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Based on the satellite attitude control method, this paper proposes an attitude control method based on neural network disturbance compensation. The paper firstly analyzes the neural network algorithm and proposes an orthogonal least squares algorithm to implement network learning. In this paper, a set of high-precision directional neural network compensation controllers is designed for the attitude control of acupuncture small satellites. The feasibility of the improved orthogonal least-squares algorithm combined with the neural network supplementary control method in satellite attitude control is verified by experiments. [ABSTRACT FROM AUTHOR]

This paper uses adaptive backstepping sliding mode control to synchronize two satellite attitude dynamics with perturbing torques. The external perturbing torques induce chaotic motion with no control inputs. The proposed control system uses Lyapunov theory and Barbalat's Lemma to guarantee the asymptotic stability of the controlled system. Simulation results confirm the effectiveness of the proposed design.

This work presents an attitude control strategy for Earth-imaging satellites. The proposed methodology is based on the closed-loop implementation of a direct method and a nonlinear programming solver for the quick determination of near-optimal attitude trajectories. The two optimization criteria considered are the target acquisition time and the maneuver energy associated to the actuation torques. For comparison purposes, we take under analysis alternative attitude controllers, such as a quaternion feedback Proportional-Derivative controller, a Linear Quadratic Regulator, a Model Predictive Control, and an adaptive Continuous Twisting Sliding Mode Controller. Optimality and accuracy of these controllers are assessed through simulations in a high-fidelity environment inclusive of orbital perturbations and refined model of the actuation device. Analysis of the results emphasizes the advantages of the proposed methodology over the alternative controllers. [ABSTRACT FROM AUTHOR]

In this paper, a variable structure PID controller with a good convergence rate and robustness for satellite attitude is proposed. In order to improve the system convergence rate, the variable structure for the proportional and differential term was designed, and an angular velocity curve with a better convergence rate was achieved by this variable structure. In addition, an integral partitioning algorithm was designed, and the system robustness to disturbance torque was improved; meanwhile, the negative effect of the integral term was avoided during the converging process. The actuator failure condition was also considered, and a fault tolerant control algorithm was designed. System stability was analyzed by the Lyapunov method, and its performance was demonstrated by numerical simulation. [ABSTRACT FROM AUTHOR]

Several novel control techniques have been created as a result of the diversity of researches which are conducted about the problem of satellite attitude control. There are always uncertainties in the problem of satellite attitude control in the space missions. Therefore, Adaptive control is a method which is taken into consideration. High computational volume is one of the problems of adaptive control technique. In this paper, a control technique which is based on optimization concepts is introduced for the problem of satellite angular velocity and attitude control. Also, it's developed based on the three-dimensional special orthogonal group, and it's not faced by a singularity problem. For comparison, the linear quadratic regulator (LQR) control technique is simulated. Finally, the results of the simulations show that the performance of the presented adaptive control technique is optimal, and this method is robust to inertia changes.

This paper describes the implementation of fast state-dependent Riccati equation (SDRE) control algorithms through the use of shallow and deep artificial neural networks (ANN). Several ANNs are trained to replicate an SDRE controller developed for a satellite attitude dynamics simulator (SADS) to display the technique's efficacy. The neural controllers have reduced computational complexity compared with the original SDRE controller, allowing its execution at a significantly higher rate. One of the neural controllers was validated using the SADS in a practical experiment. The experimental results indicate that the training error is sufficiently small for the neural controller to perform equivalently to the original SDRE controller. • A real-time SDRE controller is designed using shallow and deep neural networks. • A significant reduction in computation is achieved using the neural controllers. • Deep denoising autoencoders were trained as deep neural controllers. • The neural controller is validated through simulations and a practical experiment. • Results show that the neural controller retains its high performance. [ABSTRACT FROM AUTHOR]

In this paper, damping mode of a satellite attitude control is designed and implemented using magnetic actuators in software /hardware-in-the-loop testbed. To this end, the equivalent of Earth’s magnetic field is designed using Helmholtz coil, frictionless is made by air-bearing, and algorithms are developed on designed control board. By measuring the Earth’s magnetic field, actuator commands are generated by the damping algorithm then braking torque is produced. Some applied restrictions and special requirements such as non-simultaneous operation between magnetic sensor and magnetic actuators, air-bearing friction, initial angular velocity are considered. By identifying the air-bearing frictional model, the results are compared in software/hardware-in-the-loop. The compared results show that the ability of the designed system to perform damping mode.

