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In this paper, we try to use modeling based on singular perturbation theory, in order to control satellite attitude during the wide rolling angle maneuvering through nonlinear H∞ control strategy. Differential equations describing dynamics of the satellite are presented first, and by choosing the appropriate dynamic model for actuators and based on the standard singular perturbation model, the closed-loop system is created. Next, this model is put into the appropriate form to solve H∞ problem. Then, after solving the HJI equation, the control law is determined. Simulation results for a nominal satellite control based on our approach are finally presented.

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This paper presents a new method to design optimal thrusters’ configuration for geostationary satellite in order to reduce the fuel consumption and increase the control accuracy. The thrusters configuration generally contains information about thrusters fixed on the satellite body structure, including their location, orientation. One important factor playing a key role in thrusters’ configuration design is satellite force-torque analysis. The proposed configuration, however, should lead to fulfill specified attitude maneuver when the set of force and torque produced by satellite thruster system is adequate. For this purpose, two optimization methods using genetic algorithm (GA) and differential evolution (DE) has been applied to determine the optimal thrusters configuration on the communication satellite body. The cost function employed to minimize both the fuel consumption and error generated by thrusters installation and uncertainties. Moreover, this work allows applying some different constraints in the proposed formulation including minimization of the thruster plume impingement effect on the satellite outer structure and on the solar arrays and the second one is the satellite dimension and geometry. Simulation results show that DE outperforms GA in terms of accuracy and CPU time. Effectiveness of differential evolution algorithm is illustrated in the paper when compared with GA results.

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In this paper, LQG/LTR controller is designed for attitude control of the geostationary satellite at nominal mode. Usage actuator in this paper is the reaction wheel and control torque is determined by the LQR regulator. Usage sensors in this article are sun and earth sensors and EKF are used for estimation of noisy states. LQR controller signal has good performance, if all system's states are considered in system output feedback. But this method is ideal and does not include model noise and sensors noise. Therefore, LQG and LQG/LTR controllers are designed based on the estimated states, and are compared with LQR controller. Controllers gain coefficients are obtained based on linearization about working point. It caused to robustness and similarity of LQG and LQG/LTR response. The results show that control overshoot of LQR is greater than the others.

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In this paper, a LQG/LTR controller is proposed for attitude control a geostationary satellite at nominal phase. Basically, proposed methodology includes three parts: LQR regulator, EKF, and loop transfer recovery. Controller design is based on the linearized equations of the spacecraft dynamics using reduced quaternion model. Reduced quaternion model solve uncontrollable problem in some subspaces in the linearized state space quaternion model using all four components of quaternion. Spacecraft actuators are reaction wheels and attitude determination sensors are sun and earth sensors. LQR controller is ideal and it doesn’t account for the model uncertainty and sensor noise and it uses the feedback of the full states. To consider the model uncertainty and sensor noise, we have designed EKF which is used by LQG and LQG/LTR controllers. Controller gain coefficients are obtained using a reduced quaternion model, and based on linearization around the equilibrium point and the natural frequency of the closed loop system. To increase the robustness of the design with respect to solar radiation disturbance, singular values of LQG are approximated to Kalman filter, in LTR section. The results demonstrate that LQG/LTR performance is better than LQG’s and LQG/LTR has a good robust stability margin with respect to disturbances.

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This paper is concerned with design, develop and implementation of a quaternion based attitude control system for a rigid suborbital module which using cold gas thrusters over a short-duration mission. The quaternion controller produces a demand torque, and a pulse-width pulse-frequency (PWPF) modulator determines the necessary thruster fire signals. The effect of disturbances on module attitude has been investigated and the most significant found to be due to misalignment of thrusters effects. The system concept has been evaluated through modeling in Simulink and a rapid prototype hardware-in-the-loop platform and has been found to meet the requirements laid out for a typical module mission. The satisfactory performance of the controllers was illustrated through both numerical and hardware-in-the-loop simulations, where a system of twelve thrusters and load sensors were implemented in the hardware and disturbance effects such as thrust misalignment and sensor noise were studied. The results show the effectiveness of the proposed control method for agile attitude maneuvers of suborbital modules. The results of the HIL simulation were also used for tuning the parameters of the module’s numerical simulation that is to be used for error budgeting analyses.

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High-precision three-axis attitude control scheme is vitally important to deal with the overactuated spacecraft, as long as the overall performance through rapid response can be in general acquired. Due to the fact that the rigid-flexible spacecraft is somehow applicable, in so many academic and real environments, there is a consensus among experts of this field that the new insights in developing the present complicated systems modeling and control are highly recommended with respect to state-of-the-art. The new hybrid control scheme presented here is organized in line with the linear approach, which includes the proportional derivative based quadratic regulator and the nonlinear approach, which includes finite-time sliding mode control, as well. It should be noted that the three-axis angular rates of spacecraft under control are all dealt with in inner closed loop control and the corresponding rotation angles are also dealt with in outer closed loop control, synchronously. 

