Attitude Maneuver Research Articles

Optimal control of a tetrahedral configured reaction wheels for Quaternion rotation based on rigid satellite model

Reaction wheel is the most popular actuator for the attitude manoeuvring of a satellite system due to its compact size, reliability and able to produce precise torque. However, for redundancy purposes, the reaction wheels are assembled in a tetrahedral configuration that is able to generate torque in any direction even when one of the reaction wheel fails. The tetrahedral configured reaction wheel causes the maximum torque generation to be uneven in any direction. Hence, the quaternion rotation is completed at different rate in different direction. In order to optimize the rotation rate, an optimal control using iterative method via GPOPS toolbox. It is then compared with the traditional Eigen-axis Quaternion Feedback control. The improvement shown by the optimal control to be between 3.49% to 25.11% improvement in manoeuvre time depending on the direction of manoeuvre. The optimal control is able to outperform the traditional control method.

Constrained attitude maneuvers on [formula omitted]: Rotation space sampling, planning and low-level control

In this paper, we propose a novel framework that provides a systematic strategy to regulate the rigid body attitude on SO(3) within a generic constrained attitude zone. The proposed control scheme consists of three components: sampling, planning and low-level control. Specifically, an overlapping cell-like sampling for the attitude configuration space SO(3) is built and further reformulated to a graph model. Based on this abstraction, a complete graph search algorithm is utilized to generate a feasible path in the graph model. Both sufficient and necessary conditions on finding a feasible path are presented. Furthermore, to facilitate the control design, the point-to-point path is transformed into a smooth reference trajectory along the geodesics. Finally, a saturated low-level control law is formulated to robustly track the desired trajectory. Simulations demonstrate the effectiveness of the proposed control approach.

Attitude Maneuver Planning of Agile Satellites for Time Delay Integration Imaging

In the prevailing operations of agile satellites for time delay integration (TDI) imaging tasks, two degrees of freedom (DOFs) in the attitude maneuver are determined by tracking the transient target point, whereas the third DOF is used in the control of the drift angle, which is calculated within the sensor focal plane as a feedback. However, optimal attitude maneuver planning is rarely considered during a TDI imaging mission to improve the efficiency of agile satellites. In this paper, a general attitude maneuver planning for agile satellites in the TDI imaging of an arbitrary target strip is formulated, where the energy consumption is minimized while the TDI imaging quality is guaranteed. Coupled with the attitude dynamics, the highly nonlinear constraints required by the TDI imaging principal make this planning hard to solve. By introducing a parameterized time-mapping technique and a compact approach to the attitude maneuver that ensures zero drift angle, the problem is relieved in both complexity and computational burden. The validity of this approach is verified by various imaging simulations.

Model Predictive Sliding Control for Finite-Time Three-Axis Spacecraft Attitude Tracking

This paper deals with the robust optimal three-axis attitude tracking problem of spacecrafts. To this aim, the exceptional features of two advanced control approaches, i.e., nonlinear terminal sliding mode control (TSMC) and linear model predictive control (MPC), are combined in a compound double layer structure. The proposed approach addresses the global stability and robust attitude tracking of near-polar orbit spacecrafts actuated by a set of reaction wheels and subject to unknown disturbances, uncertainties, and actuators’ saturation. The TSMC controller is designed to guarantee the robust three-axis attitude tracking goal in the presence of disturbances and uncertainties provided that proper target points are available. The next part of the compound controller takes the advantages of the MPC technique (i.e., preknown attitude maneuvers and constraints) into account to provide the required optimal target points for the lower level TSMC controller. Despite the fact that the spacecraft has nonlinear nature, we have analytically derived conditions under which the actuators’ constraints satisfaction is guaranteed a priori . Finally, a set of simulation results are provided to illustrate the effectiveness and performance of the proposed method.