This study uses comprehensive pointing control analysis to calculate laser pointing errors for nano-satellite laser communication systems. This methodology especially enables the performance evaluation of the pointing system during the satellite pointing budget design phase. The analysis consists of a two-stage pointing approach: a body pointing stage of a satellite and a fine pointing stage within the optical payload. By simplifying each pointing stage to a single-axis model, we derive their error transfer functions, which are converted into power spectral density (PSD). The PSD of the two stages are combined using linear interpolation due to the difference in their sampling frequencies. This process yields the final laser pointing error PSD. Our method is validated against Monte-Carlo simulation results and control experiments of the fine pointing stage. The simulation and the experimental results maintained correspondence with the modeled power spectral density. These comparisons can suggest the proposed approach can predict the pointing performance of the system. Based on these results, a method to determine controller gains is introduced, aiming to minimize final pointing errors. • It presents frequency domain analyses for pointing errors of laser communication satellite. • Upsampling combines power spectral density in two types of digital system transfer functions. • It offers system design insights by the relationship between satellite attitude and pointing. [ABSTRACT FROM AUTHOR]

In this paper, the design and implementation of hardware-in-the-loop (HIL) test-bed for spacecraft attitude control are presented with respect to the practical consideration. This test-bed includes an air bearing 3-DOF table, Helmholtz coil, sun simulator, orbit simulator, reaction wheels and torqrods as actuators, gyro, GPS, AHRS, magnetometer, and processing board. In addition, online monitoring of attitude and orbit position in LabVIEW and Celestia software, data telemetry, battery package and power distribution board are developed in this case. Using this test-bed, different operational modes are evaluated and verified for satellite attitude control. The experimental results of damping mode show that the capability of decreasing the angular velocity to the desired value (0.3 deg/s). Moreover, a torque simulator to apply the disturbances and test of propulsion scenarios, online telemetry with ground station, near-optimal design of Helmholtz coil are the advantages of this test-bed.

The increasing need for testing and prototyping designs under more realistic conditions is responsible for the advancement of new types of simulation. In this scenario, one type of simulation which has gained high notoriety and applicability is the one known as Hardware-in-the-Loop (HIL). This technique allows real and virtual components of a system to be tested together, making it possible to perform tests under realistic (and even extreme) conditions without harming the real system or a prototype built only for testing. The objective of this work was to develop a low-cost HIL simulation platform to be used for many different applications, unlike most commercial ones, that are developed for one exclusive field of application, such as automotive, aerospace, power electronics, among others. Thus, the main contribution of this work is the project of a HIL platform capable of simulating different types of systems, making it possible to validate embedded control strategies designed for them. Two different applications are tested in order to validate the HIL platform: an active suspension and a satellite attitude control air bearing table, both controlled using a discrete Linear Quadratic Regulator (LQR) designed for each of them.

The single-gimbal control moment gyros (SGCMGs) steering law has been a standing topic in the field of spacecraft attitude control for several decades. A practical steering law should meet the following requirements simultaneously: powerful singularity avoidance ability, high precision torque output ability and fast singularity escape ability. Moreover, the angular velocity commands generated by the steering law should not jitter sharply. In this paper, a vector control-based singularity avoidance and escape steering law for SGCMGs is proposed to satisfy these goals. In addition, a system angular momentum state evaluation function for command torque is defined for the first time, which can be used to effectively evaluate the remaining angular momentum resources for the command torque. We demonstrate the effectiveness and superiority of the proposed steering law with detailed comparisons to existing efforts.

