Attitude Stability Of Satellite Research Articles

Rapid Attitude Stabilization of Ultra-Low Orbit Satellites Using Movable Masses and Reaction Wheels

This paper proposes a combination of movable masses and reaction wheels for the attitude control of ultra-low-orbit satellites. During the attitude control of ultra-low-orbit satellites, the satellite’s attitude is significantly influenced by the aerodynamic torque. Therefore, this paper proposes a combined control method using movable masses and reaction wheels. By adjusting the position of the movable masses to modify the relative distance between the center of mass (CM) and the center of pressure (CP), the aerodynamic torque is effectively reduced. Additionally, a finite-time terminal sliding mode extended state observer (TSMESO) is incorporated to estimate the reduced aerodynamic torque. Then, a finite-time non-singular terminal sliding mode control (NTSMC) method is introduced to achieve rapid attitude stabilization. Simulations demonstrate that under the proposed control system, the aerodynamic torque acting on the ultra-low-orbit spacecraft can be significantly reduced in a short time, and the response and high precision attitude stability can be effectively achieved.

Velocity-Free State Feedback Fault-Tolerant Control for Satellite with Actuator and Sensor Faults

A velocity-free state feedback fault-tolerant control approach is proposed for the rigid satellite attitude stabilization problem subject to velocity-free measurements and actuator and sensor faults. First, multiplicative faults and additive faults are considered in the actuator and the sensor. The faults and system states are extended into a new augmented vector. Then, an improved sliding mode observer based on the augmented vector is presented to estimate unknown system states and actuator and sensor faults simultaneously. Next, a velocity-free state feedback attitude controller is designed based on the information from the observer. The controller compensates for the effects of actuator and sensor faults and asymptotically stabilizes the attitude. Finally, simulation results demonstrate the effectiveness of the proposed scheme.

Integrated Anti-Windup Fault-Tolerant Control Architecture for Optimized Satellite Attitude Stabilization

This article presents the development of an anti-windup (AW) compensator architecture for nanosatellite systems. The design addresses the problem of fault-tolerant control (FTC) for attitude stabilization subject to actuator saturation, actuator faults, and multiple disturbances. In many practical applications, certain physical limitations in the actuator design often mean that actuator saturation is mostly unaccounted for. Based on the ESTCube-2 nanosatellite model with reaction wheels (RWs), an integrated AW–FTC architecture is designed for an optimized attitude control system. The AW compensator is formulated using the linear matrix inequality (LMI) and computed to generate an augmented solution that is appended to the linear control system. The AW compensator in the architecture is designed with suitable parameters and effective controller gain values. The AW–FTC acts on the RW momentum as control input with the observed torque response, thereby ensuring an asymptotically stable system. Simulation results are presented to demonstrate the effectiveness of this approach with finite-time stabilization under actuator faults and saturation conditions.

Natural Magneto-velocity Coordinate System for Satellite Attitude Stabilization: The Concept and Kinematic Analysis

An artificial Earth satellite with an electric charge and an intrinsic magnetic moment is considered. Due to the geomagnetic field, the satellite experiences the influence of the Lorentz and magnetic torques. To set the angular position of the satellite, we introduce natural coordinate system associated with the directions of geomagnetic induction vector and Lorentz force vector which is orthogonal both to the geomagnetic induction and relative velocity of the satellite. It is shown that such a natural magneto-velocity coordinate system is convenient for attitude stabilization of a satellite operating in the mode of scanning the Earth's surface. The properties of the trajectory of the satellite axis on the Earth's surface are analysed. The rotation tensor connecting the natural magneto-velocity and the orbital coordinate systems is obtained. The angular velocity of the natural magneto-velocity trihedron is found. Kinematic differential equations for the unit vectors of the natural magneto-velocity coordinate system are derived.