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In this paper, a method of tri-axial gravity gradient stabilization of satellite in circular orbit is proposed and investigated. In this method, only one actuator is employed. A satellite with varying-length boom is considered consisting of two rigid bodies having the freedom of moving in the boom direction. The only control input is the force between these two bodies to control the varying-length boom. The gravity gradient torque is considered as the only external torque acting on the satellite. The system is under-actuated and has Hamiltonian structure. So, the port-Hamiltonian approach is utilized. The equations of motion of the system are obtained in Hamiltonian formulation. The equilibrium points and their required control inputs are determined. The linearization around the equilibria is carried out and it can be seen that the linear dynamics of pitch-boom and roll-yaw are decoupled. Therefore, the roll-yaw dynamics is linearly uncontrollable. The method of energy shaping and damping injection is used for controller design. The conditions on the energy shaping control law to stabilize the system are determined. Further, the resulting closed-loop system is analyzed. The closed-loop system has center manifolds. Finally, the performance of the closed-loop system, convergence of state trajectory to the center manifold and its non-exponential convergence is shown by simulation.

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In the present study, a new attitude stabilization concept has been investigated for a satellite considering failure in one or more reaction wheels. In this approach control torques could be generated using only one thruster mounted on a two axis gimbal mechanism. In the other word, in the absence of reaction wheel(s), control torques are generated by applying a thruster rotating mechanism which can be turned around two axes by thruster vector. If any failure happened in reaction wheels, gimbal angles mechanisms will be added to the system as input controlling. Controller algorithm based on dynamic and kinematic equations of the satellite’s motion, has been developed in the presence of disturbances. Three-axis stabilization of the attitude in a LEO orbit satellites under disturbances has been executed by applying three reaction wheel actuators to produce torque in each direction. Disturbance torques that are commonly applied to the satellites are gravity gradient, solar radiation pressure and aerodynamics. For training the intelligent neuro-fuzzy controller, PID controller is employed. Numerical simulations show that, the recommend controlled method have acceptable results (in the presence of disturbances) and adding of a thruster actuator to the system as a redundancy, could enhance the space missions reliability and if any fault happened in the operation of reaction wheels, thruster mechanisms come in to control system , accurately, and sustained satellite stability at desirability attitude.

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An Aerial Robot or Unmanned Aerial Vehicle (UAV) is an aerial vehicle that provides its flight condition using aerodynamic forces. Also, this vehicle can be named as an autonomous robot. This robot is an under-actuated system and it is inherently unstable. Thus, the control of this nonlinear system is a problem for both practical and theoretical interest. So, the goal of this research is to contrast with highly nonlinear dynamic system of Octorotor that its control is difficult in many cases and it causes existence of instability in this Unmanned Aerial Vehicle (UAV). At the first, the structure of Octorotor is studied in this paper in order to increasing power, more carrying and increment of resistance into changing and distribution. Also, the electronic and mechanic of this robot is studied in some sections. Then, in the following, in order to attitude control of robot with introduction of dynamic system, one of the most common implemented controllers is applied on this robot. Initially, this process is done on the dynamic model of robot by Matlab/Simulink software and finally, implementation of this controller is applied on a fabricated Octorotor during a real flight in autonomous trajectory tracking in outdoor environment. At last, the study of sensors results is also shown.

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In this paper, the performance of a single-axis attitude control with pulse-width pulse-frequency (PWPF) modulation is enhanced using a modified proportional-integral-derivative (PID) controller for a rigid satellite with on-off thruster actuators. For this purpose, the well-known observer-based PID approach is utilized. The on-off thruster actuator is modeled with a constant delay followed by a second-order binomial transfer function. The modulator update frequency is limited to 40 Hz as an input to the on-off thruster actuators. In this study, the design criteria of pointing accuracy, overshoot of the attitude response, fuel consumption, and the number of thruster firings are considered for a step external disturbance (with different values). The parameters of the observer-based PID controller are tuned using parametric search method. Simulation results show that the fuel consumption and settling time of the observer-based approach are considerably decreased with respect to those of PID controller with PWPF modulator. Moreover, the overshoot of the observer-based approach is omitted. Finally, the robustness of the observer-based modified PID controller is investigated in presence of uncertainties in satellite moment of inertia and thrust level of on-off actuators.