In this paper, a variable structure PID controller with a good convergence rate and robustness for satellite attitude is proposed. In order to improve the system convergence rate, the variable structure for the proportional and differential term was designed, and an angular velocity curve with a better convergence rate was achieved by this variable structure. In addition, an integral partitioning algorithm was designed, and the system robustness to disturbance torque was improved; meanwhile, the negative effect of the integral term was avoided during the converging process. The actuator failure condition was also considered, and a fault tolerant control algorithm was designed. System stability was analyzed by the Lyapunov method, and its performance was demonstrated by numerical simulation.

An optimal thruster configuration for attitude control subsystem of a spacecraft is presented in this paper. The optimal configuration is designed according to minimum number of required thrusters for satisfying desired reliability with specific redundancy level. The genetic algorithm is employed for optimization process and feasibility of the results is evaluated using algebraic and geometry methods. The main feature of the proposed configuration among feasible configuration with minimum number of required thrusters, which has held to optimal configuration, is that this configuration has maximum reliability and minimum fuel consumption. In addition to feasibility, attitude control performance of some configurations is also examined through the simulation. The results of simulation confirm that the proposed configuration has desirable performance. It is noteworthy to mention that the configuration with maximum number of required thrusters, which is a conventional configuration such that each thruster belongs to only one control channel, has less fuel consumption than optimal configuration. However, the total mass of optimal configuration is less than that of conventional configuration due to a smaller number of thrusters.

In this study, the preferred regions of Pulse-Width Pulse-Frequency Modulator (PWPFM) are obtained analytically for the static analysis. For this purpose, a comprehensive parametric study is carried out based on the two performance indices of fuel consumption and the number of thruster firings. The preferred regions are presented by normalized relations and curves. Moreover, the exact analytical solutions of the two performance indices are obtained for a class of modulators with the assumption of constant inputs. The advantages of the present study are non dimensional analysis and obtaining the preferred regions in terms of each others, resulting in more accurate regions as opposed to inequality relations using constant values for a specified input signal. In addition, in the case of specified minimum pulse width (having the update frequency and thruster time constant), determining the preferred regions becomes more limited. In this regard, useful relations and curves based on the maximum possible value for the number of the thruster firings are derived and presented.

his paper suggests arobust optimization algorithm for the design of the satellite attitude control system in order to increase the robustness of the performance under uncertainties. A single-axis on-off attitude control with rigid dynamics is considered using Schmitt-Trigger and PID controller. The model uncertainties include the moment of inertia, thrust level, thruster delay and theexternal disturbance amplitude.A weighted combination of expected value and standard deviation of pointing error is considered as an objective function for the robust optimization. The numerical solutions show that the robust optimization reduces the variations of the objective function, i.e. it increases the robustness of the system performance compared to the deterministic optimization.

It is difficult to design the control law for flexible systems with multiple lightly damped modes. Particularly, when there exist parameter perturbations, for example, resonant frequency or damping ratio etc, the stability of the system cannot be guaranteed. Conventionally, the control design of the flexible system only focus on stabilization, without considering the performance requirement of the system. For the H∞ control, even the famous H∞ loop shaping method proposed by McFarlaned, cannot further improve the system’s performance Instead, the resulting controller is unstable, and not easy to tune on in practice. In the proposed, the system bandwidth was chosen as the performance index for the H∞ optimal control. The robustness of the system was further improved by considering the performance of the system. Furthermore, the controller is stable, and the H∞ strong stabilization problem is resolved. With this design, the control accuracy and the speed of response is improved, and also the controller can be tuned on easily. The proposed work is important for the application of flexible structures in the real world． [ABSTRACT FROM AUTHOR]

A new hierarchical method for fault detection, isolation, and identification of nonlinear systems with applications for in-orbit closed-loop controlled satellites is proposed where fault detection employs unscented Kalman filter and adaptive thresholds. Fault isolation employs multiple model approach in conjunction with Bayes’ probability theorem and adaptive window, and fault identification employs dual state/parameter estimation using unscented Kalman filter. Results show up to 97% success rate under comprehensive Monte Carlo simulations for transient abrupt faults in satellite actuators. [ABSTRACT FROM AUTHOR]

