Attitude Control Torque Research Articles

Research on Sun-Oriented Spin-Stabilized Attitude Control of Micro/Nano Satellite Using Only Magnetic Control

Sun-oriented attitude control is one of the most important attitude control modes for most micro/nano satellites, which directly affects the on-orbit energy acquisition. Therefore, it is of great importance to use the simplest sensors and actuators and the most reliable algorithm to achieve sun-oriented attitude control. A sun-oriented spin-stabilized attitude control method for micro/nano satellites using only magnetic control is proposed in this paper, in which the control progress is divided into four phases: initial damping phase, sun-aligned phase, spin-up phase, and spin-stabilized phase. The influence of the shadow zone of the orbit, offset installation of sun sensor and solar panel, limitation of the field of view of the sun sensor, and environmental disturbance torques are all considered in the proposed method. The control performance is evaluated by numerical simulations, and the simulation results show that the proposed method is applicable to the satellite equipped with a sun sensor and a 3-axis magnetometer as attitude sensors and three magnetic torquers installed orthogonally as attitude actuators. The proposed method is applicable to most Earth-orbit satellites for which the geomagnetic field can provide sufficient attitude control torque.

Integrated Power, Attitude, and Vibration Control of Gyroelastic Body

Gyroelastic body refers to a flexible structure with distributed angular momentum exchange devices, such as variable speeds control moment gyros (VSCMGs). A VSCMG contains a rotor with high rotating speeds. The direction and the magnitude of the rotor’s angular momentum can be simultaneously tuned.When distributing a set of VSCMGs on a flexible structure, they could simultaneously produce control torques for attitude control and modal forces for vibration suppression. Meanwhile, the high spinning rotor in the VSCMGs can be used for energy storage. Thus, an integrated power, attitude, and vibration control system (IPAVCS) for a free flexible structure can be obtained. In this work, the control scheme for the IPAVCS is developed. A nonlinear controller is first designed to calculate the desired attitude control torques and vibration suppression modal forces for attitude stabilization and vibration suppression. Then, a robust pseudoinverse gimbal steering law and a pseudoinverse rotor acceleration law are developed to generate the desired control input and meet the power requirement, respectively. Numerical examples are conducted to verify the effectiveness of the proposed control scheme.

Variable Mass Control and Parameter Identification of Spacecraft Orbit Refueling Process

A maintainable refueling vehicle is the future development direction of the space system. In the process of fuel filling, capture docking, and configuration transformation, the most significant factor that affects the attitude control of the system is the continuous or sudden change of the angular torque of the system. In this paper, we study the control system of the variable mass body in an on-orbit service in the process of configuration transformation. The large errors of torque of inertia may make the narrow sense TEA (torque equilibrium attitude) of the system deviate greatly from the earth-oriented attitude. In order to avoid the error caused by the partial linearization of the system model, the feedback linearization method of the nonlinear system is used to design the controller to realize the tracking of the narrow sense TEA in the process of configuration transformation. Different from the traditional attitude control method, in order to avoid the high cost of the control system caused by the change of system mass characteristics and the change of system angular torque, CMG (control moment gyroscope) angular torque is introduced into the controller. We design a joint controller of attitude control and angular torque management, which can effectively stabilize the system and reduce the angular torque saturation of the attitude control system during the on-orbit service.

Nonlinear dynamic modeling and responses of a cable dragged flexible spacecraft

The space debris removal system (SDRS) of tethered space tug is modelled as a cable dragged flexible spacecraft. The main goal of this paper is to develop a dynamic modeling approach for mode characteristics analysis and forced vibration analysis of the planar motion of a cable dragged flexible spacecraft. Solar arrays of the spacecraft are modelled as multi-beams connected by joints with additional rotating spring where the nonlinear stiffness, damping and friction are considered. Using the Global mode method (GMM), a novel analytical and low-dimensional nonlinear dynamic model is developed for vibration analysis of SDRS to enhance the design capacity for better fulfillment of space tasks. The linear and nonlinear partial differential equations that governing transverse vibration of solar arrays, transverse and longitudinal vibrations of cable are derived, along with the matching and boundary conditions. The natural frequencies and analytical global mode shapes of SDRS are determined, and orthogonality relations of the global mode shapes are established. Dynamical equations of the system are truncated to a set of ordinary differential equations with multiple-DOF. The validity of the method is verified by comparing the natural frequencies obtained from the characteristic equation with those obtained from FEM. Interesting mode localization and mode shift phenomena are observed in mode analysis. Dynamic responses of the system excitated by fluctuation of attitude control torque and short-time attitude control torque are worked out, respectively. Nonlinear behaviors are observed such as hardening, jump and super-harmonic resonances. Residual vibration of the overall system with considering the varous values of nonlinear stiffness, damping coefficient and friction coefficient has shown that the nonlinearity of joints has a great influence on the vibration of the overall system.

