Spacecraft Attitude Control System Research Articles

Multiplicative fault concept and fault observer design for reliability study of degrading reaction wheel

The reaction wheel is a typical actuator and a crucial weak link in the spacecraft attitude control system, and much attention has been paid to its reliability problem. Due to limited samples and high cost for reaction wheel life tests, a simulation method by introducing attitude coupling dynamics and multiplicative fault concept is developed to analyze the logic of electric current as a performance indicator and verify its accuracy for reliability modeling. Furthermore, a new and intrinsic performance indicator of multiplicative fault is proposed for more application scenarios of reliability modeling and an adaptive sliding mode observer is designed for fault estimation. An illustrative example shows that the performance indicator of multiplicative fault can be used for various mission scenarios but requires certain persistent excitation, while electric current is the opposite.

Hybrid nonfragile intermediate observer-based T-S fuzzy attitude control for flexible spacecraft with input saturation

This paper investigates a hybrid nonfragile intermediate observer-based T-S fuzzy attitude control strategy subject to multi-source complex disturbances to achieve integrated attitude stabilization and vibration suppression for the flexible spacecraft. A refined T-S fuzzy model for flexible spacecraft is introduced to represent the nonlinear characteristics in large-angle attitude maneuver. Then a hybrid nonfragile observer with coexisting additive and multiplicative perturbations is designed to estimate state information and disturbance simultaneously. Consequently, based on the closed-loop system, the quadratic stability and robust H∞ performance are proved via the Lyapunov stability analysis with corresponding conditions in terms of linear matrix inequalities (LMIs). It is worth mentioning that different from the limiter directly used in much literature, input saturation is disposed as a disturbance during the proof to satisfy the stability requirements. Finally, numerical simulations of a flexible spacecraft attitude control system are performed which demonstrate the effectiveness and superiority of the proposed control strategy.

Constrained multi‐observer‐based fault‐tolerant disturbance‐rejection control for rigid spacecraft

AbstractThis article develops a multi‐observer‐based fault‐tolerant disturbance‐rejection control strategy to solve the attitude stabilization problem of spacecraft subject to multisource complex disturbances, for example, external disturbance, measurement error, actuator fault, input constraint. First, two intermediate variables are introduced for multi‐observer design, so that the synergistic estimations of attitude information, actuator fault and external disturbance are obtained simultaneously. Then, a fault‐tolerant disturbance‐rejection control strategy is proposed based on the estimations, and an augmented closed‐loop system is derived. Afterwards, Lyapunov stability analysis is performed to prove the quadratic stability and robust performance, and corresponding conditions in terms of linear matrix inequalities (LMIs) is proved, where the input constraint is satisfied as well. Finally, numerical simulations of a spacecraft attitude control system are performed which demonstrate the effectiveness and superiority of the proposed multi‐observer‐based control strategy.

Spacecraft attitude control and vibration suppression using magnetically suspended control & sensitive gyroscope and radial basis function network adaptive sliding mode control

For the sake of improving the agility and super-static performance of spacecraft under external disturbances, this paper puts forward an idea for using pyramid configuration of the Magnetically Suspended Control & Sensitive Gyroscope to realize dual control of attitude maneuver and vibration suppression of spacecraft. The RBF neural network is used to reduce the chattering in sliding mode control through adaptively learning the non-linear part of the system. A band-pass filter is designed to divide the attitude deviation signal fed back by the sensors into high-frequency and low-frequency parts. The control laws of gyro gimbal and magnetic suspension rotor are designed, respectively, to realize the attitude maneuver and vibration suppression simultaneously. The adaptive laws of parameters avoid the saturation of rotor deflection angle while speeding up the attitude maneuver process of spacecraft. Compared with the traditional sliding mode control methods, the proposed one not only improves the accuracy and speed of attitude control but also improves the robustness of the spacecraft attitude control and vibration suppression system. Semi-physical simulation results show the effectiveness and superiority of the proposed method.

