Attitude Maneuver Research Articles

Time-optimal attitude planning for spacecraft with movable parts using second order cone programming

The objective of this paper is to plan a time-optimal attitude trajectory in a relatively short time for spacecraft with movable parts under complex pointing constraints. The constrained attitude planning problem is first established, then the constraints on attitude maneuvers are described. The time-optimal attitude planning for rigid spacecraft body is decoupled into geometric level and dynamic level through the addition of the path scalar. To obtain a given safe path in quaternion space, RRT*-Smart path planning and spherical and quadrangle (SQUAD) interpolation are carried out at the geometric level. The time-optimal parameterization for a given path is reformulated as a second order cone programming (SOCP) problem at the dynamic level through a nonlinear change in variables and the addition of various convexity-preserving extensions. In order to overcome the issue of high system dimension caused by the movable parts, a decoupled attitude planning method is designed to generate the quaternion trajectory of the spacecraft body and the rotation angle trajectory of the movable parts respectively. In comparison to GPOPS, the simulation results demonstrate that the proposed method can plan a suboptimal solution in milliseconds. Furthermore, the performance indicator for energy consumption is better.

Vibration Control of Innovative Lightweight Thermoplastic Composite Material via Smart Actuators for Aerospace Applications

Piezoelectric actuators and sensors can be incorporated into aerospace structures to suppress unwanted flexible oscillations. These devices need to interact with various passive structures, including innovative materials such as thermoplastic composites, which offer several advantages over traditional options. This study explores the application of a piezoelectric-based vibration control system on a lightweight carbon-reinforced thermoplastic material. Numerical and experimental investigations are conducted to assess the mechanical properties and damping behavior of the composite. As a case study, an equivalent orthotropic shell laminate is developed to facilitate finite element modeling of two composite solar panel structures equipped to a spacecraft. Moreover, an electro-mechanical formulation is implemented to integrate smart actuators and sensors onto the composite hosting structure. Finally, the efficiency of the active vibration control system is assessed when significant vibration perturbations are caused on the panels by rigid–flexible dynamics coupling during agile attitude maneuvers. The results demonstrate the damping factor of the material can be noticeably improved, making the proposed system a promising technological solution for further aerospace applications.

Dynamic Modeling and Analysis of Spacecraft with Multiple Large Flexible Structures

An analytical dynamic model is presented for a spacecraft with multiple large flexible structures. Based on the partial differential equations (PDEs) of the motion of the solar panel and deployable arm, the governing equations of the main-body and deployable antenna and the boundary conditions at each end point are used to obtain the frequency and mode shapes of the system. Then, the ordinary differential equations (ODEs) of the system can be obtained from the orthogonality relations and mode shape. The influence of the deployable antenna on the frequencies and mode shapes of the spacecraft is investigated. The frequency veering and mode interchanged phenomenon are observed with the variation of the diameter of the deployable antenna. Using the ODEs, the dynamic responses of the spacecraft are calculated to study the influence of the control torque on the attitude and position of the antenna in the attitude maneuver.

Attitude control for a full-scale flexible electric solar wind sail spacecraft on heliocentric and displaced non-Keplerian orbits

The electric solar wind sail (E-sail) spacecraft is a fuel-free propulsion system with promising potential in deep space explorations and solar activity monitoring missions. It is challenging to conduct attitude control for a full-scale flexible E-sail spacecraft due to its complex dynamical characteristics, including the spatial multi-scale, high-speed spinning, and large deformations. In this work, a double-loop attitude control strategy for the flexible E-sail spacecraft with high accuracy and efficiency is proposed. The E-sail spacecraft is firstly modeled by the referenced nodal coordinate formulation, where the rotations are described by the Lie group approach. A state-feedback controller developed from the Lyapunov stability theory is prescribed as the inner loop. The gain is obtained by solving a semidefinite programming problem with linear matrix inequality constraints, which is globally optimal and numerically reliable. A multi-variable proportional-integral-derivative controller is then introduced as the outer loop for flexibly enhancing the control performance. Several missions of the E-sail spacecraft, including attitude maneuvers on heliocentric and displaced non-Keplerian orbits, are studied to show that the desired attitude adjustments are accurately and efficiently achieved without exciting high-frequency oscillations. The double-loop control strategy has great prospects in designing and implementing orbital transformation and long-term stable Sun-orbiting missions.

