Coupled orbit and attitude dynamics of a reconfigurable spacecraft with solar radiation pressure

This work investigates the orbital and attitude dynamics of future reconfigurable multi-panel solar sails able to change their shape during a mission. This can be enabled either by changing the relative position of the individual panels, or by using articulated mechanisms and deployable, retractable and/or inflatable structures. Such a model introduces the concept of modular spacecraft of variable morphology to large gossamer spacecraft. However, this joint concept is complex in nature and requires equations for coupled orbit/attitude dynamics. Therefore, as a starting point, the system is modelled as a rigid-body dumbbell consisting of two tip masses connected by a rigid, massless panel. The system is subjected to a central gravitational force field under consideration of solar radiation pressure forces. Therefore, we assign reflectivity coefficients to the tip masses and a high area-to-mass ratio. An analytical Hamiltonian approach is used to describe the planar motion of the system in Sun-centred Keplerian and non-Keplerian circular orbits. The stability and controllability of the system is enabled through changing the reflectivity coefficients, for example through the use of electro-chromic coating on its surface. The creation of artificial unstable equilibria of the system due to the presence of solar radiation pressure and heteroclinic connections between the equilibria are investigated. We further derive a constraint for the solar radiation pressure forces to maintain the system on a circular Sun-centred orbit. It is planned that the structure is eventually capable of reconfiguring between the equilibria by a minimum actuation effort

Coupled orbit and attitude dynamics of a reconfigurable spacecraft with solar radiation pressure

This work investigates the orbital and attitude dynamics of future reconfigurable multi-panel solar sails able to change their shape during a mission. This can be enabled either by changing the relative position of the individual panels, or by using articulated mechanisms and deployable, retractable and/or inflatable structures. Such a model introduces the concept of modular spacecraft of variable morphology to large gossamer spacecraft. However, this joint concept is complex in nature and requires equations for coupled orbit/attitude dynamics. Therefore, as a starting point, the system is modelled as a rigid-body dumbbell consisting of two tip masses connected by a rigid, massless panel. The system is subjected to a central gravitational force field under consideration of solar radiation pressure forces. Therefore, we assign reflectivity coefficients to the tip masses and a high area-to-mass ratio. An analytical Hamiltonian approach is used to describe the planar motion of the system in Sun-centred Keplerian and non-Keplerian circular orbits. The stability and controllability of the system is enabled through changing the reflectivity coefficients, for example through the use of electro-chromic coating on its surface. The creation of artificial unstable equilibria of the system due to the presence of solar radiation pressure and heteroclinic connections between the equilibria are investigated. We further derive a constraint for the solar radiation pressure forces to maintain the system on a circular Sun-centred orbit. It is planned that the structure is eventually capable of reconfiguring between the equilibria by a minimum actuation effort

Suboptimal LQR-based spacecraft full motion control: Theory and experimentation

Abstract This work introduces a real time suboptimal control algorithm for six-degree-of-freedom spacecraft maneuvering based on a State-Dependent-Algebraic-Riccati-Equation (SDARE) approach and real-time linearization of the equations of motion. The control strategy is sub-optimal since the gains of the linear quadratic regulator (LQR) are re-computed at each sample time. The cost function of the proposed controller has been compared with the one obtained via a general purpose optimal control software, showing, on average, an increase in control effort of approximately 15%, compensated by real-time implementability. Lastly, the paper presents experimental tests on a hardware-in-the-loop six-degree-of-freedom spacecraft simulator, designed for testing new guidance, navigation, and control algorithms for nano-satellites in a one-g laboratory environment. The tests show the real-time feasibility of the proposed approach

Optimized Thruster Allocation Utilizing Dual Quaternions for the Asteroid Sample Return Mission (OSIRIS-REx)

As spacecraft require higher positional accuracy from the attitude control systems, new algorithm developments, along with sensor and actuator resolution and range improvements are necessary to achieve the desired science accuracies. For agile 6-Degrees of freedom (6-DOF) spacecraft with redundancy, the actuators are usually oversized or overpopulated to meet the desired slew requirements. Currently, most spacecraft utilize an over-actuated thruster system to produce 6-DOF control. This thesis presents a simulation of the OSIRIS-REx mission during the descent phase to the asteroid Bennu, with a focus on utilizing dual quaternion dynamics and a newly developed thruster allocation method. The dual quaternion based dynamics are chosen in order to demonstrate its feasibility in real-time applications. Contrary to typical plant dynamics, which decouple the spacecraft orbit and attitude dynamics, the dual quaternion description provides a compact and coupled dynamics system. Due to the coupled nature of dual quaternions, a newly developed thruster distribution matrix is implemented to take both the coupled command body forces and torques and transform them into the individual thruster frames. The developed method is based on a min-max optimization that results in a constant thruster distribution matrix. From the optimization, a minimum thrust solution is calculated for the coupled position and attitude commands. Therefore, its integration into the dual quaternion dynamics is intuitive and simplistic. The final result is a computationally fast thruster allocation solution for real-time applications