Gimbal Angular Velocity Feedforward Method for Magnetically Suspended Control Moment Gyro with Hybrid Magnetic Bearing#

Control moment gyro(CMG) has the advantages of high precision and large output moment. It is the key actuator of agile spacecraft attitude control system. Magnetically suspended CMG(MSCMG) supports high-speed rotor through magnetic bearing, which has the advantages of no friction, no wear, high precision and long life. Hybrid magnetic bearing(HMB) generates bias magnetic field by replacing bias current with permanent magnet, which greatly reduces the power consumption. However, when MSCMG output moment, moving-gimbal effect makes the control current of HMB increase sharply, which result in the rapid increase of power consumption. To solve this problem, gimbal angular velocity feedforward(GAVF) method is proposed in this paper. Firstly, the model of HMB-rotor system containing moving-gimbal effect is established. Secondly, GAVF method is introduced to make the rotor deflect by designing a feedforward matrix with gimbal angular velocity and stiffness of HMB which is determined with rotor deflection angle and rotation speed. Thus, the gyroscopic moment is mainly offset by the bias magnetic field of HMB. Finally, an adaptive compensation method based on GAVF is proposed to maintain the system performance under parameter perturbation. By the above, the power consumption of control current during moment output process is reduced. Simulation and experiment results show the effectiveness of the proposed method.

Control Moment Gyroscopes for Spacecraft Attitude Control Systems: History of Development

Control Moment Gyroscopes for Spacecraft Attitude Control Systems: History of Development

IAR-STSCKF-Based Fault Diagnosis and Reconstruction for Spacecraft Attitude Control Systems

IAR-STSCKF-Based Fault Diagnosis and Reconstruction for Spacecraft Attitude Control Systems

Dynamic modeling and open-loop analysis of a control moment gyroscope considering the influence of a flexible vibration isolator

The microvibrations that affect the output performance of control moment gyroscopes (CMGs) are important in spacecraft attitude control systems. To reduce the influence of vibrations, researchers have done much work to isolate the vibrations between the CMG and the satellite. However, the deformation of the flexibility isolator affects the direction of the angular momentum, which affects the dynamic performance of the CMG inverse. In this research, using the Lagrange’s equation, the dynamic model of the CMG-isolator coupling system with the exciting force of the dynamic unbalance torque, the CMG’s gyro torque and the gimbal motor torque is developed. Based on this model, the coupling relationship between the gimbal rotation and the isolator deformation is established. For this CMG-isolator coupling system, we conduct the simulations to study the influence of the flywheel speed and the isolator stiffness on the CMG’s output performance and its gimbal’s rotation. Finally, the experiments on a CMG prototype are conducted, which verify the dynamic model and the simulation results of the CMG-isolator coupling system.

Robust finite-time control design for attitude stabilization of spacecraft under measurement uncertainties

This paper investigates the finite-time attitude stabilization of spacecraft under multiple uncertainties. Besides inertia uncertainties, external disturbances, and actuator faults, the measurement uncertainties are particularly considered. When including the measurement uncertainties, the spacecraft attitude control system becomes a mismatching system, which brings a great challenge to the control design. To resolve this problem, a novel robust finite-time attitude control approach is proposed by using a dual-disturbance-observer-based structure. First, two finite-time disturbance observers are developed to estimate the mismatched and matched lumped disturbances in the spacecraft attitude kinematics and dynamics, respectively. Then, the robust finite-time attitude controller is synthesized by incorporating the two finite-time disturbance observers into the adding a power integrator technique. It is strictly proved that the proposed robust finite-time attitude controller can ensure the actual attitude and angular velocity stabilize to the small neighborhoods around the origin in finite time even subject to multiple uncertainties. Benefiting from the feedforward disturbance compensation, the proposed controller is not only robust against inertia uncertainties, external disturbances, and actuator faults, but also insensitive to measurement uncertainties. Lastly, the effectiveness and advantages of the proposed control approach are validated through simulations and comparisons.

