation was performed for two cases (spin-up and spin-down manoeuvres). The conclusion is based on graphical interpretation.

The effects of angular velocities on X, Y and Z axes of the TSS under the influence of combined torques from internal tether tension and external aerodynamic drag influenced during spinning manoeuvres are shown graphically.

This research focuses only on a circular orbit, which is one of the simplest orbits without many variables taken into account such as flight path angle and true anomaly. It could get quite complex for other orbit types like elliptic and parabolic orbits.

Practical implications include observing the stability rotational motion of TSS so as to perform a two-way payload exchange via momentum transfer.

In this paper, analytic solutions for a torque motion of a TSS comprising non-linear Euler’s equations of motion are established for spin-up and spin-down manoeuvres.

Flywheels can serve not only as attitude control devices, but also as energy storage devices, thereby eliminating the need for conventional batteries. Hence, a combined energy and attitude control system (CEACS) consisting of a double counter rotating flywheel assembly is proposed for small satellites in this paper. The energy level in CEACS depends mainly on the flywheels' speeds. Therefore, a specific flywheel energy management strategy has to be implemented to take into account the limitations of the flywheels, which has not been established until today.

To compare the conventional reaction wheel and battery systems with the combined energy and attitude control system. The system mass, volume and power requirements are revealed corresponding to the small satellite missions.

All the relevant system parametric equations are established. The system mass, volume and power are estimated accordingly for the conventional and the combined systems. Then, both systems are compared with respect to the typical small satellite missions.

The combined system outperforms the conventional system in most small satellite missions. However, there are some small satellite missions where the conventional systems are better in terms of the mass and volume budgets.

This research work is exclusively for small satellites in the LEO orbits.

A reasonable information for sizing the combined energy and attitude control system is established. The system mass, volume and power budgets can be extracted from this research work.

The results provide an instant answer regarding the feasibility of the combined energy and attitude control system for small satellites.