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This paper presents a completely practical control approach for quadrotor drone. Quadrotor is modelled using Euler-Newton equations. For stabilization and control of quadrotor a classic PID controller has been designed and implemented on the plant and a fuzzy controller is used to adjust the controller parameters. Considering that quadrotor is a nonlinear system, using classic controllers for the plant is not effective enough. Therefor using fuzzy system which is a nonlinear controller is effective for the nonlinear plant. According to the desire set point, fuzzy system adjusts the controller gain values to improve the performance of quadrotor and it leads to better results than classical PID controller. To study the performance of fuzzy PID controller on attitude control of the system, a quadrotor is installed to the designed stand. The system consists of accelerometer and gyroscope sensors and a microcontroller which is used to design fuzzy PID attitude controller for the quadrotor. Considering that the experimental data has lots of errors and noises, Kalman filter is used to reduce the noises. Finally using the Kalman filter leads to better estimation of the quadrotor angle position and the fuzzy PID controller performs the desired motions successfully.

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In this paper, the preferred regions of pulse-width pulse-frequency (PWPF) modulator parameters are obtained based on zero-input, static, and dynamic analysis in the presence of sensor noise as an input noise to PWPF modulator. The design parameters are reduced to 3 by using the quasi-normalized equations of PWPF modulator. Therefore, the results are applicable for grouped parameters, regardless of the value of each parameter, separately. Moreover, the computational burden is highly decreased, especially in a statistical analysis. The input noise of the modulator is constructed by a low pass filter driven by a white Gaussian noise. The fuel consumption and number of thruster firings are considered as performance indices. The modulator output frequency is also limited to 50 Hz. The preferred regions of quasi-normalized system are extracted based on eliminating the upper 30% (and 10%) of the plotted graphs for the above-mentioned performance indices. Finally, the preferred regions can simply be viewed in our resulting curves, i.e., normalized hysteresis plotted versus normalized PWPF on-threshold for different values of modulator time constant. Each of these curves is plotted for a specified value of input noise power spectral density.

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Attitude control of the UAV’s is basis of the of many control systems such as position control, trajectory traking, traking moving targets and obstacle avoidance. Hence, one of the most important parts of the UAV's control is designing an appropriate and efficient controller, so that system being able to eliminates or reduces external disturbances, mechanical underactuation, changes in the model or physical parameter and interactions between its subsystems. In this paper, the attitude control problem is studed. For this purpose, the dynamics model of a quadrotor is derived by using Newton-Euler mtethod and the required parameters of the model such as moment of inertia, thrust and drag torque coefficient identified by experimental methods and an actual physical sample. Then, modidied PID and sliding mode controllers are designed to provide attitude traking for quadrotor and performance of these controllers is investigated in the presence of disturbance and sensors noise. Finally, the desgned cotrollers are implemented on a real 3DOF system and the experimental results are compared with the simulation results.

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In this paper, a linear dynamic model for a reaction wheel is identified using experimental analysis. To do this, online input-output data of reaction wheel is sent and received by CAN protocol working with the frequency of one mega bit per second. The experimental hardware consists of reaction wheel, processing board, CAN protocol, and LabVIEW monitoring. Modeling assumes the reaction wheel and its inner control circuit as a black box and takes into account the practical considerations. Initially, behavior of the reaction wheel is examined using test signals for velocity and acceleration as inputs. After that, the test signals are replaced by Chirp and PRBS signals and the output results are saved. According the results obtained in the tests, ARMAX and ARX linear dynamic models are assigned to the motor and different orders of these models are compared with each other to reach the appropriate order of the models. Furthermore, a delay is also incorporated in the model and its proper order is determined by the simulations. Finally, to validate the proposed model, the outputs of the model and plant are compared followed by exerting a new test signal. The results indicate a good agreement between the proposed model and the practical behavior.

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Maneuvering with the highest speed and low power has always been a challenge to design a satellite and spacecraft control system. In this paper, apart from the complexity of modeling actuators, different control methods were used to control the satellite attitude in the presence of uncertainties and disturbances in satellites, in order to obtain an explicit response to minimize the EULERINT criterion. The EULERINT criterion is the integral of the Euler angles between the body axes and the target around Euler's axis over time and somehow interprets the speed of the satellite maneuver in the three control axes. First, using the proportional-derivative control, the comparison of the EULERINT criterion in the application of different kinematic representations (Euler, quaternion vectors and direction cosine matrix equations) in linear and nonlinear models of the satellite was carried out. Then the comparison of the EULERINT criterion between the different methods was presented using the quaternion kinematic, which has the least amount of EULERINT, through changing the proportional-derivative controller to linear-quadratic regulator controllers, pole placement, adaptive, fuzzy, and adaptive-fuzzy. The comparison was conducted to achieve the best control method in terms of frequency response, the lowest EULERINT and the least control effort to control the attitude of the satellite in the presence of disturbance and uncertainty.