Stabilization of a large flexible spacecraft using robust adaptive sliding hypersurface and finite element approach

A novel adaptive sliding hypersurface with an adjustable dead zone scheme for attitude control of a spacecraft with large flexible sun oriented solar panels is employed to overcome the difficulties arising from the measurement of flexible dynamics coordinates. Equations of motion for multi-axis attitude maneuver of a large rigid-flexible system are derived utilizing Lagrangian formulation based on a hybrid system of coordinates. The flexible appendages are modeled as elastic plates to examine flexural deformations employing the finite element method. A proposed synthesized hybrid-sliding surface in conjunction with an adaptation law can perform desired mission properly, in which the excitation of high-frequency flexible modes, external disturbances, and system parameter uncertainties are taken into account. It is noteworthy that the design of the proposed controller does not require the exact model of the system’s dynamic and can be performed for every general nonlinear time-varying system in the real environment. Within the Lyapunov analysis and Barbalat’s lemma, the stability of the global closed-loop system is guaranteed. The comparative results expressed as flexible structures deflection shows significant improvements in terms of robustness of the highly flexible system to the vibration energy excitation.

Finite-time attitude maneuvering and vibration suppression of flexible spacecraft

The finite-time attitude maneuvering control and vibration suppression for the flexible spacecraft with external disturbances, inertia uncertainties and input saturation are investigated in this paper. Firstly, the input shaping technique is applied to suppress most of the vibration caused by flexible appendages. Then, for the nominal attitude control system, a multivariable continuous sliding mode controller is applied to guarantee the finite-time convergence. Next, unlike the conventional discontinuous design of integral sliding mode, a continuous compensated controller with modified multivariable adaptive twisting algorithm is proposed to reject the lumped uncertainty, including disturbances, uncertainties, saturation and residual vibration. The controller can realize finite time convergence and the chattering suppression, while there’s no need for the priori knowledge of the lumped uncertainty. A rigorous proof of the finite time stability of the closed-loop system is derived by the Lyapunov method. Finally, the efficiency of the proposed method is illustrated by numerical simulations.

Dynamic Near-Optimal Control Allocation for Spacecraft Attitude Control Using a Hybrid Configuration of Actuators

This paper proposes a novel dynamic near-optimal control allocation scheme with combination of a saturated baseline controller for spacecraft attitude control using single-gimbal control moment gyros (CMGs) and reaction wheels. First, a saturated controller is proposed to stabilize the nominal system in the presence of actuation mismatch. Aided by a state-dependent variable, a dynamic control allocator is then proposed that allows for smooth switching between two actuation sets. Unlike the previous static constraint optimization formulations, the control allocation augments its performance function with penalty terms in order to enforce individual input constraints and configuration singularity avoidance. Moreover, this dynamic control allocation is implemented with an online update law, which has a modest computational complexity compared to its numerical optimization counterpart. The closed-loop boundedness is guaranteed by a constructive Lyapunov-design method. Simulation results demonstrate that during the attitude maneuvers, the input saturation constraint and the active avoidance of CMG singularities are enforced with a relatively small computational cost.

Optimal design and robust analysis of a net of active devices for micro-vibration control of an on-orbit large space antenna

Large deployable antennas are required for the advancement of space communications, Earth observation, radio astronomy and deep space exploration. Most contemporary space antennas have exceeded the size of launching vehicles, leading to the necessity of stowed concepts to overcome the limitation. Many structural models have been investigated by different organizations. Generally, mesh deployable reflectors are currently more mature compared to other foldable solutions. This will be the topic of this paper. Orbital disturbances, perturbations originated by on-board sources and thermal deformations affecting the deployed configuration can deteriorate the accuracy of the communications system. Undesired dynamic behaviour of structural components has to be predicted and counteracted. Therefore, vibration control is a key technology to correct the distortions altering the proper functioning of the system. An intelligent active structure is introduced as a structure configured with distributed actuators and sensors and guided by a controller to modify the dynamic response of the system. In this paper, the supporting structure of a very large mesh reflector is described. A FEM formulation is adopted to assemble the frame and it is validated by comparing it with commercial codes. According to the adopted model, the active elements can be embedded in the middle of the truss elements. Of course, active control of all the devices at the same time requires a control effort which could be not affordable in space applications. However, the effectiveness is not the same for all the actuators. In this study, two cascade optimization procedures are performed to assess the best positioning and velocity feedback gains of the actuators that must be controlled to damp vibrations. In particular, a gradient-based technique is used after both a Genetic Algorithm and a Reinforcement Learning methodology to find the best set of gains for the controller. The objective function is set as a weighted sum of power consumption of the actuators. As a study case, the spacecraft implements a velocity feedback strategy when performing a generic attitude manoeuvre to coordinate the simultaneous action of the devices to ensure the damping performance of the system is enhanced. In addition, the impact of damaged actuators and uncertainties on the mechanical properties of the passive structure has been discussed and analysed.