Attitude stabilization of slosh-coupled satellite with flexible appendage

In this paper, a robust control scheme is proposed for flexible satellite attitude stabilization in the presence of multiple disturbances including the disturbances caused by flexible appendages with the sloshing dynamics of the liquid propellant and external disturbances from space environment. A composite control scheme is designed by combining a composite disturbance observer (CDO) with feed-forward controller. First, a CDO is constructed to estimate the lumped disturbances, where the lumped disturbances in the system are equivalent into two parts: a slow varying disturbance and periodic disturbance. Then a PD controller is adopted to stabilize the system and restrain the residue disturbance. Finally, the simulation results for flexible satellite verify that the proposed method can make the satellite asymptotically stable and demonstrate the effectiveness of the proposed scheme.

Mass-Ejecting Reaction Wheel for CubeSat Small Orbit Corrections

CubeSats usually carry no propulsion system because of difficulties in scaling down existing propulsion technologies to meet the stringent size, volume, and power limits. The most challenging case in this context is the one-unit CubeSat: the use of which has been rapidly growing. To increase CubeSat functionality by allowing small orbit corrections, particularly due to orbit injection errors, this paper develops a new propulsion method relying on an ejection of masses from a reaction wheel. Using the reaction wheel’s kinetic energy, propellant masses are released from the wheel’s outer circumference using an impulsive application of an electric current. This paper models the dynamics of the satellite, the modified reaction wheel, and the ejected masses; evaluates the pros and cons of the proposed method; and provides a methodology for constructing a model of a mass-ejecting reaction wheel designed according to given specifications of a desired velocity change and satellite attitude stability. The system is useful for small orbit corrections, which can be used (for example) for CubeSat formation establishment after orbital injection.

Fault-tolerant control allocation of a flexible satellite with an infinite-dimensional model

Fault-tolerant control allocation (FTCA) strategy is proposed for attitude stabilization of a flexible satellite with actuator redundancy. The control scheme is based on the infinite-dimensional model of a flexible satellite with no discretization, so the spillover instability is eliminated. This is one of the important benefits of the proposed control scheme over the previous FTCA schemes that have been used for the flexible satellite. The proposed scheme contains two modules. The first module provides a virtual control law to meet stabilization and vibration control objectives in the presence of uncertainties and external disturbances. There is no need to implement in-domain actuators on panels to stabilize their vibration. In this module, the virtual control is designed using adaptive integral sliding mode approach where the sliding surface includes angular velocities, internal reaction torques, and nominal control for healthy system. The second module, based on fault/failure information and using a control allocation scheme, provides redistribution of the virtual control law among the available actuators. Due to simultaneous actuator faults and control constraints, there is an error between the actual virtual control and the designed control that affects the overall system stability. To eliminate this error, gain of the virtual control signal is adjusted by an adaptive updating law. The closed-loop system stability is guaranteed for small changes in a neighborhood of the sliding surface with simultaneous vibration damping. A numerical example illustrates the effectiveness of the proposed control strategy.

Averaging technique in the problem of satellite attitude stabilization in indirect position in the orbital reference frame with the use of Lorentz torque

A dynamically symmetric satellite in a circular orbit of small inclination is considered. The problem of the Lorentzian stabilization of the satellite in the orbital coordinate system in the indirect equilibrium position is solved under the conditions of the perturbing effect of the gravitational torque. To solve this problem, which is characterized by incomplete control, a method of averaging differential equations is developed. Using the original construction of the unsteady Lyapunov function, we obtain sufficient conditions for the asymptotic stability of the programmed satellite motion mode in the form of constructive inequalities with respect to the control parameters.

Adaptive fuzzy Jacobian control of spacecraft combined attitude and Sun tracking system

PurposeThe 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.Design/methodology/approachTo 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.FindingsNumerical results show the effectiveness of the proposed adaptive control scheme in simultaneously tracking the desired attitude and the Sun.Practical implicationsReaction 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.Originality/valueThe proposed Jacobian control scheme is inspired by the knowledge of Jacobian matrix in the trajectory tracking of robotic manipulators.