An optimal thruster configuration for attitude control subsystem of a spacecraft is presented in this paper. The optimal configuration is designed according to minimum number of required thrusters for satisfying desired reliability with specific redundancy level. The genetic algorithm is employed for optimization process and feasibility of the results is evaluated using algebraic and geometry methods. The main feature of the proposed configuration among feasible configuration with minimum number of required thrusters, which has held to optimal configuration, is that this configuration has maximum reliability and minimum fuel consumption. In addition to feasibility, attitude control performance of some configurations is also examined through the simulation. The results of simulation confirm that the proposed configuration has desirable performance. It is noteworthy to mention that the configuration with maximum number of required thrusters, which is a conventional configuration such that each thruster belongs to only one control channel, has less fuel consumption than optimal configuration. However, the total mass of optimal configuration is less than that of conventional configuration due to a smaller number of thrusters. [ABSTRACT FROM AUTHOR]

This paper addresses the global stability and robust attitude tracking problem of a near polar orbit satellite subject to unknown disturbances and uncertainties. It is assumed that the satellite is fully actuated by a set of reaction wheels (RW) as control actuators because of their relative simplicity, versatility and high accuracy. The terminal sliding mode control (TSMC) approach is utilized in a two‐level architecture to achieve control objectives. In the lower layer a detumbling‐like controller is designed which guarantees the finite‐time detumbling and tracking of the desired angular velocities and based on this result a robust attitude tracking controller is designed in the upper layer to achieve 3‐axis attitude tracking in the presence of unknown disturbances and bounded uncertainties. Robust stability and tracking properties of designed controllers are proved using the Lyapunov theory. Finally, a set of numerical simulation results are provided to illustrate the effectiveness and performance of the proposed control method. [ABSTRACT FROM AUTHOR]

Achieving precise attitude control in satellites equipped with two reaction wheels remains a significant challenge. During the last decade, model predictive control has emerged as a promising technique for satellite attitude control. However, this paper takes a step by introducing the novel concept of linear time-varying fractional-order model predictive control specifically tailored for satellites with dual reaction wheels, enabling them to achieve desired orientations more effectively. This study begins by elucidating the nonlinear dynamics and kinematic equations governing the satellite's behavior. To facilitate control design, the satellite linear time-varying model is derived by linearizing the nonlinear equations around the operating point. Leveraging this linearized representation, a linear time-varying fractional-order model predictive control method is designed, leveraging the benefits of fractional-order cost functions. Finally, the simulation results show the remarkable performance of the linear time-varying fractional-order model predictive control approach, particularly when confronted with input constraints. Notably, the proposed method outperforms the traditional linear time-varying model predictive control method by achieving an 80% improvement in tracking speed, while simultaneously reducing the applied torque by 67%. Additionally, the proposed approach decreases the mean absolute error of the satellite attitude angles by over 80% compared to the conventional linear time-varying model predictive control method. This research not only addresses the pressing challenge of attitude control in satellites with dual reaction wheels but also presents a novel and highly effective approach that surpasses existing methods in terms of performance and efficiency. As such, it holds immense potential for the field of satellite attitude control, making it an indispensable read for engineers and researchers working in the domain of aerospace engineering and control systems. [ABSTRACT FROM AUTHOR]

This work presents the application and comparison of two satellite attitude control methods, using the satellite simulation platform developed by ITA, the MuSat, as a test bench. One of the algorithms under test applies feedback linearization, a method that was already validated using this test bed. The second control algorithm, developed during this research, applies control by state-dependent Riccati equation (SDRE) technique. A numerical simulation using a computational model of the MuSat, followed by an experiment using the physical platforms, is used to compare the control performance of both algorithms. The system in which the control is applied consists in a frictionless spherical air bearing that is able to rotate in three axes, allowing for validating attitude control algorithms in laboratory. The experimental results obtained indicate that the system had a better control performance under the SDRE algorithm than the algorithm based on feedback linearization. [ABSTRACT FROM AUTHOR]

