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The purpose of this paper is to resolve complex nonlinear dynamical problems of the pitching axis of solar sail in body coordinate system compared with inertial coordinate system. And saturation condition of controlled torque of vane in the orbit with big eccentricity ration, uncertainty and external disturbance under complex space background are considered.

The pitch dynamics of the sailcraft in the prescribed elliptic earth orbits is established considering the torques by the control vanes, gravity gradient and offset between the center-of-mass (cm) and center-of-pressure (cp). The maximal torques afforded by the control vanes are numerically determined for the sailcraft at any position with any pitch angle, which will be used as the restriction of the attitude control torques. The finite/infinite time adaptive sliding mode saturation controller and Bang–Bang–Radial Basis Function (RBF) controller are designed for the sailcraft with restricted attitude control torques. The model uncertainty and the input error (the error between real input and ideal control law input) are solved using the RBF network.

The finite true anomaly adaptive sliding mode saturation controller performed better than the other two controllers by comparing the numerical results in the paper. The control torque saturation, the model uncertainty and the external disturbance were also effectively solved using the infinite and finite time adaptive sliding mode saturation controllers by analyzing the numerical simulations. The stabilization of the pitch motion was accomplished within half orbit period.

The complex accurate dynamics can be approximated using the RBF network. The controllers can be applied to stabilization of spacecraft attitude dynamics with uncertainties in complex space environment.

Advanced control method is used in this paper; saturation of controlled torque of vane is resolved when the orbit with big eccentricity ration is considered and uncertainty and external disturbance under complex space background are settled. Moreover, complex and accurate nonlinear dynamical model of pitching axis of solar sail in body coordinate system compared with inertial coordinate system is provided.

This paper aims to describe a novel type of attitude control system (ACS) in different configurations. This servomechanism is compared with control moment gyro (CMG) in significant parameters of performance for ACS of rigid satellite.

This new actuator is the fluid containing one or more rings and fluid flow is supplied by pump. The required torque control is obtained by managing fluid angular velocity. The cube-shaped satellite with three rings of fluid in the principle axes is considered for modeling. The satellite is considered rigid and nonlinear dynamics equation is used for it. In addition, the failure of the pyramid-shaped satellite with an additional ring fluid is discussed.

The controller model for four fluid rings has more complexity than for three fluid rings. The simulation results illustrated that four fluid rings need less energy for stabilization than three fluid rings. The performance of this type of actuator is compared with CMG. At last, it is demonstrated that performance parameters are improved with fluid ring actuator.

Fluid ring actuator can be affected by environmental pressure and temperature. Therefore, freezing and boiling temperature of the fluid should be considered in system designation.

Fluid ring servomechanism can be used as ACS in rigid satellites. This actuator is compared by CMG, the prevalent actuator. It has less displacement attitude maneuver.

The results provide the feasibility and advantages of using fluid rings as satellite ACS. The quaternion error controller is used for this model to enhance its performance.

Attitude determination and control subsystem (ADCS) is a vital part of earth observation satellites (EO-Satellites) that governs the satellite’s rotational motion and pointing. In designing such a complicated sub-system, many parameters including mission, system and performance requirements (PRs), as well as system design parameters (DPs), should be considered. Design cycles which prolong the time-duration and consequently increase the cost of the design process are due to the dependence of these parameters to each other. This paper aims to describe a rapid-sizing method based on the design for performance strategy, which could minimize the design cycles imposed by conventional methods.

The proposed technique is an adaptation from that used in the aircraft industries for aircraft design and provides a ball-park figure with little engineering man-hours. The authors have shown how such a design technique could be generalized to cover the EO-satellites platform ADCS. The authors divided the system requirements into five categories, including maneuverability, agility, accuracy, stability and durability. These requirements have been formulated as functions of spatial resolution that is the highest level of EO-missions PRs. To size, the ADCS main components, parametric characteristics of the matching diagram were determined by means of the design drivers.

Integrating the design boundaries based on the PRs in critical phases of the mission allowed selecting the best point in the design space as the baseline design with only two iterations. The ADCS of an operational agile EO-satellite is sized using the proposed method. The results show that the proposed method can significantly reduce the complexity and time duration of the performance sizing process of ADCS in EO-satellites with an acceptable level of accuracy.

Rapid performance sizing of EO-satellites ADCS using matching diagram technique and consequently, a drastic reduction in design time via minimization of design cycles makes this study novel and represents a valuable contribution in this field.