Integrated attitude and shape control for OrigamiSats with variable surface reflectivity

OrigamiSats, a new concept in solar sailing, are origami spacecraft with reflective panels that, when flat, operate as a conventional solar sail. Shape reconfiguration, i.e. “folding” of the origami design, allows the OrigamiSat to change operational modes, performing different functions as per mission requirements. For example, a flat OrigamiSat could be reconfigured into the shape of a parabolic reflector, before returning to the flat configuration when required to operate again as a solar sail, providing propellant-free propulsion. The attitude dynamics and shape reconfiguration of OrigamiSats are known to be highly coupled, thus presenting a challenge from a control perspective. This paper investigates the problem of integrating attitude and shape control of a Miura-fold pattern OrigamiSat through the use of variable reflectivity, allowing differences in solar radiation pressure to be used to enact shape reconfiguration and attitude manoeuvres. A closed-loop feedback controller is presented which combines and balances the attitude and shape control requirements, and gain-scheduling is implemented to address some specific features of the system dynamics. Numerical simulations of the multibody dynamics of the system are used to test the proposed controller and simulations of some example manoeuvres are performed which demonstrate the system’s performance.

Specific Tracking Control of Rotating Target Spacecraft Under Safe Motion Constraints

This paper considers a space mission that a chaser spacecraft approaches and tracks a rotating target at close range under safe motion constraints. The control objective is to keep the chaser hovering at the desired relative position above the specific point of the target and ensure the chaser's line-of-sight pointing to the target. During the proximity process of position and attitude maneuver, two kinds of safe geometric constraints are essential to be considered. One is collision avoidance for position constraint, and the other is the pointing forbidden zone for attitude constraint. Superquadrics curved surface is used to describe the position forbidden. Pointing forbidden is represented by inequality constraint based on SO(3). Combining the artificial potential function with backstepping control methodology and adaptive sliding-mode, a safety robust control law is designed. The proposed control scheme guarantees the chaser to track the target under safe geometric constraints. In addition, by using the dynamic specific point under observation, the proposed control scheme can be applied to other observation missions like close flying around. Finally, simulation results illustrate the effectiveness of the control scheme.

Finite-Time Refined Antidisturbance Velocity Tracking Control for Gimbal System of Control Moment Gyros With Harmonic Drive

This paper is concerned with high-precision and fast velocity-tracking control for the harmonic drive-based gimbal system of control moment gyro (CMG). As a major challenge in the controller design, the simultaneous presence of multiple matched and mismatched disturbances should be carefully addressed. Specifically, the rotor imbalance inherent in the CMG will cause vibration of the output torque that significantly degrades the tracking performance. On the other hand, the integration of harmonic drives will lead to nonlinear torsional stiffness and introduce additional frictions into the control system. Furthermore, practical tasks such as fast attitude maneuver have posed higher requirement on the rapidity of the anti-disturbance control schemes. To overcome the aforementioned difficulties, we propose a finite-time refined anti-disturbance velocity tracking controller, which consists of a finite-time harmonic disturbance observer (FTHDO) to compensate for the imbalance vibration, an adaptive finite-time extended state observer (AFTESO) to handle the remaining disturbances with unknown rate of change, and a finite-time controller to ensure fast convergence of the tracking error. Both numerical simulation and hardware-in-the-loop (HIL) experimental results show that the proposed scheme outperforms the existing methods in terms of accuracy and rapidity.

Antiunwinding Sliding Mode Control for Rigid Spacecraft Based on Modified Rodrigues Parameters

In this paper, attitude maneuver control without unwinding phenomenon is investigated for rigid spacecraft via modified Rodrigues parameters. A novel switching function is constructed by a hyperbolic sine function to guarantee that the sliding surface contains two stable equilibria. A sliding mode controller is developed to ensure that the sliding surface is reachable. Moreover, a dynamic parameter is designed to guarantee unwinding-free performance of the closed-loop system when the system states outside the sliding surface. In addition, a boundary layer is introduced to deal with the chattering problem. Further, the unwinding-free performance within the boundary layer is guaranteed by imposing some conditions on the parameters of the switching function. Finally, comparison simulations are performed to verify the effectiveness and superiority of the presented control law.