Increasing in dimensions of the satellites and using light movingstructures, causes flexibility and uncertainty in their models. Therefor to control the attitude of the satellites, should use those methods which resist against the plant’s model uncertainty and could reject the disturbance and the measurementnoise. One of these methods is the robust control. But due to the location of the poles in the dynamic equations of the satellite, the design of robust controllers faces some problems. In this paper, using aninternal feedback, the dynamic equations are changed so that the poles are located in a more proper place. And then,considering flexibility affects as uncertainty and also, uncertainty in inertia matrix of the satellite, a H∞ controller, and finally to improve the performance, a µ-controller will be designed for the new equations. But these two controllers will be analyzed and compared for the primary equations and not for the new equations.For comparison, a classical controller is also designed forthe primary system.

Abstract The performance of an attitude control system is impacted by the inertia change, especially for small satellites. Although there are various control methods for satellites with inertia uncertainties, very few controllers are designed with consideration of time-varying inertia. In this paper, a linear parameter varying (LPV) model is established for a satellite with time-varying inertia. Moreover, because of their common occurrences, the actuator faults and saturation are considered during the controller design. A gain-scheduled controller is developed to guarantee the steady-state and transient performance of the system by limiting the steady-state variance and regional pole constraints. The simulations indicate that the proposed controller improves the stability performance by adjusting the gain with the inertia variations. [ABSTRACT FROM AUTHOR]

This paper proposes an integrated sensor fault detection and recovery for the satellite attitude control system. By introducing a nonlinear observer, the healthy sensor measurements are provided. Considering attitude dynamics and kinematic, a novel observer is developed to detect the fault in angular rate as well as attitude sensors individually or simultaneously. There is no limit on type and configuration of attitude sensors. By designing a state feedback based control signal and Lyapunov stability criterion, the uniformly ultimately boundedness of tracking errors in the presence of sensor faults is guaranteed. Finally, simulation results are presented to illustrate the performance of the integrated scheme. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

Satellite mission life are primarily dictated by the state of health of its Reaction Wheel Assembly (RWA), especially for commercial GEO satellites. Propulsion based actuating devices can't replace or accommodate for the loss of RWA because thrusters based pointing control is not as accurate as the RWA. The RWA is the primary set of actuators (as compared to thrusters for orbit maintenance and maneuvering) mainly responsible for the satellite mission for accurately and precisely pointing its payloads to the right targets to conduct its mission operations. The RWA, consisting of either a set of four individual wheels mounted in a pyramid or three in an orthogonal configuration, is mainly the primary control actuators of the satellite during its normal operations. Future GEO satellites will be required to achieve much longer lives than their typical 15 years normal life expectancy. Driven by customers' demands/goals and competitive market have challenged Attitude Control Subsystems (ACS) engineers to develop better ACS algorithms to address such a stringent requirement. There are two main directions to design satellite's under-actuated control subsystem: (1) Attitude Feedback with Zero Momentum Principle and (2) Attitude Control by Angular Velocity Tracking via Small Time Local Controllability concept. Successful applications of these control laws have been largely demonstrated via simulation for the rest to rest case. Limited accuracy and oscillatory behaviors are observed in three axes for non-zeros wheel momentum while realistic loss of a wheel scenario (i.e., fully actuated to under-actuated) has not been closely examined! This study revisits the under-actuated control design with detailed set ups of multiple scenarios reflecting real life operating conditions which have put current under-actuated control laws mentioned earlier into a re-evaluation mode since rest to rest case is not adequate to truly represent an on orbit failure of a single wheel. The study is intended to facilitate the ACS community to further develop a more practical underactuated control law and present a path to extend these current thinking to address a more realistic reconfigurable ACS subject to a dynamic transition from a 3 RWs mode to 2 RWs mode. [ABSTRACT FROM AUTHOR]