Averaging technique in the problem of Lorentz attitude stabilization of an Earth-pointing satellite

An Earth-pointing satellite with Lorentz attitude control system is under consideration. The problem of the satellite attitude stabilization in the orbital reference frame in the presence of disturbances and complicated by underactuation is studied. The averaging technique is developed and applied in the problem. A rigorous mathematical justification is proposed, which makes it possible to consider a satellite with unequal moments of inertia and to obtain specific estimates for the control parameters that provide a solution to the problem of satellite attitude stabilization with the use of the Lorentz control torque. The octupole approximation of the Earth's magnetic field is used; the periodic functions of time arise in geomagnetic field induction. With the aid of nonlinear approach to the problem it is proved that Lorentz attitude control system ensures asymptotic stability of the satellite equilibrium under the perturbing action of disturbing torques provided dissipative torque is operating alongside with the restoring one. The proof is based on the Lyapunov direct method and the averaging technique. In the paper, original constructions of nonstationary Lyapunov functions are proposed. These constructions take into account the structure of perturbations acting on the considered system. With the aid of the proposed Lyapunov functions, sufficient conditions for the asymptotic stability of the satellite equilibrium position are derived. These conditions are formulated in terms of explicit inequalities on the control parameters. Thus, a constructive approach to the design of stabilizing controls is formulated.

Observer-Based Backstepping Fault-tolerant Control for Spacecraft with Reaction-wheel Failures

In this paper, nonlinear satellite system with actuator faults is considered. To solve the problem of attitude stabilization of satellite under external disturbances and actuator faults, an observer is designed firstly. And the output of the observer is proved to be uniformly ultimately bounded. By using the estimated information of the observer, a backstepping fault-tolerant controller is derived. The proposed controller could achieve attitude stabilization in the presence of actuator faults. Finally, a simulation of attitude stabilization is carried out to illustrate effective of the fault-tolerant controller

Deployment of flexible space tether system with satellite attitude stabilization

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 J2 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.

Reinforcement Learning-Based Satellite Attitude Stabilization Method for Non-Cooperative Target Capturing.

When a satellite performs complex tasks such as discarding a payload or capturing a non-cooperative target, it will encounter sudden changes in the attitude and mass parameters, causing unstable flying and rolling of the satellite. In such circumstances, the change of the movement and mass characteristics are unpredictable. Thus, the traditional attitude control methods are unable to stabilize the satellite since they are dependent on the mass parameters of the controlled object. In this paper, we proposed a reinforcement learning method to re-stabilize the attitude of a satellite under such circumstances. Specifically, we discretize the continuous control torque, and build a neural network model that can output the discretized control torque to control the satellite. A dynamics simulation environment of the satellite is built, and the deep Q Network algorithm is then performed to train the neural network in this simulation environment. The reward of the training is the stabilization of the satellite. Simulation experiments illustrate that, with the iteration of training progresses, the neural network model gradually learned to re-stabilize the attitude of a satellite after unknown disturbance. As a contrast, the traditional PD (Proportion Differential) controller was unable to re-stabilize the satellite due to its dependence on the mass parameters. The proposed method adopts self-learning to control satellite attitudes, shows considerable intelligence and certain universality, and has a strong application potential for future intelligent control of satellites performing complex space tasks.

Time efficient sliding mode controller based on Bang–Bang logic for satellite attitude control

In order to improve the convergence rate of standard sliding mode controller, time efficient controllers based on Bang–Bang logic for satellite attitude stabilization and tracking control are developed in this paper. The time efficient open-loop control algorithm Bang–Bang control is combined with closed-loop sliding mode control to improve system robustness. A two-stage structure sliding mode with a fixed angular velocity stage and a fixed deceleration stage is proposed in this paper. The sliding mode parameter is real-time updating hence the modified sliding mode could have Bang–Bang character. The system inertia matrix uncertainty and disturbance torque is discussed and the controller proposed in this paper is robust to the perturbation. The control torque constraint is also discussed and the constraint on control parameters is given. The performance of the controller is demonstrated by numerical simulation.