The hybrid attitude maneuver control of the three-axis stable liquid-filled spacecraft with large-deformation appendage

The hybrid attitude maneuver control of the three-axis stable liquid-filled spacecraft with large-deformation appendage

Generating Low-Earth Orbit Satellite Attitude Maneuver Profiles Using Deep Neural Networks

To perform Earth observations, low-Earth orbit (LEO) satellites require attitude maneuvers, which can be classified into two types: maintenance of a target-pointing attitude and maneuvering between target-pointing attitudes. The former depends on the observation target, while the latter has nonlinear characteristics and must consider various conditions. Therefore, generating an optimal reference attitude profile is difficult. Mission performance and satellite antenna position-to-ground communication are also determined by the maneuver profile between the target-pointing attitudes. Generating a reference maneuver profile with small errors before target pointing can enhance the quality of the observation images and increase the maximum possible number of missions and accuracy of ground contact. Therefore, herein we proposed a technique for optimizing the maneuver profile between target-pointing attitudes based on data-based learning. We used a deep neural network based on bidirectional long short-term memory to model the quaternion profiles of LEO satellites. This model was used to predict the maneuvers between target-pointing attitudes. After predicting the attitude profile, it was differentiated to obtain the time and angular acceleration profiles. The optimal maneuver reference profile was obtained by Bayesian-based optimization. To verify the performance of the proposed technique, the results of maneuvers in the 2–68° range were analyzed.

Output Feedback Control for Spacecraft Attitude System with Practical Predefined-Time Stability Based on Anti-Windup Compensator

This paper investigates the problem of output feedback attitude control for rigid spacecraft subject to inertia matrix uncertainty, space disturbance, and input saturation. Firstly, a model transformation is adopted to convert an attitude system with immeasurable angular velocity into a new system. All states of the new converted system are measurable and available for feedback; however, the system contains mismatched uncertainty resulting from the coordinate transformation. Then, an adaptive nonsingular back-stepping control with practical predefined-time convergence is designed. To resolve the problem of input saturation, an anti-windup compensator is developed. It is analytically proved that the spacecraft attitude and angular velocity are practical predefined-time stable, such that the convergence time is a given tunable constant. The simulation results reveal that the proposed control framework provides rapid attitude maneuver and actuator saturation elimination.

Finite-Time Attitude Optimization Maneuver Control for Coupled Spacecraft Under Attitude Measurement Errors and Actuator Faults

The large-angle attitude maneuver trajectory design and tracking control for flexible spacecraft with slosh structure under attitude measurement errors and actuator faults are investigated in this article. Considering that rapid attitude maneuvers can lead to severe flexible vibrations and liquid sloshing, which in turn affects system stability. Therefore, under the requirement of fast maneuvering, the attitude trajectory is optimized to maximize the system’s stability by establishing a new indicator function, while satisfying multiple physical constraints. Furthermore, the novel adaptive multivariable command filtering backstepping fault-tolerant controller is proposed for the coupled spacecraft, to achieve the finite-time attitude tracking to the optimized trajectory, despite the measurement errors and actuator faults. The merits lie in the modified auxiliary signals and virtual commands design, which can compensate the filter errors, caused by the applied filters to deal with the “explosion of terms” of conventional backstepping control, in finite time with better control performance. The numerical simulations are performed to verify the effectiveness of the proposed method.

Optimal attitude maneuvers in the presence of prohibited directions

AbstractWe consider the problem of realigning a space telescope from one observation target to the next in the presence of prohibited directions. More specifically, we want to steer the space telescope – modelled as a gyrostat – from an initial attitude at rest to a final attitude at rest within a fixed time interval. During the motion, the line of sight of the telescope must be kept away from forbidden directions towards bright objects like the sun, moon or earth due to power or thermal requirements. The kinematics of the spacecraft motion are governed by a differential equation on the rotation group SO(3). Treating the angular velocities as control variables, this equation takes the form of a controlled dynamical system. To ensure reorientation maneuvers satisfying these pointing constraints, we introduce a cost functional penalizing proximity of the line of sight of the telescope to any of the forbidden directions. Furthermore, we include penalty terms which provide a smooth motion of the satellite and ensure the execution of a rest‐to‐rest maneuver. The chosen cost functional is minimized over all possible trajectories of the controlled dynamical system between the prescribed initial and target attitudes, which leads to an optimal control problem on SO(3), which is solved by applying a version of Pontryagin's Maximum Principle tailor‐made for optimal control problems on Lie groups. Parametrizing SO(3) in terms of Cardan angles, the solution is formulated as a boundary‐value problem on Euclidean space and hence can be solved numerically by conventional methods. The existence of two first integrals is established and exploited to reduce the computational effort. The applicability of this approach is shown in concrete examples.