Lyapunov differential equations and inequalities for stability and stabilization of linear time-varying systems

This paper studies exponential stability and stabilization of linear time-varying systems without assuming that the coefficient matrix is bounded, which is generally necessary in the existing literature. First, by introducing the notions of uniformly exponentially stable and uniformly exponentially expanding functions, some Lyapunov differential inequalities based characterizations for exponential stability of linear time-varying systems are established. Second, it is shown that the linear time-varying system is both uniformly exponentially stable and uniformly exponentially expanding if and only if the corresponding Lyapunov differential equation has a unique positive definite solution. Third, some degenerated Lyapunov differential equations, where the Q matrix is only positive semi-definite, are also re-examined. Finally, a weighted controllability Gramian, which contains a weighting time-varying function that can be properly designed to reveal a trade-off between the convergence rate/regulation time of the closed-loop system and the control effort, based stabilization approach is established to ensure that the closed-loop system is both uniformly exponentially stable and uniformly exponentially expanding. The effectiveness of the proposed approach is illustrated by the design of the spacecraft attitude control system with magnetic actuation.

Rapid search method for a spacecraft attitude maneuver path with multiple constraints

The rapid attitude maneuver capability of spacecraft is required in many space missions. Due to the complex attitude constraints and limited satellite resources, the rapid on-orbit attitude maneuver is very challenging for spacecraft attitude control systems. To address this problem, a rapid attitude path search method named Constraint Path Mapping (CPM) is proposed to quickly obtain a safe attitude path. In the CPM method, the attitude constraint safety-mapping strategy is designed to map the attitude constraint space to the attitude safety space, allowing for efficient planning and improved processing of complex attitude constraints. For resource use, this paper adopts the Euler forward search and node safety search to avoid algorithm redundancy caused by the repeated processing of the constraints. Finally, the numerical simulation is performed to validate the feasibility and effectiveness of the proposed method. In particular, this method is expected to be tested on a reconnaissance satellite launched by the Shanghai Institute of Satellite Engineering in March 2021.

Modeling for the Electromagnetic Dynamic Distortion Effect of the Induction-Based Electrodynamic Suspension Reaction Spheres

The electrodynamic suspension reaction sphere (EDSRS) is regarded as a novel actuator for the spacecraft attitude control system. The omnidirectional rotation of the EDSRS spherical rotor brings about the dynamic distortion of the electromagnetic force and torque characteristics. With the mechanical structure of the EDSRS simplified and equivalent to multi-degrees of freedom (DOF) single-sided linear induction motor (LIM), this article first proposes a 3-D LIM analytical electromagnetic force model by employing the electromagnetic physical quantities in the 2-D Fourier series expansion form. Then, an analytical electromagnetic force and torque model considering the electromagnetic dynamic distortion effect (EDDE) of the EDSRS is established based on the previous LIM force model according to the motion decomposition principle. Finally, the proposed analytical EDDE model is experimentally validated on the developed laboratory prototype.

ICESat-2 Early Mission Synopsis and Observatory Performance.

The Advanced Topographic Laser Altimetry System (ATLAS) onboard the NASA Ice, Cloud, and land Elevation Satellite‐2 (ICESat‐2) is the newest Earth observing satellite for global elevation studies. The primary objectives for ICESat‐2 follow the objectives of its predecessor, ICESat and also focus on providing cryospheric measurements to determine ice sheet mass balance, and monitor both sea ice thickness and extent. However, the global observations support secondary science objectives as well such as biomass estimation, inland water elevation, sea state height and aerosol concentrations. Since launch of ICESat‐2, ATLAS has collected more than a trillion measurements. This study provides a mission overview, a description of the operational components that enable the altimeter products for science, on‐orbit observatory performance, and assessment of the spacecraft attitude control systems that enable repeat measurements to within 10 m and pointing control within ±45 m. These metrics should be considered for ground‐based validation campaigns or science investigations.

