MIMO active vibration control of magnetically suspended flywheels for satellite IPAC service

Theory and simulation results have demonstrated that four, variable speed flywheels could potentially provide the energy storage and attitude control functions of existing batteries and control moment gyros (CMGs) on a satellite. Past modeling and control algorithms were based on the assumption of rigidity in the flywheel’s bearings and the satellite structure. This dissertation provides simulation results and theory which eliminates this assumption utilizing control algorithms for active vibration control (AVC), flywheel shaft levitation and integrated power transfer and attitude control (IPAC) that are effective even with low stiffness active magnetic bearings (AMB), and flexible satellite appendages. The flywheel AVC and levitation tasks are provided by a multi input multi output (MIMO) control law that enhances stability by reducing the dependence of the forward and backward gyroscopic poles with changes in flywheel speed. The control law is shown to be effective even for (1) Large polar to transverse inertia ratios which increases the stored energy density while causing the poles to become more speed dependent and, (2) Low bandwidth controllers shaped to suppress high frequency noise. These two main tasks could be successfully achieved by MIMO (Gyroscopic) control algorithm, which is unique approach. The vibration control mass (VCM) is designed to reduce the vibrations of flexible appendages of the satellite. During IPAC maneuver, the oscillation of flywheel spin speeds, torque motions and satellite appendages are significantly reduced compared without VCM. Several different properties are demonstrated to obtain optimal VCM. Notch, band-pass and low-pass filters are implemented in the AMB system to reduce and cancel high frequency, dynamic bearing forces and motor torques due to flywheel mass imbalance. The transmitted forces and torques to satellite are considerably decreased in the present of both notch and band-pass filter stages. Successful IPAC simulation results are presented with a 12 [%] of initial attitude error, large polar to transverse inertia ratio (IP / IT), structural flexibility and unbalance mass disturbance. Two variable speed control moment gyros (VSCMGs) are utilized to demonstrate simultaneous attitude control and power transfer instead of using four standard pyramid configurations. Launching weights including payload and costs can be significantly reduced

Optimal guidance and control in space technology

In this thesis, we deal with several optimal guidance and control problems of the spacecrafts arising from the study of lunar exploration. The research is composed of three parts: 1. Optimal guidance for the lunar module soft landing, 2. Spacecraft attitude control system design basing on double gimbal control moment gyroscopes (DGCMGs), and 3. Synchronization motion control for a class of nonlinear system.To achieve a precise pinpoint lunar module soft landing, we first derive a three dimensional dynamics to describe the motion of the module for the powered descent part by introducing three coordinate frames with consideration of the moon rotation. Then, we move on to construct an optimal guidance law to achieve the lunar module soft landing which is treated as a continuously powered descent process with a constraint on the angle of the module between its longitudinal axis and the moon surface. When the module reaches the landing target, the terminal attitude of the module should be within an allowable small deviation from being vertical with reference to lunar surface. The fuel consumption and the terminal time should also be minimized. The optimal descent trajectory of the lunar module is calculated by using the control parameterization technique in conjunction with a time scaling transform. By these two methods, the optimal control problem is approximated by a sequence of optimal parameter selection problems which can be solved by existing gradient-based optimization methods. MISER 3.3, a general purpose optimal control software package, was developed based on these methods. We make use of this optimal control software package to solve our problem. The optimal trajectory tracking problem, where a desired trajectory is to be tracked with the least fuel consumption in the minimum time, is also considered and solved.With the consideration of some unpredicted situations, such as initial point perturbations, we move on to construct a nonlinear optimal feedback control law for the powered deceleration phase of the lunar module soft landing. The motion of the lunar module is described in the three dimensional coordinate system. Based on the nonlinear dynamics of the module, we obtain the form of an optimal closed loop control law, where a feedback gain matrix is involved. It is then shown that this feedback gain matrix satisfies a Riccati-like matrix differential equation. The optimal control problem is first solved as an open loop optimal control problem by using the time scaling transform and the control parameterization method. By virtue of the relationship between the optimal open loop control and the optimal closed loop control along the optimal trajectory, we present a practical method to calculate an approximate optimal feedback gain matrix, without having to solve an optimal control problem involving the complex Riccati-like matrix differential equation coupled with the original system dynamics.To realize the spacecraft large angle attitude maneuvers, we derive an exact general mathematical description of spacecraft attitude motion driven by DGCMGs system. Then, a nonlinear control law is designed based on the second method of Lyapunov and the stability of the attitude control system is established during the design process. A singularity robustness plus null motion steering law is designed to realize the control law. Principle of DGCMGs’ singularity is proved, and the singularity analysis of the orthogonally mounted three DGCMGs system and that of the parallel mounted four DGCMGs system are presented.Finally, we consider a new class of nonlinear optimal tracking and synchronizing control problems subject to control constraints, where the motions of two distinct objects are required to achieve synchronization at the minimum time while achieving the optimal tracking of a reference target. We first provide a rigorous mathematical formulation for this class of optimal control problems. A new result ensuring the synchronization of the two distinct objects is obtained. On this basis, a computational method is developed for constructing an optimal switching control law under which the motions of the two distinct objects will achieve synchronization at the minimum time while achieving the optimal tracking of a reference target. This computational method is developed based on novel applications of the control parameterization method and a time scaling transform. A practical problem arising from the study of the angular velocity tracking and synchronization of two spacecrafts during their formation flight is formulated and solved by the method proposed