Simulated Annealing-Based Heuristic for Multiple Agile Satellites Scheduling Under Cloud Coverage Uncertainty

Agile satellites are the new generation of Earth observation satellites (EOSs) with stronger attitude maneuvering capability. Since optical remote sensing instruments cannot see through the cloud, the cloud coverage brings a significant influence on the satellite observation missions. The scheduling of multiple agile EOSs (AEOSs) is already complicated due to strong satellite maneuverability and onboard satellite energy constraints. Moreover, introducing cloud coverage uncertainty further increases scheduling complexity. Motivated by these challenges, we address the multiple AEOSs scheduling problem under cloud coverage uncertainty where the objective aims to maximize the entire observation profit. A chance constraint programming model is adopted to describe the uncertainty initially, and the observation profit under cloud coverage uncertainty is then calculated via a sample approximation method. To overcome the solving difficulty, an improved simulated annealing (ISA)-based heuristic including a fast insertion strategy is proposed for large-scale observation missions. Finally, extensive experiments are conducted to verify the effectiveness and efficiency of the proposed method. Experimental results show that the ISA-based heuristic outperforms other algorithms for the multiple AEOSs scheduling problem under cloud coverage uncertainty, which validates the proposed algorithm.

Nonlinear attitude maneuvering of a flexible spacecraft for space debris tracking and collision avoidance

This work addresses the problem of attitude maneuvering of a spacecraft equipped with two flexible solar panels, in response to an alert in two distinct operational scenarios: (a) detection of an approaching debris, and (b) collision avoidance of an impacting debris. In scenario (a) the approaching debris is assumed to be off a collision course with the satellite. As a result, the latter can track the debris, to provide supplementary observations (in addition to those given by ground stations) that allow improving the estimation of its trajectory. Instead, in scenario (b) an impact is predicted, therefore the satellite shall perform a collision avoidance maneuver. This is designed with the objective of minimizing propellant consumption, while ensuring a miss distance greater than a specified threshold value. This research considers the overall dynamics by modeling the spacecraft as a multibody structure, with the use of the Kane's method, which simplifies the process of deriving all the governing equations, while identifying their minimum number. Flexibility is introduced using a modal decomposition technique, under the assumption of small amplitude oscillations. In the two scenarios of interest, efficient strategies are introduced to rotate the solar panels, for the purpose of either maximizing their irradiation or minimizing their oscillations. A new nonlinear reduced-attitude control law, which enjoys global stability properties, is proposed, and is shown to improve the performance of attitude maneuvers when only a single axis must be driven toward a desired direction. Actuation is modeled as well, by assuming the use of a pyramidal array of four single-gimbal control momentum gyroscopes. Their steering law includes singularity avoidance, based on the singular value decomposition of the actuation matrix. Numerical simulations demonstrate that the agile attitude maneuvering strategies proposed in this study allow achieving the operational objectives in the two scenarios of interest, with limited elastic oscillations.

Spacecraft attitude maneuver planning with multi–sensor pointing constraints using improved RRT–star algorithm

During the spacecraft attitude maneuvers for the observation of mission targets, multiple instruments may face different pointing constraints, and they are rarely considered by the current attitude maneuver planning methods. In this case, this paper proposes an improved rapidly–exploring random trees star (RRT*) algorithm to obtain discrete quaternion path nodes while satisfying the pointing constraints of multiple sensors. Next, the temporal properties of these path nodes are defined by a trapezoidal path planning method, and bounded constraints on the angular velocity and control moment during the attitude maneuver are introduced. Then, the continuous quaternion curves through the discrete quaternion path nodes are calculated by a segmented cubic spline interpolation function. Finally, the angular velocity and control moment are calculated through the inverse dynamics. The simulation results verify the feasibility of the method. It can satisfy complex constraints and obtain the attitude optimization path, and alleviate the large control moment variations in the start and stop phases of the rest–to–rest maneuver.

Anti-unwinding adaptive super-twisting attitude control for spacecraft

This article addresses the attitude maneuver control without unwinding based on variable gain(i.e. adaptive) super-twisting algorithm and the modified Rodrigues parameters(MRPs). First, a selection logic for the MRP set and its shadow set is designed in order to construct an attitude control system model for rigid spacecraft. Second, a nonlinear sliding function is presented, and on the sliding phase, finite-time convergence and anti-unwinding performance are guaranteed. As a result, a first order sliding mode controller is created to ensure finite-time convergence and anti-unwinding performance on the reaching phase. To alleviate the chattering and improve convergence speed, an adaptive super-twisting algorithm is presented, and the finite-time convergence and anti-unwinding performance are guaranteed on the reaching phase. Furthermore, the adaptive gains are designed based on dual-layer adaptive law, which allows the gains to be increased and decreased as appropriate. It is further demonstrated that when the controller gains are time-varying according to adaptive law, the convergence property and anti-unwinding performance are guaranteed. Numerical simulations are carried out to illustrate the closed-loop system's performance under the proposed control scheme.

