Control Moment Gyros Cluster Research Articles

Deep reinforcement learning-based attitude control for spacecraft using control moment gyros

This paper addresses the development of an attitude control system that steers control moment gyros (CMGs) based on deep reinforcement learning (DRL) for agile spacecraft. The proposed DRL-based attitude control system learns CMG steering strategies to achieve the desired attitude, thus potentially bypassing the singularity issues inherent in the CMG cluster. In particular, it is designed in two-phases to apply the DRL technique efficiently. In the first phase, the attitude control is performed based on DRL up to a certain tolerance, after which it switches to conventional control and steering law for stabilization in the second phase. The rapid pointing capability of the proposed DRL-based attitude control system is demonstrated for an agile spacecraft equipped with pyramid-type single gimbal control moment gyros. Additionally, in realistic scenarios of pointing multiple targets on the ground, the momentum vector recovery that the CMG system needs to consider is also briefly discussed.

Fault detection and isolation of control moment gyros for satellite attitude control subsystem

Abstract Control moment gyros, as one of the most commonly used actuators onboard satellites, are prone to faults and failures. The ability to detect faults, isolate their location, and identify their severity can enhance mission success rate and reduce maintenance and damage costs extensively. Therefore, in this paper, a model-based fault detection, isolation, and identification scheme is proposed and evaluated. Firstly, an adaptive threshold fault detection algorithm is proposed, using Unscented Kalman filters in conjunction with residual and innovation sequences. Secondly, a fault isolation and identification algorithm is proposed using a binary grid search method to adapt joint estimation Kalman filter’s covariance matrix once a fault is detected. Extensive Monte Carlo simulations are conducted to evaluate the performance of the proposed schemes for a 3-axis stabilized satellite with a four-single-gimbal control moment gyro cluster. Results show superior performance of the proposed methods with faster tracking and more accuracy.

Envelope oriented singularity robust steering law of control moment gyros for spacecraft attitude maneuver

The problem of steering pyramid control moment gyro (CMG) cluster for fast spacecraft attitude maneuver along eigenaxis is investigated. A novel steering law is proposed to continuously attempt to reduce the difference between the current gimbal angle and the desired one corresponding to the angular momentum envelop of the CMG cluster. The proposed steering law can be decomposed into two parts: the first one is a singularity robust term to keep maneuverability and produce control torque, and the other is a null motion term to rearrange the gimbal angles toward momentum envelope. By involving this steering law, it is expected to possess both rapid angular momentum exchange and singularity avoidance ability. In addition, by introducing a new limit vector on attitude error, classical cascade-saturation control algorithm is revised to guarantee spacecraft eigenaxis rotation. Both open-loop steering law test and closed-loop attitude maneuver simulations are performed to evaluate the efficacy of the proposed methods.

Steering law of control moment gyros using artificial potential function approach

A Control Moment Gyro (CMG) cluster for an agile spacecraft usually suffers from a singularity problem where it does not meet a desired torque. To solve the problem, a proposed steering law consists of SR (singularity robustness) inverse and SR null motion to compensate each other. The SR null motion solves the singularity problem that the SR inversestays in the singularities. Especially, an APF (artificial potential function) is included in the SR null motion to avoid internal singularities and to reach external singularity simultaneously. The APF newly considers the distance between the current gimbal angle and the gimbal angle corresponding to all singularities. All the combinations of singular gimbal angles are regarded as a target and obstacles of the potential function. After a spacecraft maneuver, an initial gimbal angle is regarded as the target to secure the repeatability of next maneuver. When avoiding or escaping the obstacles corresponding to the unwanted singular gimbal angles, the proposed steering law rapidly changes the gimbal angle by taking torque error. To overcome the local minima problem which is generally a serious drawback in using the potential function, an algorithm for changing parameters of the potential function are proposed in the steering law. To avoid the internal singularity, moreover, the present study introduces a configuration change of the CMG cluster. The proposed steering law is mathematicallyanalyzed and verified through numerical simulations. It shows that the CMG cluster can avoid an internal singularity rapidly, resulting in faster maneuver time of the spacecraft.