The quad‐rotor is an under‐actuation, strong coupled nonlinear system with parameters uncertainty, unmodeled disturbance and drive capability boundedness. The purpose of the paper is to design a flight control system to regulate the aircraft track the desired trajectory and keep the attitude angles stable on account of these issues.

Considering the dynamics of a quad‐rotor, the closed‐loop flight control system is divided into two nested loops: the translational outer‐loop and the attitude inner‐loop. In the outer‐loop, the translational controller, which exports the desired attitude angles to the inner‐loop, is designed based on bounded control technique. In consideration of the influence of uncertain rotational inertia and external disturbance, the backstepping sliding mode approach with adaptive gains is used in the inner‐loop. The switching control strategy based on the sign functions of sliding surface is introduced into the design procedure with respect to the input saturation.

The validity of the proposed flight control system was verified through numerical simulation and prototype flight experiment in this paper. Furthermore, with relation to the flying, the motor speed is kept in the predetermined scope.

This article introduces a new flight control system designed for a quad‐rotor.

The purpose of this paper is to describe the general idea, design, and implementation of control system for general aviation aircraft which reduces pilot workload.

Proposed indirect flight control system framework is intended to simplify piloting, reduce pilot workload, and allow low‐end general aviation aircraft to operate under deteriorated meteorological conditions. Classical control theory is used for the design of the flight control laws. Although not inherently robust, controllers with classical control logic are made sufficiently stable using a correct and updated controller structure.

Despite controversies between perception of a modern manned aerial vehicle and limitations imposed by legacy airworthiness codes it is shown that a pilot workload reducing system can be successfully implemented onboard of a low‐end general aviation aircraft.

Hi‐level control laws and optimization of handling qualities can lead to unfavourable and unpredictable forms of man‐machine interactions, e.g. pilot‐induced oscillations.

General aviation aircraft are mostly flown by a single pilot, who could benefit from an intelligent system or “virtual copilot” assisting in or supervising the aircraft's safe operation under any conditions. Aircraft with this capability represents a next step in the evolution that might ultimately lead to trajectory‐based free‐flight concept of aircraft operations.

The paper introduces a safety enhanced digital flight control system on board small general aviation aircraft.

To provide a progress report into research conducted to establish guidelines for the development of guidance vision aids.

The first stage of the research is to establish a coherent engineering basis for the methods of (visual) motion perception and control to inform the design of pilot aids that will support flight in degraded visual conditions, particularly when close to the ground. The next stage will then be to construct and evaluate synthetic displays that recover the visual cues necessary to allow flight in degraded visual conditions for a range of manoeuvres using the flight simulation facilities at the University of Liverpool (UoL). The research is guided by tau (time to contact) theory from the field of ecological psychology.

The closure of spatial gaps for a number of aircraft manoeuvres are presented in the tau domain. Analysis of the landing flare manoeuvre suggest that both a constant rate of change of tau strategy and an intrinsic tau‐guidance strategy will yield benefits in terms of touchdown descent rate if presented as display symbology.

Results are presented from trials where only one professional pilot was used. Results from a wider population of pilots need to be analysed to ensure that the observed trends are generic.

The reported results are being used in the next phase of the research project to inform the design of a guidance vision‐aid for the flare manoeuvre. These displays will be tested in flight simulation trials.

The research takes a theory of motion perception and applies it to aircraft guidance display technology.

The purpose of this paper is to present analysis and primary evaluation of different control laws implemented on experimental indirect (fly‐by‐wire) flight control system designed for perspective general aviation aircraft.

The control law tests have been accomplished on the flight simulation stand equipped with side‐stick, throttle lever and flight instrument display. Every evaluator was caring out 2‐4 five min instrument flights (IR) according to command shown on the screen. PZL‐110 general aviation aircraft properties and seven modes of control system operation were modeled and examined.

Results of evaluation by 45 commercial pilots are analyzed and handling qualities of the small aircraft equipped with the indirect flight control system (fly‐by‐wire) have been examined. In this way, the most convenient control law was chosen for design the user‐friendly, human‐centered, simplified software‐based flight control system.

The result of research can be implemented on real indirect flight control system dedicated to general aviation aircraft.

This paper presents the practical approach for analysis of handling qualities of general aviation aircraft equipped with indirect flight control system. This kind of works concern to military and transport airplanes are known, however there are no published work in the area of small aircraft so far.