Reaction Wheel Torque Distribution Algorithm for Enhancing the Agility of the Spacecraft using Null motion

Earth observation satellites require rapid attitude maneuvering and stringent pointing accuracy for imaging payloads. Four numbers of Reaction Wheel (RW) are usually configured in standard tetrahedral configuration for attitude manoeuvring as well as for fine pointing. The Reaction wheel maximum torque is one of the critical parameters that determine the agility of the spacecraft. Three axis attitude control torque signal is distributed to 4 Reaction wheels using standard pseudo inverse solution which minimizes the sum of the square of the individual wheel torque rather than the maximum of the individual wheel torque. The pseudo inverse solution leads to wheel saturation even when the total momentum does not reach the envelope and there by limiting the agility of the spacecraft. In this paper, new approach is adopted to optimally utilize the full capacities of RWs and thereby enhancing the agility of the spacecraft. The new algorithm is based on the null motion of the reaction wheel torque alignment matrix. The proper choice of the null vector results in equal wheel torque demand when combining with the pseudo inverse solution. The maneuver & control algorithms are integrated together with new distribution method and have been demonstrated for typical rest to rest and rate to rate maneuver.

Attitude control of the low earth orbit CubeSat using a moving mass actuator

This paper investigates an attitude control method for the CubeSat using a moving mass actuator to solve the problem of the strong aerodynamic disturbance in low Earth orbit. The rotational and translational equations are derived for the CubeSat with three moving masses, and their dynamic effects are analyzed. A magnetorquer is used to prevent the underactuation of the attitude control system. The movement of moving masses is slowed down by using a discrete double-loop Proportion Integral Differential control method, thereby reducing the fast time-varying additional disturbance. A nonlinear observer is used for the precise estimation of the slow time-varying disturbance. Notably, the ideal attitude control torque is allocated to two actuators by using the proposed control allocation algorithm. Numerical simulation indicates that the attitude convergence accuracy is up to ±0.1° despite the uncertain dynamics, unknown disturbances, and dynamic effects. The results verify the feasibility of the proposed control method.

Aerodynamic/reaction-jet compound control of hypersonic reentry vehicle using sliding mode control and neural learning

Considering the saturation of the aerodynamic surface (ADS) and the limitation of the actuator input, the paper investigates the aerodynamic/reaction-jet compound attitude control for the hypersonic reentry vehicle subject to poor aerodynamic maneuverability. The attitude control torque is preferentially allocated to the ADS. If the ADS is saturated, the remaining control torque is distributed to the reaction control system (RCS) through the control distribution algorithm to assist the ADS. For the dynamics uncertainty in the calculation of control torque, the terminal sliding mode (TSM) controller based on the back-stepping frame is designed to improve the robust performance and achieve the finite-time convergence. Furthermore, the predefined-time TSM controller is constructed with the online-data neural learning and the disturbance observer, which guarantee the predefined-time convergence and complete the effective approximation of system dynamics uncertainty. The stability of the attitude system is mathematically proved via Lyapunov function, while the simulation results are provided to show the effectiveness of the proposed controllers.

Design of Sliding Mode Techniques for a CMG-based Testbed Attitude Control System

Abstract Precise pointing accuracy and rapid maneuvering are two key features for attitude control missions of small spacecraft. Control moment gyroscopes (CMGs) are applied as ideal actuator for large torque output capability but are usually limited due to the problem of inherent mechanical singularity. This paper proposes a robust attitude control methodology, based on Sliding Mode Control (SMC) techniques, in presence of CMG practical restrictions and disturbances. Two second-order SMC techniques are designed, to guarantee accuracy and limited convergence time. Moreover, attitude control torques are generated by means of four single gimbal CMGs in pyramidal configuration, considering the design of an experimental testbed. The effectiveness of the proposed methodologies are shown in simulations, for different mission scenarios, including singularity points.