Trajectory planning of dual-arm space robots for target capturing and base manoeuvring

This paper presents a trajectory planning method for a dual-arm space robot that is free-floating and receives no actuation from the base spacecraft. The task is to capture a target by a mission arm and accomplish base manoeuvring with the coupling effects by a manoeuvring arm. Three manoeuvring cases for the base are discussed concerning attitude regulation, attitude tracking and position tracking. In all cases, an optimization method is applied to solve for the joint velocity of the arms, which ensures that the joint states of the mission arm obtain the desired states for capturing and control the motion of the base. Moreover, the collision avoidance of the manoeuvring arm with the base and mission arm is considered an inequality constraint. In particular, the joint velocity of the mission arm is represented by a new variable to remove the disturbance caused by the mission arm on the attitude of the base which makes it easier to complete the attitude manoeuvring case. Finally, numerical simulations are carried out by the planar and spatial dual-arm robots to validate the efficiency of the control strategy during both capturing and base manoeuvring.

Relative Motion Dynamics Modeling and Control of DFP Vibration Isolation Spacecraft

In order to better meet the future high precision task requirements, the DFP(Disturbance-Free Payload) spacecraft composed of non-contact PM(Payload Module)and SM(Support Module)is taken as the object to study the relative motion dynamics modeling and control between the two modules and verify the system vibration isolation performance. Firstly, the force and torque expressions of the two modules are derived by simplifying the configuration and analyzing the stress. In view of that couple effect, the relative motion dynamics equations between two modules of DFP spacecraft with high model accuracy, and simple and uniform format are established with the dual quaternions. Based on this, the PD control law is designed, and the relative motion of PM and SM could meet DFP spacecraft working requirements when the measurability of control quantity and the measurement error of sensors were taken into account. Simulation results verify the advantage of vibration isolation and attitude maneuverability of DFP spacecraft.

Robust finite-time adaptive control algorithm for satellite fast attitude maneuver

A robust adaptive finite-time controller for satellite fast attitude maneuver is proposed in this paper. The finite-time sliding mode is proposed based on standard sliding mode, hence the terminal convergence rate could be largely improved, and the inherent robustness to some typical perturbations is maintained. A new method is proposed to deal with the singularity issue based on the property of Euler rotation. In order to deal with inertia matrix uncertainty, finite-time adaptive law for inertia matrix estimation variables is proposed. Considering that the variable estimation system has no direct feedback, an auxiliary state converging slower than system is designed to achieve finite-time stability. The overall global finite-time stability is proved by Lyapunov method and the performance of the controller is demonstrated by numerical simulation results.

Semi-active vibration control of a rotating flexible plate using stiffness and damping actively tunable joint

Lighter structures are increasingly required for flexible spacecraft. However, low frequency and large amplitude vibration problems are unavoidable due to external disturbances or attitude maneuvering especially when working under a microgravity environment. Therefore, a new methodology involving a semi-active technique using an actively tunable joint with variable stiffness and damping control, capable of handling such issues is proposed in this study. The incentive active joint was conceived with a compact structure, based on the electromagnetic direct drive principle. First, a dynamic model of rotating flexible plate was established. Then, a prototype was fabricated and tested. Finally, on the one hand, numerical simulations and experimental results indicated that, when the joint torsional stiffness changed, a frequency shift phenomenon occurred. On the other hand, two types of noncontact periodic vibration excitation methods along the rotation direction were proposed and the experimental results validated that the major frequency bandwidth of interference signal was 0.08–0.52 Hz with a significant vibration attenuation of 5.5–31.5 dB in effective bandwidth. Moreover, in the low frequency range (0.08–0.43 Hz), variable damping was found to be the main factor and variable stiffness was the secondary factor. However, in the high frequency range (around 0.52 Hz), variable stiffness was dominant and variable damping was inferior. These findings are expected to effectively suppress the low frequency and large amplitude vibration of solar panels with flexible joints.