Because they are inexpensive platforms for satellites, CubeSats have become a low-cost way for universities and even developing countries to have access to space technology. This paper presents the ITASAT design, particularly the Attitude Determination and Control Subsystem, the Onboard Software, and the Assembly,Integration and Testing program. The ITASAT is a 6U CubeSat nano-satellite in development at the Instituto Tecnológico de Aeronáutica, in São José dos Campos, Brazil. The platform and its subsystems will be provided by industry while the payloads are being designed and developed by the principal investigators. The ITASAT Attitude Determination and Control Subsystem will rely on a 3-axis magnetometer, 6 analog cosine sun sensors, 3-axis MEMS gyroscopes, 3 magnetic torque coils, and 3 reaction wheels. The Attitude Determination and Control Subsystem operating modes, control laws, and embedded software are under the responsibility of the Instituto Tecnológico de Aeronáutica. A Kalman filter shall be employed to estimate the quaternion attitude and gyroscope biases from sensor measurements. The Attitude Determination and Control Subsystem operating modes are the nominal mode, with geocentric pointing attitude control and the stabilization mode, in which only the satellite angular velocity is controlled. The nominal mode will be split into 2 sub-modes: reaction wheel control plus magnetic wheel desaturation and 3-axis magnetic attitude control. Simulation results have shown that the attitude can be controlled with 1-degree accuracy in nominal mode with the reaction wheels, but these errors grow as much as 20 degrees or higher with the 3-axis magnetic control.

Recently, new actuators known as fluidic momentum controllers (FMC) have been proposed for satellite attitude control. This actuator has many advantages such as high applied torque to weight ratio, easiness in assembly, low transmitted vibration and so on respect to the other momentum exchange devices like momentum/reaction wheels or CMGs. However, one of the main problems in implementing such actuators is complexity in mathematic modeling of them. This issue makes many researchers to use a simplified model for control system designing without to consider uncertainty in this simplified model. In this paper for eliminating of this problem, an adaptive sliding mode control (SMC) has been used in this article. This control method is not also robust respect to uncertainties but also can estimate over threshold of them without necessity to use larger and heavier actuators to be sure of satellite stability. According to the results, we can observe that proposed control system is capable to reach satellite to the desire attitude in minimum time and without overshoot.

In this paper attitude control system of nanosatellite based on Single Gimbal Control Moment Gyroscope (SGCMG) is presented. A LQR/LQG method is developed for stability of satellite and a feedback quaternion strategy is used for maneuvering mode. In the stabilization mode LQR/LQG controllers are designed with linearization of nonlinear dynamic equation of satellite and control moment gyroscope, so that in other reseach didn’t use this controller in the stabilization mode of this system. In the maneuvering mode a feedback quaternion controller applyed for nonlinear system. Numerical simulations are provided to show the efficiency of the proposed controller for a nanosatellite with four single gimbal control moment gyroscope pyramid cluster. Results of simulations shown that LQR/LQG method is more accurate in compared with feedback quaternion controller.

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.

This article presents software for a satellite position and attitude control system based on mathematical modeling. The software is developed using Matlab for conceptual design; an added benefit is that it shortens design time and decrease design costs. Providing interactive modules for different actuators with various configurations and control algorithms makes this toolbox suitable for analyzing the effect of design parameters on satellite response and stability. Moreover, this toolbox contains adjustable modules for date and orbital parameters, which helps users improve their realization of satellite position effects on different matters such as eclipse, magnetic field, and satellite communication with ground station. Taking position data, the software computes disturbance torques and gives users the ability to analyze satellite attitude control performance. The software’s ability to show simulation results as a set of graphics and text windows makes it more user friendly. [ABSTRACT FROM PUBLISHER]