Fault Modeling, Estimation, and Fault-Tolerant Steering Logic Design for Single-Gimbal Control Moment Gyro

This brief addresses the single-gimbal control moment gyro (SGCMG) fault modeling, estimation, and tolerant-control steering logic design problem, aiming at enhancing the reliability and safety of spacecraft attitude control systems. The SGCMG is modeled as a two-loop system, including a wheel speed control loop and a gimbal rate control loop. Each loop contains an electrical motor (EM) and its corresponding variable speed drive (VSD), which may suffer from faults. By analyzing and modeling potential faults of the EM-VSD system, the SGCMG fault model is further developed. Then, a local adaptive fault estimator is proposed to reconstruct the total time-varying fault effects of each SGCMG. It is proven that the gimbal angle estimation error and fault estimation error converge to small compact sets containing zero. Moreover, leveraging estimated fault effects, a fault-tolerant steering logic is further developed to allocate the commanded attitude control torque properly such that the gimbal rate constraints are satisfied, and fault effects are compensated. To verify the proposed fault estimator and fault-tolerant steering logic, numerical simulations are carried out on an SGCMG-actuated spacecraft.

Optimal control of approaching target for tethered space robot based on non-singular terminal sliding mode method

In the tethered space robot approaching phase, the tether can provide attitude control torque and orbit control force, it can reduce the fuel consumption of the gripper, but there is operation coupling for the tether is not always pass through the centroid of the gripper. And the field of view of the camera is limited, the tracking error of attitude and position can make the target deviate from the camera view field. Aiming to above problems, a position and attitude coordinated control scheme based on non-singular terminal sliding mode method is designed for the approaching task. The dynamic model of tethered space robot based on lumped mass model and elastic rod model are established, respectively. Combining the dynamic model of tethered space robot based on elastic rod model with the relative motion model of the target, a Gaussian pseudo spectral motion planning model is established, the planning model takes into account the constraints of control input and the field of view of gripper. Solving the programming model, the optimal approaching trajectory of the tethered space robot is obtained. Then, a trajectory tracking controller is designed based on non-singular terminal sliding mode method. The simulation results show that the proposed control scheme can realize the accurate position and attitude tracking of the optimal trajectory.

Compound fault-tolerant attitude control for hypersonic vehicle with reaction control systems in reentry phase

In this paper, a novel compound fault-tolerant attitude control (FTC) scheme is proposed for reentry hypersonic vehicles with aerodynamic surfaces and reaction control systems (RCS) in the presence of parameter uncertainties, external disturbances and aerodynamic surfaces faults. Aerodynamic surfaces work as the primary actuators and RCS serve as auxiliary actuators. When aerodynamic surfaces cannot provide the required attitude control torque due to low dynamic pressure or faults, RCS are activated to assist aerodynamic surfaces to generate the residual torque. A nonlinear disturbance observer-based sliding mode controller is designed to calculate the required attitude control torque which can handle the parametric uncertainties and external disturbances together. The quadratic programming method is applied to obtain the optimal aerodynamic surfaces deflections from the required control torque. An innovative fuzzy rule-based decision-making system is design to solve the RCS control allocation problem, which is conceptually easy to understand and computationally efficiently compared with existing approaches. Based on quantized control theory, the closed-loop control system stability is rigorously analyzed. Simulation results are given to demonstrate the effectiveness and efficiency of developed FTC scheme.