Dynamic Modeling and Attitude–Vibration Cooperative Control for a Large-Scale Flexible Spacecraft

Modern spacecraft usually have larger and more flexible appendages whose vibration becomes more and more prominent, and it has a great influence on the precision of spacecraft attitude. Therefore, the cooperative control of attitude maneuvering and structural vibration of the system has become a significant issue in the spacecraft design process. We developed a low-dimensional and high-precision mathematical model for a large-scale flexible spacecraft (LSFS) equipped with a pair of hinged solar arrays in this paper. The analytic global modes are used to obtain the rigid–flexible coupling discrete dynamic model, and the governing equations with multiple DOFs for the system are derived by using the Hamiltonian principle. The rigid–flexible coupled oscillating responses of LSFS under the three-axis attitude-driving torque pulse during the in-orbit attitude maneuvering process are investigated. A study on the flexibility of the hinge was also conducted. Based on the simplified and accurate dynamic model of the system, we can obtain a state-space model for LSFS conveniently, and the cooperative control schemes for rigid motion and flexible oscillation control are designed by using the LQR, PD, and PD + IS algorithms. The simulation results show that three cooperative controllers can realize spacecraft attitude adjustment and synchronously eliminate flexible oscillation successfully. By comparison, the PD + IS controller is simpler so that it is suitable for the real-time attitude–vibration cooperative control of spacecraft.

Simulation of Spacecraft Moon Landing Control in Virtual Environment Complexes

The paper considers the task for simulation of final stage spacecraft landing on the Moon in virtual environment systems. To solve this task, methods and algorithms are proposed for the lunar module motion control with the implementation of fast attitude maneuvers and minimum fuel consumption during its deceleration. The spacecraft control is based on virtual sensors feedback and makes it possible to implement stabilization, reorientation, deceleration, maneuvers, hovering and soft landing of the spacecraft on the Moon. The work involves virtual reality technologies with the implementation of human interaction with a computer-synthesized environment. In this case, to control the spacecraft in manual mode, virtual hands are used that copy the movements of the operator’s hands and act on the elements of virtual controls (joystick, buttons, etc.) inside the spacecraft model. Approbation of methods and algorithms proposed in the paper was carried out in our software package of virtual environment system on the example for landing simulation of virtual model Orel spacecraft in semi-automatic mode. In this software package the spacecraft control in manual mode is implemented by data which transit from Oculus Rift CV1 VR headset and Oculus Touch controllers designed for tracking the operator’s head and hands, as well as displaying synthesized stereopair to his eyes. The simulation of spacecraft landing on the Moon was carried out for stages that begin immediately after the basic deceleration at an altitude of about 2 km and including the free fall of the lunar vehicle, its verticalization, horizontal and vertical deceleration, hovering, and soft landing. The results of approbation showed the adequacy and quality of the solutions proposed in the paper, which can be further used to create simulators designed to train cosmonauts how to control a spacecraft during landing on the Moon.

Satellite Attitude Control Using Double-Gimbal Variable-Speed Control Moment Gyroscope: Single-Loop Control Formulation

A single-loop control architecture is proposed to achieve singularity-free attitude control of satellites actuated using a double-gimbal variable-speed control moment gyroscope (DGVSCMG). Satellite and DGVSCMG combined rotational dynamics is modeled in the framework of geometric mechanics. It is shown that the combined attitude dynamics is in control-affine form when the motor torques of the DGVSCMG are taken as control inputs. A single-loop hybrid robust nonsingular fast finite-time sliding mode control (NFFTSMC) is designed to ensure an unwinding-free attitude maneuver. In the proposed single-loop formulation, the respective motor torques of the DGVSCMG are computed directly using only one NFFTSMC without requiring the steering law, resulting in mathematical singularity-free attitude control. A novel sliding surface is also proposed to ensure that the rotor angular speed converges to its nominal value after the maneuver. This eliminates the need of repeated desaturation of the rotor spin rate, unlike all the steering-law-based control architecture. Unwinding-free global stability of the system using the proposed hybrid NFFTSMC is proved for the sliding and reaching phases. Simulation results are presented to substantiate the claim made, assuming the rigid satellite to be equipped with only one DGVSCMG.