Optimal Attitude Control Management For A Cubesat

CubeSats have become popular among universities, research organizations, and government agencies due to their low cost, small size, and light weight. Their standardized configurations further reduce the development time and ensure more frequent launch opportunities. Early cubesat missions focused on hardware validation and simple communication missions, with little requirement for pointing accuracy. Most of these used magnetic torque rods or coils for attitude stabilization. However, the intrinsic problems associated with magnetic torque systems, such as the lack of three-axis control and low pointing accuracy, make them unsuitable for more advanced missions such as detailed imaging and on-orbit inspection. Three-axis control in a cubesat can be achieved by combining magnetic torque coils with other devices such as thrusters, but the lifetime is limited by the fuel source onboard. To maximize the mission lifetime, a fast attitude control management algorithm that could optimally manage the usage of the magnetic and thruster torques is desirable. Therefore, a recently developed method, the B-Spline-augmented virtual motion camouflage, is presented in this defense to solve the problem. This approach provides results which are very close to those obtained through other popular nonlinear constrained optimal control methods with a significantly reduced computational time. Simulation results are presented to validate the capabilities of the method in this application.

Roger D. Werking Head, Attitude Determination and Control Section National Aeronautics and Space Administration/ Goddard Space Flight Center Extensiye work has been done for many years in the areas of attitude determination, attitude prediction, and attitude control. During this time, it has been difficult to obtain reference material that provided a comprehensive overview of attitude support activities. This lack of reference material has made it difficult for those not intimately involved in attitude functions to become acquainted with the ideas and activities which are essential to understanding the various aspects of spacecraft attitude support. As a result, I felt the need for a document which could be used by a variety of persons to obtain an understanding of the work which has been done in support of spacecraft attitude objectives. It is believed that this book, prepared by the Computer Sciences Corporation under the able direction of Dr. James Wertz, provides this type of reference. This book can serve as a reference for individuals involved in mission planning, attitude determination, and attitude dynamics; an introductory textbook for stu dents and professionals starting in this field; an information source for experimen ters or others involved in spacecraft-related work who need information on spacecraft orientation and how it is determined, but who have neither the time nor the resources to pursue the varied literature on this subject; and a tool for encouraging those who could expand this discipline to do so, because much remains to be done to satisfy future needs.

Due to the volume and power limitations of a small satellite, careful consideration must be taken while designing an attitude control system for 3-axis stabilization. Placing redundancy in the system proves difficult and utilizing power hungry, high accuracy, active actuators is not a viable option. Thus, it is customary to find dependable, passive actuators used in conjunction with small scale active control components. This document describes the application of Elastic Memory Composite materials in the construction of a flexible spacecraft appendage, such as a gravity gradient boom. Assumed modes methods are used with Finite Element Modeling information to obtain the equations of motion for the system while assuming free-free boundary conditions. A discussion is provided to illustrate how cantilever mode shapes are not always the best assumption when modeling small flexible spacecraft. A key point of interest is first resonant modes may be needed in the system design plant in spite of these modes being greater than one order of magnitude in frequency when compared to the crossover frequency of the controller. LQG/LTR optimal control techniques are implemented to compute attitude control gains while controller robustness considerations determine appropriate reduced order controllers and which flexible modes to include in the design model. Key satellite designer concerns in the areas of computer processor sizing, material uncertainty impacts on the system model, and system performance variations resulting from appendage length modifications are addressed.

"A common goal of satellite control systems is to reduce the time required to change a spacecraft's attitude, which maximizes its mission capability. Time- optimal attitude control algorithms increase the agility of satellites such as imaging spacecraft, thus allowing a greater frequency of image collection. Eigenaxis based maneuvering, though common in industry and academia, fails to produce the minimum-time solution for actual satellites. Solving the optimal control problem is often challenging and requires evaluating multiple maneuver paths to ensure the shortest path is found for each spacecraft configuration. One of the primary difficulties in predicting optimal control benefits stems from the wide range of satellite configurations and infinite variation in inertia. To mitigate this issue, this research aims to determine an analytical relationship between satellite inertia and the time savings of using optimal control rather than eigenaxis maneuvering on spacecraft with a NASA standard four reaction-wheel configuration. To accomplish this, the development of a script using DIDO R optimization software determines minimum-time paths for satellite maneuvers. Each path was independently verified and validated using Pontryagin's minimization principle to ensure that they are physically feasible and that each solution is optimal. Additionally, this work demonstrates that inertia ratios can be used to characterize the attitude control performance of any spacecraft, allowing for the analysis of satellite inertias and their relationship to maneuver time reduction regardless of the scale of the spacecraft. The calculation of the agility envelope volume is then utilized in conjunction with the DIDO R script and various inertia ratios in order to investigate the mathematical relationship between satellite inertia and time savings from optimal control. The result of this work is a design-space tool that can be used by engineers to help determine whether or not to implement time-optimal control algorithms on any spacecraft in a simple and effective way."--from abstract.