Gimbal-Angle Vectors of the Nonredundant CMG Cluster

This paper deals with the method using the preferred gimbal angles of a control moment gyro (CMG) cluster for controlling spacecraft attitude. To apply the method to the nonredundant CMG cluster, analytical gimbal-angle solutions for the zero angular momentum state are derived, and the gimbal-angle vectors for the nonzero angular momentum states are studied by a numerical method. It will be shown that the number of the gimbal-angle vectors is determined from the given skew angle and the angular momentum state of the CMG cluster. Through numerical examples, it is shown that the method using the preferred gimbal-angle is an efficient approach to avoid internal singularities for the nonredundant CMG cluster.

Coupled dynamic analysis of a single gimbal control moment gyro cluster integrated with an isolation system

Control moment gyros (CMGs) are widely used as actuators for attitude control in spacecraft. However, micro-vibrations produced by CMGs will degrade the pointing performance of high-sensitivity instruments on-board the spacecraft. This paper addresses dynamic modelling and performs an analysis on the micro-vibration isolation for a single gimbal CMG (SGCMG) cluster. First, an analytical model was developed to describe both the coupled SGCMG cluster and the multi-axis isolation system that can express the dynamic outputs. This analytical model accurately reflects the mass and inertia properties, the gyroscopic effects and flexible modes of the coupled system, which can be generalized for isolation applications of SGCMG clusters. Second, the analytical model was validated using MSC.NASTRAN software based on the finite element technique. The dynamic characteristics of the coupled system are affected by the mass distribution and the gyroscopic effects of the SGCMGs. The gyroscopic effects produced by the rotary flywheel will stiffen or soften several of the structural modes of the coupled system. In addition, the gyroscopic effect of each SGCMG can interact with or counteract that of others, which induce vibration modes coupled together. Finally, the performance of the passive isolation was analysed. It was demonstrated that the gyroscopic effects should be considered in isolation studies on SGCMG clusters; otherwise, the isolation performance will be underestimated if they are ignored. • We developed an analytical model for vibration isolation of SGCMG cluster. • The model describes the coupled dynamics of the SGCMG cluster and isolation system. • The model is universal for SGCMG cluster for some kind of mounting configuration. • Dynamic characteristics and isolation performance of the isolation system are influenced by the gyroscopic effects. • The passive isolation filters much of the disturbance and will not affect attitude control.

An improved installation for control moment gyros and its applications on reconfiguration and singular escape

A cluster of control moment gyros (CMGs) is widely used as the actuator of attitude control for modern satellite. However, to use CMGs will induce some accessional problems, like the singularity and the failure of them. In this paper, a rotator for the support bracket of each control moment gyro (CMG) is adopted and its application on reconfiguration and singular escape is discussed. The first step puts forward a novel reconfiguration scheme for the CMGs and applied on a satellite, which can adjust the installation angles of the remaining CMGs and make the CMGs remain axial symmetry. Then the characteristic of the output torque is studied in the process of reconfiguration of the CMGs. Based on this characteristic, a special steering law of the CMGs is designed to counteract the disturbance from the reconfiguration. The feasibility of singularity avoidance via the reconfiguration is also discussed by calculating the derivative of the singularity measurement to the installation angle. A singularity avoidance steering law is proposed based on the above calculation.

인공위성을 위한 제어모멘트자이로의 설계시 고려요소 및 배치형상 개선방안

CMG는 인공위성의 자세제어에 필수적인 구동기로 각광받고 있다. 본 논문에서는 CMG의 특이점 회피를 위해 제한된 각운동량 범위(2H)내에서 사용할 경우, CMG의 설계시 고려할 요소를 확인해본다. 아울러 3개의 CMG가 사용될 경우에 대해 필요한 분석과 해석을 수행하였다. 또한 4개의 CMG가 사용될 경우, CMG의 특이점 회피를 위해 기존에 사용하였던 설치방식과 다른 개선된 설치방식을 제안함으로써 위성 운용시 자세기동 빈도가 많은 롤 또는 피치축의 한쪽 축방향 회전에 대한 기동 성능향상 방안을 확인하였다. Nowadays, CMG(Control Moment Gyros) becomes one of the essential actuators for satellite attitude control. The method to define the key requirements of CMG is suggested to avoid CMG's singularity problem for the limited envelope of angular momentum of 2H. Furthermore, the analysis and simulation are carried out to provide a necessary guideline when three CMGs are used for spacecraft control purpose. An improved configuration of redundant four CMG cluster, slightly different from the conventional configuration, is proposed not only to avoid the CMG singularity problem, but to improve agility about roll or pitch-axis.

