Development of a Simulation System with Integrated In Situ Monitoring Capabilities for the Lubrication State of Rolling Elements in Space Control Moment Gyro Bearings

Distributed Optimization Method for Spacecraft Attitude Control and Vibration Suppression

During an impulsive orbital maneuver, thrust vector misalignment from the center of mass (C.M) generates a large exogenous disturbance torque that results in attitude deviation. This paper aims not to use the reaction control system (RCS) for the spacecraft attitude control. In order to reject the large disturbance very fast, a new control system is proposed. In this method, the large disturbance torque is rejected quickly while the RCS is not employed. The control system is based on one degree of freedom (1DoF) gimbaled-thruster, spin-stabilization, and two control moment gyros (CMG). The nonlinear two-body dynamics of the mentioned spacecraft is formulated. Because RCS is not used, this method is an efficient and implementable method for attitude control of small spacecraft. Numerical simulation shows that thrust vector deviation converges to zero despite disturbance torques. By this method, the disturbance is rejected very fast; thus an accurate orbital velocity change can be obtained. This method can eliminate the initial attitude deviation easily in addition to disturbance rejection. The results show the good performance and superiority of the proposed method compared to some other thrusting maneuver methods.

In the flexible spacecraft with control moment gyroscopes, there are multiple disturbances including not only internal disturbances from actuators and flexible appendages, but also external disturbances from space environment. These disturbances are characterized by a wide frequency range and may degrade attitude control performance to a great extent. In this paper, the lumped disturbance is modeled as a harmonic plus a polynomial model, and an extended harmonic disturbance observer (EHDO) is proposed to estimate the total disturbance. Since the rotor dynamic imbalance disturbance from control moment gyroscopes is described by an internal harmonic model, the lumped disturbance can be estimated precisely via EHDO even with a lower bandwidth. Afterwards, a backstepping-based composite controller is designed to compensate the disturbances by combining the output of EHDO and realize high-precision attitude control for flexible spacecraft with control moment gyroscopes. Simulation results are presented to demonstrate the effectiveness of the proposed method.

Dynamic modelling and observation of micro-vibrations generated by a Single Gimbal Control Moment Gyro

Steering control law for double-gimbal scissored-pair CMG

The control moment gyroscope (CMG), which has the capability to produce large, continuously adjustable torques, is a key attitude control actuator for a spacecraft. The rotor spins at a constant rate about the rotor axis. Any rotation of the gimbal causes a change in the angular momentum of the spacecraft, which produces the gyroscopic moment for attitude control. On September, 2011, 200Nms CMGs were carried in orbit onboard Tiangong target spacecraft of China, which was the first inorbit application of independently developed CMG of China. 200Nms CMGs are designed for the attitude control of middle size spacecraft. Since then, even larger CMGs were under development to meet the demand of the attitude control for even larger spacecraft, such as the space station of China. In this research, three techniques are developed: System redundancy technique, for all electrical components and control electronics; Long life flywheel shafting lubrication technique, with screw cap shaped oil reservoir for 15 years long life lubrication; High stiffness gimbal drive and high accurate control technology, with harmonic reducer, and disturbance observer based terminal sliding mode control. The prototype of 1000Nms CMG was developed, and has successfully passed the qualification level environmental test.

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

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.