Constructive schemes to spacecraft attitude control with low communication frequency using sampled-data and encryption approaches

PurposeRecent spacecraft attitude control systems tend to use wireless communication for cost-saving and distributed mission purposes while encountering limited communication resources and data exposure issues. This paper aims to study the attitude control problem with low communication frequency under the sampled-data.Design/methodology/approachThe authors propose constructive control system structures based on quantization and event-triggered methods for intra-spacecraft and multi-spacecraft systems, and they also provide potential solutions to shield the control system's data security. The proposed control architectures can effectively save communication resources for both intra-spacecraft and multi-spacecraft systems.FindingsThe proposed control architectures no longer require sensors with trigger-ing mechanism and can achieve distributed control schemes. This paper also provides proposals of employing the public key encryption to secure the data in control-loop, which is transmitted by the event-triggered control mechanism.Practical implicationsSpacecraft attempts to use wireless communication, yet the attitude control system does not follow up promptly to accommodate these variations. Compared with existing approaches, the proposed control structures can save communication resources of control-loop in multi-sections effectively, and systematically, by rationally configuring the location of quantization and event-triggered mechanisms.Originality/valueThis paper presents several new control schemes and a necessary condition for the employment of encryption algorithms for control systems based on event-based communication.

Fault-Tolerant Attitude Control for Rigid Spacecraft Without Angular Velocity Measurements

In this paper, a fault-tolerant control scheme is proposed for the rigid spacecraft attitude control system subject to external disturbances, multiple system uncertainties, and actuator faults. The angular velocity measurement is unavailable, which increases the complexity of the problem. An observer is first designed based on the super-twisting sliding mode method, which can provide accurate estimates of the angular velocity in finite time. Then, an adaptive fault-tolerant controller is proposed based on neural networks using the information from the observer. It is shown that the attitude orientations converge to the desired values exponentially. Finally, a simulation example is utilized to verify the effectiveness of the proposed scheme.

ОБЕСПЕЧЕНИЕ ЖИВУЧЕСТИ СИСТЕМЫ УПРАВЛЕНИЯ КОСМИЧЕСКИМ АППАРАТОМ ПРИ КРИТИЧЕСКИХ ОТКАЗАХ РЕАКТИВНЫХ МАХОВИКОВ

A method for ensuring the survivability of the spacecraft attitude control system with a minimally redundant cluster of flywheels by General Electric scheme and a magnetic drive in the event of the flywheels' failures is presented. The results of computer simulation were obtained and it was found that in case of failure of any two flywheels, the Earth survey satellite retains the ability to scanning observation given targets.

H∞ Control Design for a Rigid-Flexible Spacecraft under a Fuel Slosh Influence

The spacecraft Attitude Control System (ACS) performance and robustness depend on the interactioneffects between the fuel slosh motion, the panel's flexible motion, and the spacecraft rigid motion, mainly duringtranslational and/or rotational maneuvers. In regards to satellite pointing accuracy flexibility and fuel, slosh is thetwo most important effects that should be considered in the satellite ACS design since their interactions can damage the ACS performance and robustness. Once, the lowest vibration frequencies, normally of the sloshing modeare about six times less than of the ACS bandwidth. Therefore, there is a strong possibility that this mode can destabilize the ACS pointing accuracy. This phenomenon is called spillover because the control effort spills over outside the control bandwidth. As a result, the designer needs to explore the limits between the conflicting requirements of performance, that is, increase of the bandwidth without introduction noise in the ACS keeping the systemrobustness to parameters variation. In this paper, one applies the H infinity control method which can deal withthese two design requirements (performance and robustness) considering the controller error pointing that may belimited by the minimum time necessary to suppress disturbances that affect the satellite attitude acquisition. Theequations of motions are obtained considering the Lagrange method for small flexible deformations and a mechanical model of liquid sloshing which allows modeling and investigating the longitudinal dynamic characteristics of apartially filled liquid tank during a pitch maneuver, satisfying performance and robustness requirements.