Fast Attitude Maneuver of a Flexible Spacecraft with Passive Vibration Control Using Shunted Piezoelectric Transducers

This paper is concerned with designing a bang-bang control input to perform a quick rotational maneuver of a rigid spacecraft hub connected with flexible appendages. The control design is based on only the rigid body mode making it very simple to design and at the same time achieve the quickest maneuver possible. The induced vibrations are suppressed using piezoelectric transducers bonded to the appendages and connected to an electric circuit with the objective of converting the vibrational energy to electrical energy and then dissipating it using passive electric elements, such as a resistance and an inductor. The proposed control design method is applied to a spacecraft containing a rigid hub and flexible appendages. The attitude control torque is produced using either the reaction wheels contained inside the rigid hub or jet thrusters mounted outside it. The control design process starts with deriving the nonlinear partial differential equations of motion for the spacecraft using Hamilton’s principle which accounts for the electromechanical coupling and the presence of resistive or resistive-inductive circuits. To simplify the analysis, the nonlinear ordinary differential equations of motion are then obtained using the assumed mode method. The effectiveness of the control design method is numerically tested on a spacecraft that is required to perform a quick attitude maneuver and, simultaneously, suppress the induced vibrations. The simulations show a quick and accurate maneuver has been achieved combined with very low levels of vibrations resulting from the reduced coupling between flexible and rigid motions as well as the damping added as a result of the passive shunt circuit. Furthermore, the resistance-inductance shunt circuit is shown to be more effective in damping the vibrations than the resistance shunt circuit.

RBF network based adaptive sliding mode control for solar sails

PurposeThe 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.Design/methodology/approachThe 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.FindingsThe 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.Practical implicationsThe 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.Originality/valueAdvanced 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.

Magnetic attitude control torque generation of a gravity gradient stabilized satellite

Magnetic torquer is used to generate a magnetic dipole moment onboard satellites whereby a control torque for attitude control purposes is generated when it couples with the geomagnetic field. This technique has been considered very attractive for satellites operated in Low Earth Orbit (LEO) as the strength of the geomagnetic field is relatively high below the altitude of 1000 km. This paper presents the algorithm used to generate required magnetic dipole moment by 3 magnetic torquers mounted onboard a gravity gradient stabilized satellite operated at an altitude of 540 km with nadir pointing mission. As the geomagnetic field cannot be altered and its magnitude and direction vary with respect to the orbit altitude and inclination, a comparison study of attitude control torque generation performance with various orbit inclination is performed where the structured control algorithm is simulated for 13°, 33° and 53° orbit inclinations to see how the variation of the satellite orbit affects the satellite's attitude control torque generation. Results from simulation show that the higher orbit inclination generates optimum magnetic attitude control torque for accurate nadir pointing mission.

A decentralized cooperative control scheme for a distributed space transportation system

In this paper, a new concept for a distributed space transportation system is proposed and a corresponding decentralized cooperative control scheme is investigated. First, for a leaderless team of homogeneous space robots transporting a large object in orbit, a systematic control architecture that includes information flow is developed. Second, based on relative orbit dynamics, a rendezvous guidance law and a rigid formation control law are designed, and the necessary communication topology for the space robot team is discussed. Third, to guarantee the consensus of the motion of the large object with the robot team, both orbital maneuver control and attitude control for the large object are studied Emphasis is placed not on the attitude control law but on the force distribution problem, for which an algorithm exploiting a special property of trigonometric functions is proposed to transfer the necessary attitude control torque to the distributed forces. To support the above control method, an estimation of the motion of the formation center contributed by each robot in a decentralized manner is developed using a Kalman filter. Fourth, the robustness of the system to the failure of one robot is analyzed, and four effective typical fault response modes are proposed. Finally, numerical simulations validated the performance, robustness and practicability of the developed control scheme.

Torque distribution algorithm for effective use of reaction wheel torques and angular momentums