This book discusses all spacecraft attitude control-related topics: spacecraft (including attitude measurements, actuator, and disturbance torques), modeling, spacecraft attitude determination and estimation, and spacecraft attitude controls. Unlike other books addressing these topics, this book focuses on quaternion-based methods because of its many merits. The book lays a brief, but necessary background on rotation sequence representations and frequently used reference frames that form the foundation of spacecraft attitude description. It then discusses the fundamentals of attitude determination using vector measurements, various efficient (including very recently developed) attitude determination algorithms, and the instruments and methods of popular vector measurements. With available attitude measurements, attitude control designs for inertial point and nadir pointing are presented in terms of required torques which are independent of actuators in use. Given the required control torques, some actuators are not able to generate the accurate control torques, therefore, spacecraft attitude control design methods with achievable torques for these actuators (for example, magnetic torque bars and control moment gyros) are provided. Some rigorous controllability results are provided. The book also includes attitude control in some special maneuvers, such as orbital-raising, docking and rendezvous, that are normally not discussed in similar books. Almost all design methods are based on state-spaced modern control approaches, such as linear quadratic optimal control, robust pole assignment control, model predictive control, and gain scheduling control. Applications of these methods to spacecraft attitude control problems are provided. Appendices are provided for readers who are not familiar with these topics.

Some Problems of Optimum Three-Dimensional Attitude Control of Spacecraft

The attitude dynamics of a spacecraft with a variable speed control moment gyroscope (VSCMG), in the presence of external torques and internal inputs, is derived using variational principles. A complete dynamics model, that relaxes some of the assumptions made in prior literature on control moment gyroscopes, is obtained. A non-standard VSCMG model, that has an offset between the center of the gimbal axis and the center of the rotor (flywheel) is considered. The dynamics equations show the complex nonlinear coupling between the internal degrees of freedom associated with the VSCMG and the spacecraft base body’s attitude degrees of freedom. Some of this coupling is induced by the non-zero offset between the gimbal axis and the rotor center. This dynamics model is then generalized to include the effects of multiple control moment gyroscopes placed in the base body with non-parallel gimbal axes. It is shown that the dynamical coupling can improve the control authority on the angular momentum of the base body of the spacecraft using changes in the momentum variables of the VSCMG. Numerical simulations confirm the use of these VSCMGs for attitude control for a given de-tumbling maneuver.

To deal with bounded parametric uncertainties and external disturbances existed in a flexible spacecraft with control moment gyroscopes as actuator, an adaptive robust attitude control algorithm is presented in this paper. With the proposed control law, robust stability is achieved, as well as system chattering is weakened. A steering law with modified robust coefficient is also proposed in this paper. Stability of the closed-loop control system is proved via Lyapunov technique. Numerical simulations are included to demonstrate the effectiveness of the proposed approach for the flexible spacecraft attitude maneuver with control moment gyroscopes.

Design and Dynamic Analysis of Micro-vibration Isolator for Single Gimbal Control Moment Gyro

Control moment gyro (CMG) is a critical component for attitude control of large spacecraft. To avoid its accidental failure caused by the performance degradation of high-speed rotating bearings, a metric, which is referred to as the health indicator (HI), should be constructed and monitored. Although the HI construction problem have been widely studied to evaluate the degradation process of machinery, the poor quality of telemetry data makes it impossible to apply existing solutions to the bearings in CMG. Therefore, a novel HI construction method based on energy theory and physics-inspired machine learning is proposed. Generally, the energy balance relationship of system would change with the degradation of its components. Based on this idea, a physics-inspired convolutional neural network model is established to formulate the relationship between monitoring signals and total energy consumed by driving motor. Then a transfer learning strategy is developed to capture the changes of model parameters and perform HI construction. Compared with existing approaches, the proposed method does not rely on vibration signals or high-frequency data, and can extract degradation information obscured by controller and other influences. A real dataset collected from an aerospace CMG is utilized to verify the effectiveness of proposed method. Results show that the constructed HIs have good performance in monotonicity, robustness, and time correlation. The application of constructed HIs for CMG anomaly detection is also discussed.

Control moment gyros (CMGs) offer the advantage of being able to generate large torque, therefore; they are commonly used to provide attitude control in heavy satellites and agile attitude control in small spacecrafts. On the other hand, a CMG possesses singularity, owing to which it becomes impossible to generate a desired torque if a CMG reaches singularity. In this paper, we propose a method of attitude control that considers the avoidance of singularity by nonlinear model predictive control (NMPC).