An active fault-tolerant satellite attitude control scheme based on fault effect classification is presented at the occurrence of faults associated with torques. In this paper, the flexibility and practicability of the fault-tolerant scheme are top priorities. Faults are modeled and divided into additive and multiplicative ones in order to estimate and deal with them specifically and exactly. The additive faults, including additive part of flywheel faults and other uncertain fault torques, are estimated by additive fault estimator and compensated on the basis of nominal controller, whereas the multiplicative faults, denoting torque gain parameter faults of flywheels, are estimated by multiplicative fault estimator and the estimated fault parameters are used for dynamic torque command distribution of flywheels. The final simulation examples and performance comparison of three fault-tolerant schemes show that the proposed scheme based on fault effect classification is an effective, flexible and saving-energy fault-tolerant satellite attitude control scheme. It possesses an engineering value for improving reliability and prolonging on-orbit working lifetime of satellites. [ABSTRACT FROM AUTHOR]

Because they are inexpensive platforms for satellites, CubeSats have become a low-cost way for universities and even developing countries to have access to space technology. This paper presents the ITASAT design, particularly the Attitude Determination and Control Subsystem, the Onboard Software, and the Assembly, Integration and Testing program. The ITASAT is a 6U CubeSat nano-satellite in development at the Instituto Tecnológico de Aeronáutica, in São José dos Campos, Brazil. The platform and its subsystems will be provided by industry while the payloads are being designed and developed by the principal investigators. The ITASAT Attitude Determination and Control Subsystem will rely on a 3-axis magnetometer, 6 analog cosine sun sensors, 3-axis MEMS gyroscopes, 3 magnetic torque coils, and 3 reaction wheels. The Attitude Determination and Control Subsystem operating modes, control laws, and embedded software are under the responsibility of the Instituto Tecnológico de Aeronáutica. A Kalman filter shall be employed to estimate the quaternion attitude and gyroscope biases from sensor measurements. The Attitude Determination and Control Subsystem operating modes are the nominal mode, with geocentric pointing attitude control and the stabilization mode, in which only the satellite angular velocity is controlled. The nominal mode will be split into 2 sub-modes: reaction wheel control plus magnetic wheel desaturation and 3-axis magnetic attitude control. Simulation results have shown that the attitude can be controlled with 1-degree accuracy in nominal mode with the reaction wheels, but these errors grow as much as 20 degrees or higher with the 3-axis magnetic control. [ABSTRACT FROM AUTHOR]

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This article has discussed the development of a three-axis attitude digital controller for an artificial satellite using a digital signal processor. The main motivation of this study is the attitude control system of the satellite Multi-Mission Platform, developed by the Brazilian National Institute for Space Research for application in different sort of missions. The controller design was based on the theory of the Linear Quadratic Gaussian Regulator, synthesized from the linearized model of the motion of the satellite, i.e., the kinematics and dynamics of attitude. The attitude actuators considered in this study are pairs of cold gas jets powered by a pulse width/pulse frequency modulator. In the first stage of the project development, a system controller for continuous time was studied with the aim of testing the adequacy of the adopted control. The next steps had included an analysis of discretization techniques, the setting time of sampling rate, and the testing of the digital version of the Linear Quadratic Gaussian Regulator controller in the MATLAB/SIMULINK. To fulfill the study, the controller was implemented in a digital signal processor, specifically the Blackfin BF537 from Analog Devices, along with the pulse width/pulse frequency modulator. The validation tests used a scheme of co-simulation, where the model of the satellite was simulated in MATLAB/SIMULINK, while the controller and modulator were processed in the digital signal processor with a tool called Processor-In-the-Loop, which acted as a data communication link between both environments.function and required time to achieve a given mission accuracy are determined, and results are provided as illustration.

This Paper presents a dynamic model of a micro-satellite in Mesbah class. At this model aerodynamic torque and solar radiation pressure torques are considered as disturbance torques. Gravity gradient torque is assumed as stabilizing torque and acts as a passive controller. Magnetic torquers act as an active controller. There are three methods of optimization of power consumption; first using LQR controller, secondly using the mapping function (which is suggested to ensure that the generated magnetic moment by the coils is perpendicular to the local magnetic field vector), and finally powering on control system over the earth stations only for the purpose of power saving.