Exact steering in control of moment gyroscope systems

Single Gimbal Control Moment Gyroscopes (CMGs) are thought to be efficient actuators for the attitude control of the new generation of small and agile satellites. CMGs belong to a class of actuators known as momentum exchange devices. This thesis presents a detailed formulation of three-axis attitude dynamics and control of a satellite equipped with a cluster of n momentum exchange devices (which include CMGs and reaction wheels).EThOS - Electronic Theses Online ServiceGBUnited Kingdo

A Comparison Study on Control Moment Gyroscope Arrays and Steering Laws

Current reaction wheels and magnetorquers for microsatellite are limited by low slew rate and heavily depends on orbital parameters for coverage area. Control moment gyroscope (CMG) clusters offer an alternative solution for high slew rates and rapid retargeting. Though CMGs are often used in large space missions, their use in microsatellites is limited due to the stringent mass budget. Most literature reports only on pyramid configuration, and there are no definite cross-comparison studies between various CMG clusters and steering laws. In this research, a generic tool in Matlab and Simulink is developed to further understand CMG configurations and steering laws for a microsat mission. Various steering laws necessary for mitigating singularities in CMG clusters are compared in two distinct missions. The simulation results were evaluated based on the pointing accuracy, platform jitter, and pointing stability achieved by the spacecraft for each combination of CMG clusters, and steering laws and trajectories. The simulation results demonstrate that the pyramid cluster is marginally better than the rooftop cluster in pointing accuracy. The comparison of steering laws shows that, counterintuitively, Singularity Robust steering law, which passes through singularities, outperforms both Moore-Penrose and Local Gradient methods for almost all evaluation criteria for the two missions it was tested on. The simulation results would aid systems engineers in designing low-cost actuation systems and corresponding control software, which can increase the data acquisition rate of remote sensing missions

Optimal Attitude Control of Agile Spacecraft Using Combined Reaction Wheel and Control Moment Gyroscope Arrays

This dissertation explores the benefits of combined control moment gyroscope (CMG) and reaction wheel array (RWA) actuation for agile spacecraft. Agile spacecraft are capable of slewing to multiple targets in minimum time. CMGs provide the largest torque capability of current momentum exchange actuation devices but also introduce singularity events in operation. RWAs produce less torque capability than CMGs but can achieve greater pointing accuracy. In this research, a combined RWA and CMG (RWCMG) system is evaluated using analytical simulations and hardware experiments. A closed-loop control scheme is developed which takes advantage of the strengths of each actuator set.The CMGs perform slews for a representative target field. Borrowing from variable-speed CMG theory, a system of switching between CMG and RWA actuation allows the RWA to assume control of the spacecraft when desired pointing tolerance is met for a given target. During collection, the CMG gimbals may travel along null motion trajectories towardpreferred angles to prepare for the next slew. Preferred gimbal angles are pre-computed off-line using optimization techniques or set based on look-up tables. Logic is developed to ensure CMG gimbal angles travel the shortest path to the preferred values. The proportional-integral-derivative, quaternion feedback, and nonlinear Lyapunov-based controllers are assessed for the RWCMG system. Extended and unscented Kalman filter techniques are explored for improved accuracy in analytical simulation. Results of RWCMG hardware experiments show improvements in slew capability, pointing accuracy, and singularity avoidance compared to traditional CMG-only systems

Variable Vector Countermeasure Suit (V2Suit) for Space Habitation and Exploration

The Variable Vector Countermeasure Suit (V2Suit) for Space Habitation and Exploration is a visionary system concept that will revolutionize space missions by providing a platform for integrating sensors and actuators with daily astronaut intravehicular activities to improve human health and performance. The V2Suit uses control moment gyroscopes (CMGs) within a miniaturized module placed on body segments to provide a viscous resistance during movements _ a countermeasure to the sensorimotor and musculoskeletal adaptation performance decrements that manifest themselves while living and working in microgravity and during gravitational transitions during long-duration spaceflight, including post-flight recovery and rehabilitation. Through an integrated design, system initialization, and control systems approach the V2Suit is capable of generating this viscous resistance along an arbitrarily specified direction of down. When movements are made, for example, parallel to that down direction a resistance is applied, and when the movement is perpendicular to that direction no resistance is applied. The V2Suit proposes to be a countermeasure to this spaceflight-related adaptation and de-conditioning and the unique sensorimotor characteristics associated with living and working in 0-G, which are critical for future long-duration space missions. This NIAC Phase II project leveraged the study results from Phase I and focused on detailing several aspects of the V2Suit concept, including a wearable CMG architecture, control steering laws, human-system integration evaluations, developing a brassboard prototype unit as a proof-of-concept, as well as evaluating the concept in the context of future space exploration missions. A human mission to Mars, such as that outlined in the Mars Design Reference Architecture 5.0, provides a framework for determining the concept of operations and requirements for the V2Suit system. Mars DRA 5.0 includes approximately 180 day 0-G transits to- and from- Mars, as well as a 500 day stay on the surface (~3/8-G) (Figure 3). Accordingly, there are four gravitational transitions associated with this mission: 1-G to 0-G (Earth launch), 0-G to 3/8-G (Mars landing), 3/8-G to 0-G (Mars launch), and 0-G to 1-G (Earth landing). This reference mission provided the basis for developing high-level operational requirements to guide the subsequent study and design of the key V2Suit components