A person-following nanosatellite for in-cabin astronaut assistance: System design and deep-learning-based astronaut visual tracking implementation

An intelligent, person-following nanosatellite is under development for in-cabin astronaut assistance in the China Space Station. It is named the Intelligent Formation Personal Satellite (IFPS). The satellite weighs 2.0 kg and is shaped as a sphere of diameter 230 mm. Fans and MEMS flywheels are used for its position and attitude maneuvering as inside the Space Station cabin it is a weightless and standard atmospheric pressure environment. The RGB-D camera and IMU based visual-inertial SLAM method is used to support its localization and navigation. The on-board information processing and computing hardware primarily consists of an embedded AI microprocessor and an FPGA. The satellite is designed to fly autonomously and follow the designated astronaut to offer immediate assistance. Thus, efficient and robust astronaut visual tracking is the most important prerequisite for supporting its basic person-following operating mode. We achieved this by further improving our previously proposed tracking algorithm that consists of a deep convolution neural network (DCNN)-based detection module and a probabilistic-model-based tracking module. The DCNN in the detection module was further improved through optimizations of lightweight network architecture design, parameters model compression and inference acceleration. While maintaining the originally high detection accuracy, the DCNN was optimized significantly in terms of memory, computation and power consumption to quite meet the engineering constraints in the development of IFPS. The complete pipeline of the astronaut visual tracking algorithm was also designed and implemented in the embedded AI microprocessor for online tracking application. Experimental results demonstrated the effectiveness of the proposed efficient and robust astronaut detection and tracking algorithm.

Semi-active vibration control of two flexible plates using an innovative joint mechanism

Abstract With the advancement of space technology, space flexible jointed appendages have been applied widely. The low-frequency vibration problem is inevitable because of the external disturbances or attitude maneuvering in micro-gravity environment. In this study, an innovative joint mechanism was proposed and a new methodology for semi-active technique of joint variable stiffness control was utilized to suppress the low-frequency vibration of flexible jointed appendages. The joint mechanism, based on the direct drive principle, exhibits a compact structure and integrates with an electro-permanent-magnet driving mechanism. First, the concept design of joint mechanism is introduced. Second, dynamic equations of active joint and two flexible plates are obtained using Lagrange equation and assumed mode method. Finally, two types of non-contact periodic vibration excitation methods along rotation direction are proposed. A self-made electro–magnetic vibration exciter was employed to achieve single-frequency excitation, and two fans were placed on both sides of flexible appendages to induce multi-frequency excitation. According to numerical simulations and experimental results, when the joint torsional stiffness changes, frequency-shift phenomenon occurs, and the natural frequency of the system varies subsequently. Furthermore, the experimental results demonstrated that the major frequency bandwidth of interference signal was 0.06–0.63 Hz with a significant vibration attenuation of 5.41–26.94 dB for joint-1 and 5.63–16.48 dB for joint-2 in effective bandwidth. The low frequency vibration of flexible appendages system can be attenuated robustly, which verifies the effectiveness of this semi-active methodology.

Improved Spacecraft Magnetic Attitude Maneuvering

A magnetometer is essential in spacecraft magnetic attitude control due to the need for magnetic field information to compute the control command. The measurements of the magnetometer, however, are usually affected by other electric currents in the spacecraft, especially those of the magnetic coils when they are turned on during actuation. As a result, magnetic rods and magnetometers are usually turned on at alternate times, resulting in a reduced duty cycle of the magnetic rods, and hence longer maneuver times. This paper presents a magnetic attitude control system with extended duty cycle and low magnetometer measurements frequency. Instead of measurements, a computed magnetic field strength vector is used for updating the control command during the duty cycle. Using the measured spacecraft rotational motion, and knowing the control torque command, it is possible to compute the magnetic field strength vector. These computations are corrected using magnetometer measurements at a lower rate. The Tikhonov regularization approach is implemented to solve the singular magnetic torque system. This algorithm is demonstrated via Monte Carlo numerical simulations to have faster attitude maneuvers and lower power consumption by the magnetic rods. Real data obtained from the Cascade Smallsat and Ionospheric Polar Explorer spacecraft are used for validation of the proposed approach.