Robust finite time control algorithm for satellite attitude control

Finite time controllers robust to inertia matrix uncertainty for satellite attitude stabilization control and attitude tracking control are developed in this paper. A three-stage finite sliding mode is constructed to improve system convergence rate. The singularity issue is solved by using the property of Euler Axis when it's paralleled to angular velocity. Based on this finite time sliding mode, finite time controller is developed to ensure the system state could reach the sliding mode in finite time. System inertia matrix uncertainty and disturbance torque is considered in this paper. The control torque constraint is also considered to ensure the norm of control torque does not exceed system upper bound. Finite time stability of the controller is proved and the controller performance is demonstrated by simulation results.

Time Efficient Robust PID Plus Controller for Satellite Attitude Stabilization Control Considering Angular Velocity and Control Torque Constraint

AbstractA time efficient robust proportional, integral, differential (PID) plus controller for satellite attitude stabilization control is proposed in this paper. A time efficient sliding mode with...

Attitude Stabilization Control of Flexible Satellites With High Accuracy: An Estimator-Based Approach

This paper presents a novel nonlinear control approach for attitude stabilization of flexible satellites. It is developed by using a nonlinear estimator technique. An estimator is proposed to estimate external disturbances and unknown flexible vibrations acting on the satellite. Applying the estimated value, a compensation control effort that compensates for such effects in real-time is derived and added to the nominal controller to accomplish the desired attitude maneuver. The proposed controller guarantees the closed-loop system to be asymptotically stable, and provides high-accuracy control, external disturbance rejection, and robustness to flexible vibrations. The key feature of this approach is that the controller has no conservativeness. The effectiveness of the approach is verified through both simulation and experimental results.

Three-axis attitude stabilization of a flexible satellite using non-linear PD controller

In this paper, after complete modelling of a flexible satellite equipped with a control moment gyroscope (CMG) actuator, it is shown that a PD-like controller can globally asymptotically stabilize this satellite by using Lyapunov’s direct method. Despite the simplicity, simulations show that the controller can stabilize the flexible satellite in a three-axis manoeuvre even in the presence of external disturbances. Then, using a non-linear variable gains PD controller, which only uses angular velocity of the rigid body and the attitude parameters as the inputs, the performance of the control system is improved in some important aspects such as reducing maximum control torque, reducing maximum peak of deflection of the appendages and increasing robustness of the controller against the orbital disturbances. In addition, locally asymptotically stability of the non-linear variable gain PD controller is guaranteed using a novel Lyapunov candidate function. Considering the difficulty in measuring the appendages’ deflection and the primarily existence of parameter uncertainties, and as this controller is independent of changes in these parameters, such a control system is very useful and applicable. In order to validate the system’s mathematical model and the control system performance, an exact model of the satellite is constructed in the ADAMS/View software that is linked to the MATLAB software. The efficacy of the proposed approach is demonstrated by several numerical examples.

Velocity-Free Fault-Tolerant and Uncertainty Attenuation Control for a Class of Nonlinear Systems

This paper addresses a difficult problem of velocity-free uncertain attenuation control for a class of nonlinear systems with external disturbance and multiple actuator faults. With only the output measurement available for feedback, a sliding-mode observer (SMO) is proposed to reconstruct the full states. The reconstructed signal can approximate the true value to any accuracy. An adaptive version of the observer is further presented to handle a class of structured uncertainties in the system. Together with the system output feedback, the reconstructed state is used to synthesize a velocity-free controller. All states in the closed-loop system are guaranteed to be uniformly ultimately bounded (UUB). System uncertainty and external disturbances are attenuated. Actuator fault is accommodated. An example with application the approach to satellite attitude stabilization maneuver is presented to verify the effectiveness of the proposed scheme.

Methodology to Analyze Attitude Stability of Satellites Subjected to Aerodynamic Torques

A method to analyze the stability of satellites flying in low Earth orbits is presented. Starting from the classical rotational equations of motion, including gravity gradient, the influence of aerodynamic torques is studied both in the pitch and roll–yaw axes. Applying concepts used to assess aircraft stability, such as stability derivatives, several ideas and recommendations are presented. The work aims to link geometrical properties with attitude response, providing a way to design satellites that actively takes into account environmental torques, either by minimizing them or by utilizing them to meet mission goals.