Attitude Dynamics/Control of a Dual-BodySpacecraft with Variable-Speed Control Moment Gyros

The dynamics equations of a spacecraft consisting of two bodies mutually rotating around a common gimbal axis are derived by the use of the Newton‐Euler approach. One of the bodies contains a cluster of single-gimbal variable-speed control moment gyros. The equations include all of the inertia terms and are written in a general form, valid for any cluster configurations and any number of actuators in the cluster. A guidance algorithm has been developed under the assumtion that the two bodies of the spacecraft are optically coupled telescopes that relay laser signals. The reference maneuver is found by the imposition of the connectivity between the source and the target on the ground. A new nonlinear control law is designed for the spacecraft attitude and joint rotation by the use of Lyapunov’s direct method. An acceleration-based steering law is used for the variable-speed control moment gyros. The analytical results are tested by numerical simulations conducted for both regulation and tracking cases. I. Introduction T HE dynamics and control of multibody spacecraft are a challenging problem because of the complexity of the dynamics equations and the time-varying inertia of the system. The problem becomes even more interesting when gimbaled momentum exchange devices are considered to control attitude. Control moment gyros (CMGs) are unique among attitude control actuators because they can provide high output torque without using expendable fuels and can provide a level of precision and continuity unachievable with jet thrusters. Indeed, CMGs have been used for decades on space stations and on military spacecraft when fast slewing capability and high pointing accuracy were required. The use of CMGs is also currently under consideration for several future civil spacecraft requiring high agility (as in Refs. 1 and 2). A main drawback to the use of CMGs is the presence of singular gimbal-angle configurations at which the CMG cluster is unable to produce the required torque, or, in some cases, any torque at all. 3−5 Many previous studies have considered the problem of the dynamics and control of spacecraft by the use of single-gimbal CMGs.

Singularity Analysis of Variable Speed Control Moment Gyros

Single-gimbal control moment gyros (CMGs) have many advantages over other actuators for attitude control of spacecraft. For instance, they act as torque amplifiers and, thus, are suitable for slew maneuvers. However, their use as torque actuators is hindered by the presence of singularities, that, when encountered, do not allow a CMG cluster to generate torques about arbitrary directions. One method to overcome this drawback is to use variable-speed single-gimbal control moment gyros (VSCMGs). Whereas the wheel speed of a conventional CMG is constant, VSCMGs are allowed to have variable wheel speed. Therefore, VSCMGs have extra degrees of freedom that can be used to achieve additional objectives, such as singularity avoidance and/or power tracking, as well as attitude control. The singularity problem of a VSCMGs cluster is studied in detail for the cases of attitude tracking, with and without a power tracking requirement. A null motion method to avoid singularities is presented, and a criterion is developed to determine the momentum region over which this method will successfully avoid singularities. This criterion can be used to size the wheels and develop appropriate momentum damping strategies tailored to the specific mission requirements. I. Introduction A CONTROL moment gyro (CMG) is a device used as an actuator for attitude control of spacecraft. It generates torques through angular momentum transfer to and from the main spacecraft body. This is achieved by changing the direction of the angular momentum vector of a gimballed flywheel. Because a CMG operates in a continuous manner, contrary to the gas jet’s on/off operation, it can achieve precise attitude control. Moreover, as with other momentum exchange devices, for example, reaction wheels, it does not consume any propellant, thus prolonging the operational life of the spacecraft. Single-gimbal CMGs essentially act as torque amplifiers. 1 This torque amplification property makes them especially advantageous as attitude control actuators for large space spacecraft and space structures, for example, a space station. In fact, single-gimbal and double-gimbal CMGs have been used for attitude control of the Skylab, the MIR and the International Space Station (ISS). An obstacle when using a CMG system in practice is the existence of singular gimbal angle states for which the CMGs cannot generate a torque along arbitrary directions. At each singular state, all admissible torque directions lie on a two-dimensional surface in the three-dimensional angular momentum space; therefore, the CMG system cannot generate a torque normal to this surface. The CMG singularities can be classified into two categories: 1) external or saturated singularities in which the total angular momentum sum of the CMGs lies on the maximum momentum envelope and 2) internal singularities in which the total momentum lies inside this envelope. The external singularities can be easily anticipated from the given CMG configuration and mission profile; therefore, they can be taken into account at the design step. A properly designed momentum management scheme can also relieve the external singularity problem. The internal CMG singularities, on the other hand, are in general difficult to anticipate. Avoiding such internal singularities has, thus, been a long-standing problem in the CMG