This study aims to establish the laser links between satellites among large-scale distributed satellite systems; a combined attitude control strategy containing two stages is proposed in this paper.

These two stages are: one is the attitude initial pointing control to change the attitude of satellite pointing to the other satellite based on the position information of each satellite; the other one is the high precision attitude tracking control to scan the uncertainty cone because the initial pointing control accuracy is not enough to establish the laser link. At the initial pointing control stage, a method to determine the target attitude of each satellite is presented based on the position information of each satellite, and the fuzzy adaptive control algorithm is used to control the satellites to its calculated attitude. Then, at the high precision attitude tracking control stage, a strategy for laser link acquisition and scanning the uncertainty cone by the lasers of the spacecraft is proposed, and an angular velocity tracking scanning controller is designed while the convergence of the attitude tracking error is ensured through Lyapunov–Krasovskii theory.

Simulations are conducted to verify the effectiveness of the proposed control algorithm, and the laser link for a large-scale distributed satellite system with super long distance is achieved through a combined attitude control strategy.

A combined attitude control strategy is valid for a large-scale distributed satellite system with super long distance.

A combined attitude control strategy can be used to achieve laser link acquisition for a large-scale distributed satellite system like space gravitational wave detection.

A combined attitude control strategy can provide a way to solve the typical problem that pointing control accuracy is not enough to establish the laser link for a large-scale distributed satellite system.

The purpose of the paper is to design a new attitude stabilization system for a microsatellite based on single gimbal control moment gyro (SGCMG) in which the gimbal rates are selected as controller parameters.

In the stability mode, linear quadratic regulator (LQR) and linear quadratic Gaussian (LQG) control strategies are presented with the gimbal rates as a controller parameters. Instead of developing a control torque to solve the attitude problem, the attitude controller is developed in terms of the control moment gyroscope gimbal angular velocities. Attitude control torques are generated by means of a four SGCMG pyramid cluster.

Numerical simulation results are provided to show the efficiency of the proposed controllers. Simulation results show that this method could stabilize satellite from initial condition with large angles and with more accuracy in comparison with feedback quaternion and proportional-integral-derivative controllers. These results show the effect of filtering the noisy signal in the LQG controller. LQG in comparison to LQR is more realistic.

The LQR method is more appropriate for the systems that have project models reasonably exact and ideal sensors/actuators. LQG is more realistic, and it can be used when not all of the states are available or when the system presents noises. LQR/LQG controller can be used in the stabilization mode of satellite attitude control.

The originality of this paper is designing a new attitude stabilization system for an agile microsatellite using LQR and LQG controllers.

The purpose of this paper is agile attitude control design with the novel three-dimensional (3D) magnetically suspended wheel (MSW) that is the preferred type for agile maneuvering compared to conventional control moment gyro due to frictionless, low vibration and long lifetime. This system does not require a separate steering law for pyramid arrangement to derive tilt angles. It is also conducting an agile maneuver with high accuracy despite the high-frequency disturbances.

In this paper, a disturbance observer-based attitude stabilization method is proposed for an agile satellite with a pyramid cluster of the novel 3D magnetically suspended wheel actuator. This strategy includes a disturbance observer and a linear quadratic regulator controller. The rotor shaft deflection of MSW is actively controlled to reduce vibration and producing gyro torque. The deflection angle of the pyramid cluster MSWs considered as control parameters. The closed-loop stability is proved by using the Lyapunov strategy. The efficiency and performance of the offered method verified by numerical simulation via MATLAB/SIMULINK software.

According to simulation results, the disturbance observer-based control controller stabilized the system with high accuracy and optimal tilt angles without any extra steering law equation. Hence, the system speed is increased, and the system error is minimized without separate steering law.

The magnetically suspended wheel is a new kind of inertia actuator for attitude control that has several benefits such as frictionless, high-speed rotor, clean environment and low vibration compared to the traditional wheel. It has complex nonlinear dynamics that cause have complicated controller design. The proposed strategy stabilizes the system and conducting an agile maneuver with high precision despite the high-frequency disturbances. It is applicable for some missions requiring high accuracies, like Earth observation and the solar observation mission that require a very accurate pointing control and a long lifetime.

This paper is the initial paper to design a pyramid array for magnetically suspended wheels. Compared to other research, this method doesn’t need a separate steering law of the MSWs cluster and presented optimal tilt angles with less computational. Also, it designs a disturbance observer-based controller for this system that proposed high accuracy and agile stabilization.