Stable Control of Nutation and Precession for the Radial Four-Degree-of-Freedom AMB-Rotor System Considering Strong Gyroscopic Effects

The stable control of the active magnetic bearing rotor (AMB-rotor) system is the key technology for a magnetically suspended control moment gyroscope, which is used in the spacecraft attitude control system. The nutation and precession instability caused by strong gyroscopic effects is the principal issue of the AMB-rotor system. In order to solve the gyroscopic effects problem, in this article, a state-space model of the radial four-degree-of-freedom AMB-rotor system is established, and the strong gyroscopic effects term varying with the rotor speed is regarded as a bounded perturbation. The regional pole assignment method of an uncertain system is used to design a robust controller to maintain the system stability over the full rotor speed range. However, the controller designed by the regional pole assignment method does not consider the influence of closed-loop stiffness of the magnetic bearing, which will lead to the start-up oscillation and structural mode self-excited oscillation problems. Therefore, on the basis of the regional pole assignment, the system stiffness is further constrained and a robust controller is solved. Experiments show that the designed controller can not only deal with the start-up oscillation and the structural modal vibration problems, but also can achieve the stable control of the nutation and precession.

Assessing debris strikes in spacecraft telemetry: Development and comparison of various techniques

Debris strikes on operational spacecraft are becoming more common due to increasing numbers of space objects. Sample return missions indicate hundreds of minor strikes, but rigorous analysis is often only performed when a strike causes an anomaly in spacecraft performance. Developing techniques to identify and assess minor strikes that do not immediately cause anomalous behavior can help to validate models for debris populations, perform risk assessments, and aid in the attribution of future anomalies. This study introduces debris strikes to a spacecraft dynamics simulation and assesses the effect on spacecraft telemetry. Various signal processing and change detection techniques are used to identify strikes in noisy telemetry and estimate strike parameters. Matched filter wavelets are developed to identify the effects on state telemetry, where errors are autonomously corrected by the spacecraft attitude control system. A bank of matched filters is used to estimate the parameters of the strike based on a priori knowledge of the spacecraft's response characteristics. A sequential probability ratio test is used to highlight abrupt changes in the spacecraft's angular momentum. Monte-Carlo analyses are conducted to characterize the performance of these algorithms. The results of the various techniques are compared in terms of correctly identifying the debris strikes and accurately estimating the strike parameters. Developing the capability to catalog and characterize minor debris strikes allows any spacecraft to be used as an in situ debris sensor.

A method of using fuzzy hypergraphs to evaluate structural and technological survivability of attitude control systems for unmanned spacecraft

Successful completion of a mission by an unmanned spacecraft (USC), both under nominal operational conditions both under examined contingencies and unexamined off-nominal situations is possible through designing survivability into the USC onboard system (OS). An analysis of current methods for evaluating USC OS survivability during their configuration management and reconfiguration under conditions of examined in-flight contingencies widely used in the design and development of the said USC has shown that these methods are not acceptable for evaluating the USC OS survivability in case of unexamined off-nominal situations in flight. This calls for development of conceptually novel methodological and procedural framework for evaluating structural survivability of USC OS configurations that take into account the level of participation of functional elements (FE) and OS subsystems in the USC control operations under various scenarios of the mission plan implementation. The paper proposes an original approach to evaluating the structural and technological survivability of the USC OS based on a fuzzy hypergraph formal representation of the operations to control the USC attitude, where the edges of the hypergraph connect the FE and OS subsystems that support the implementation of this or that specific control process. The paper also shows how one could use for the quantitative evaluation of the structural and technological survivability of a specific USC OS configuration the results of differentiation of a fuzzy hypergraph that could be visualized as a fuzzy hypergraph of technological independence of OS FE. Such an approach makes it possible to analyze the effects of FE on OS, identify the most critical elements, which have the lowest technological independence under mission plan implementation conditions, which could be used for providing a rationale for the required level of structural and functional redundancy of USC elements and subsystems introduced during various phases in its life cycle. Keywords: unmanned spacecraft, onboard systems survivability, fuzzy hypergraph derivative.