Event-Triggered Adaptive Attitude Tracking Control for Spacecraft With Unknown Actuator Faults

This paper is devoted to attitude tracking control of fractionated spacecraft with wireless communication. We consider the practical case that the spacecraft suffers from uncertain inertia parameters, external disturbances, and even unknown and time-varying actuator faults. Within the framework of the backstepping method, a novel event-triggered adaptive fault-tolerant control scheme is proposed. In our design, an event-triggering mechanism is introduced to determine the time instants for communication, which successfully avoids continuous communication and Zeno phenomenon. Then, with the aid of a bound estimation approach and a smooth function, the impacts of the actuator faults, as well as the network-induced error, are effectively compensated for. Moreover, by employing the prescribed performance control technique, it is shown that the attitude tracking errors can converge to predefined arbitrarily small residual sets with prescribed convergence rate and maximum overshoot, no matter if there exist unknown actuator faults. Compared with conventional adaptive attitude control schemes, the proposed scheme significantly reduces the communication burden, while providing high reliability and stable, rapid, and accurate response for attitude maneuvers. Simulation results are presented to illustrate the effectiveness of the proposed scheme.

Active Fault-Tolerant Control System Design for Spacecraft Attitude Maneuvers with Actuator Saturation and Faults

This paper designs an active fault-tolerant control system for spacecraft attitude control in the presence of actuator faults, fault estimation errors, and control input constraints. The developed fault-tolerant control system is able to detect the actuator fault without false alarms caused by external disturbances, and also estimate the total fault effects accurately through an indirect fault identification approach, in which an auxiliary variable is utilized to build the relation between fault and system states. Once the fault identification is completed with certain degree of reconstruction accuracy, a fault-tolerant backstepping controller using the nonlinear virtual control input is reconfigured to accommodate the detected actuator faults effectively, in spite of actuator saturation limitations and fault estimation errors. Numerical simulation is carried out to demonstrate that the proposed active fault-tolerant control system is successful in fault detection, identification, and controller reconfiguration for handling actuator faults in attitude control systems.

Control of non-cooperative spacecraft in final phase proximity operations under input constraints

This paper addresses the approaching trajectory control problem for the final phase of non-cooperative spacecraft proximity operations in the presence of disturbances and even input saturation. As a stepping stone, a novel line-of-sight tracking model is developed to describe attitude maneuver towards the target in the proximity. Then, a two-stage artificial potential function is employed to enable the pursuer spacecraft to comply the safety trajectory during approaching the target. In each stage, gradients of potential function are analyzed and the local minimum problem associated with potential function can be addressed. Furthermore, a saturated control law is proposed to ensure a reliable real-time tracking measurement of the target, and the controller rigorously enforces safe path and actuator magnitude constraints. The associated stability proof is accomplished through Lyapunov method. Simulations are carried out to evaluate the effectiveness of the proposed method.

Characterization of deployable ultrathin composite boom for microsatellites excited by attitude maneuvers

Composite booms are often adopted onboard spacecraft to modify inertia properties and/or to move instrument as far as possible from the influence of the platform bus. In the frame of the current trend towards smaller spacecraft, also microsatellites are increasingly equipped with these deployable appendages. Due to limited volume and mass available on these platforms, the design of the boom is more challenging and a validation of the proposed system is crucial. The paper details, from the structural dynamics point of view, the numerical and experimental validation campaign for a 1-meter boom purposely designed and built in composite material for a microsatellite platform. The experimental validation was carried out by using a free-floating platform and while simulating the loads experienced during an in-orbit attitude re-orientation maneuver. The complete architecture of the platform, including solar panels’ appendages with related flexibility issues as well as the thrusters for actuation, provides for the credibility and the interest of the validation outcome.

Solar sailing CubeSat attitude control method with satellite as moving mass

Low-cost solar sailing CubeSats for deep-space exploration are relatively inexpensive, have a short development time, and can be launched more often. However, attitude control of a solar sailing CubeSat is challenging due to the low controllability of the CubeSat itself. In this paper, a new attitude control method for the solar sailing CubeSat where solar radiation pressure torque, which is a torque generated by the satellite's center of mass offset with respect to its center of pressure, is used as the controlling torque. In this study, a mechanical structure to adjust the center of mass with the CubeSat as a movable mass is presented. The CubeSat's position on the solar sail's surface plane can be adjusted to control the attitude about the roll and pitch axes. Therefore, a dynamics model for the solar sailing CubeSat is derived and an attitude controller is developed. The method is demonstrated for the case of a 10 × 10 m2 sail with a 12-U CubeSat. The simulation results show that the solar sail attitude maneuver and sun pointing stability control are feasible when the initial and desired states satisfy the constraints. As a semi-active control method, the system can be combined with a micro flywheel control system for the attitude stability control of the CubeSat.