In attitude control of spacecraft using more than three reaction wheels, the distribution of the attitude control torque to the wheels is not unique because of the redundancy. There are several wheel torque distribution algorithms which optimize the wheel torques or other factors. In particular, the optimal torque distribution algorithm is acknowledged as algorithm which minimizes the maximum wheel torque. This algorithm is advantageous to make maximum use of the wheel torques, because each wheel torque must be lower than the wheel torque capability and torque is the primary driver in many cases. However, as a result of minimizing the maximum wheel torque, the distribution of the wheel angular momentums is not calculated by a similar formula for the wheel torques distribution. In other words, the wheel angular momentums cannot be derived from the current attitude angular momentum. When certain wheel reaches maximum angular momentum earlier than the other wheels, this prohibits maximum use of the other wheels' capability. Therefore, minimizing the maximum wheel torque is not always effective when other constraint such as angular momentum matters.Recently, it has become more important that both wheel torques and angular momentums are used more effectively in order to improve the performance of the spacecraft agility, such as the high angular acceleration and rate, by using minimum spacecraft resources (i.e. minimum number of wheels which satisfies certain agility requirements). In this paper, shown is the wheel torque distribution algorithm which is effective in terms of both the wheel torques and angular momentums as much as possible. In the proposed algorithm, the wheel torques/angular momentums distributed from the current attitude torque/angular momentum can be optimal for particular direction like the spacecraft X/Y/Z axis. In addition, it is shown by numerical simulation that the proposed algorithm improves the usage of attitude control angular momentum by up to 60% compared to the optimal torque distribution algorithm.

Attitude coordination of multi-HUG formation based on multibody system theory

Application of multiple hybrid underwater gliders (HUGs) is a promising method for large scale, long-term ocean survey. Attitude coordination has become a requisite for task execution of multi-HUG formation. In this paper, a multibody model is presented for attitude coordination among agents in the HUG formation. The HUG formation is regarded as a multi-rigid body system. The interaction between agents in the formation is described by artificial potential field (APF) approach. Attitude control torque is composed of a conservative torque generated by orientation potential field and a dissipative term related with angular velocity. Dynamic modeling of the multibody system is presented to analyze the dynamic process of the HUG formation. Numerical calculation is carried out to simulate attitude synchronization with two kinds of formation topologies. Results show that attitude synchronization can be fulfilled based on the multibody method described in this paper. It is also indicated that different topologies affect attitude control quality with respect to energy consumption and adjusting time. Low level topology should be adopted during formation control scheme design to achieve a better control effect.

LQR/LQG attitude stabilization of an agile microsatellite with CMG

PurposeThe 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.Design/methodology/approachIn 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.FindingsNumerical 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.Practical implicationsThe 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.Originality/valueThe originality of this paper is designing a new attitude stabilization system for an agile microsatellite using LQR and LQG controllers.

A Vector Measurement-based Angular Velocity Estimation Scheme for Maneuvering Spacecraft

A new practical approach to estimate the body angular velocity of maneuvering spacecraft using only vector measurements is presented. Several algorithms have been introduced in previous studies to estimate the angular velocity directly from vector measurements at two time instants. However, these direct methods are based on the constant angular velocity assumption, and estimation results may be invalid for attitude maneuvers. In this paper, an estimation scheme to consider attitude disturbances and control torques is proposed. The effects of angular velocity variation on estimation results are quantitatively evaluated, and an algorithm to minimize estimation errors is designed by selecting the optimal time interval between vector measurements. Without losing the simplicity of direct methods, the design parameters of the algorithm are restricted to the expected covariance of disturbances and the maximum angular acceleration. By applying the proposed estimation scheme, gyroscopes can be directly replaced by attitude sensors such as star trackers.

Three-axis attitude control by two-step rotations using only magnetic torquers in a low Earth orbit near the magnetic equator

This study proposes a novel method for three-axis attitude control using only magnetic torquers (MTQs). Previously, MTQs have been utilized for attitude control in many low Earth orbit satellites. Although MTQs are useful for achieving attitude control at low cost and high reliability without the need for propellant, these electromagnetic coils cannot be used to generate an attitude control torque about the geomagnetic field vector. Thus, conventional attitude control methods using MTQs assume the magnetic field changes in an orbital period so that the satellite can generate a required attitude control torque after waiting for a change in the magnetic field direction. However, in a near magnetic equatorial orbit, the magnetic field does not change in an inertial reference frame. Thus, satellites cannot generate a required attitude control torque in a single orbital period with only MTQs. This study proposes a method for achieving a rotation about the geomagnetic field vector by generating a torque that is perpendicular to it. First, this study shows that the three-axis attitude control using only MTQs is feasible with a two-step rotation. Then, the study proposes a method for controlling the attitude with the two-step rotation using a PD controller. Finally, the proposed method is assessed by examining the results of numerical simulations.