AUTONOMOUS SPACECRAFT RENDEZVOUS WITH A TUMBLING OBJECT: APPLIED REACHABILITY ANALYSIS AND GUIDANCE AND CONTROL STRATEGIES

Rendezvous and proximity operations are an essential component of both military and commercial space missions and are rising in complexity. This dissertation presents an applied reachability analysis and develops a computationally feasible autonomous guidance algorithm for the purpose of spacecraft rendezvous and proximity maneuvers around a tumbling object. Recent advancements enable the use of more sophisticated, computation-based algorithms, instead of traditional control methods. These algorithms are desirable for autonomous applications due to their ability to optimize performance and explicitly handle constraints (e.g., safety, control limits). In an autonomous setting, however, some important questions must be answered before an algorithm implementation can be realized. First, the feasibility of a maneuver is addressed by analyzing the fundamental spacecraft relative dynamics. Particularly, a set of initial relative states is computed and visualized from which the desired rendezvous state can be reached (i.e., backward reachability analysis). Second, with the knowledge that a maneuver is feasible, the Model Predictive Control (MPC) framework is utilized to design a stabilizing feedback control law that optimizes performance and incorporates constraints such as control saturation limits and collision avoidance. The MPC algorithm offers a computationally efficient guidance strategy that could potentially be implemented in real-time on-board a spacecraft.http://archive.org/details/autonomousspacec1094560364Major, United States Air ForceApproved for public release; distribution is unlimited

Deflecting small asteroids using laser ablation : Deep space navigation and asteroid orbit control for LightTouch2 Mission

This paper presents a low-cost, low mass, mission design to successfully intercept and deflect a small and faint, 4 m in diameter asteroid. Intended to be launched after 2025, the laser-ablating mission, LightTouch2 will be used to deflect the orbit of the asteroid by at least 1 m/s. This will be achieved with a total mission lifetime of less than three years. Analysis includes the initial approach of the spacecraft, the operations of the laser at an optimal spacecraft-to-asteroid distance of 50 m and the relative orbit of the spacecraft that flies in formation with the asteroid. Analysis includes line-of-sight measurements with radiometric tracking from ground station to improve the trajectory estimate and observability of the spacecraft, collision avoidance and mapping strategies. The spacecraft will also need optimal discrete control. This is achieved by impulse-bit manoeuvres used to account for the perturbations caused by the resultant thrust on the asteroid, plume impingement, laser recoil and solar radiation pressure. The spacecraft controls its trajectory within a 1 m box from the reference trajectory to enable the laser to optimally focussing the laser beam. The proposed approach uses an unscented Kalman filter to estimate the relative spacecraft-asteroid position, velocity and perturbative acceleration

Spacecraft Alignment Determination and Control for Dual Spacecraft Precision Formation Flying

Many proposed formation flying missions seek to advance the state of the art in spacecraft science imaging by utilizing precision dual spacecraft formation flying to enable a virtual space telescope. Using precision dual spacecraft alignment, very long focal lengths can be achieved by locating the optics on one spacecraft and the detector on the other. Proposed science missions include astrophysics concepts with spacecraft separations from 1000 km to 25,000 km, such as the Milli-Arc-Second Structure Imager (MASSIM) and the New Worlds Observer, and Heliophysics concepts for solar coronagraphs and X-ray imaging with smaller separations (50m 500m). All of these proposed missions require advances in guidance, navigation, and control (GNC) for precision formation flying. In particular, very precise astrometric alignment control and estimation is required for precise inertial pointing of the virtual space telescope to enable science imaging orders of magnitude better than can be achieved with conventional single spacecraft instruments. This work develops design architectures, algorithms, and performance analysis of proposed GNC systems for precision dual spacecraft astrometric alignment. These systems employ a variety of GNC sensors and actuators, including laser-based alignment and ranging systems, optical imaging sensors (e.g. guide star telescope), inertial measurement units (IMU), as well as micro-thruster and precision stabilized platforms. A comprehensive GNC performance analysis is given for Heliophysics dual spacecraft PFF imaging mission concept

Proximity Navigation of Highly Constrained Spacecraft

Bandit is a 3-kg automated spacecraft in development at Washington University in St. Louis. Bandit's primary mission is to demonstrate proximity navigation, including docking, around a 25-kg student-built host spacecraft. However, because of extreme constraints in mass, power and volume, traditional sensing and actuation methods are not available. In particular, Bandit carries only 8 fixed-magnitude cold-gas thrusters to control its 6 DOF motion. Bandit lacks true inertial sensing, and the ability to sense position relative to the host has error bounds that approach the size of the Bandit itself. Some of the navigation problems are addressed through an extremely robust, error-tolerant soft dock. In addition, we have identified a control methodology that performs well in this constrained environment: behavior-based velocity potential functions, which use a minimum-seeking method similar to Lyapunov functions. We have also adapted the discrete Kalman filter for use on Bandit for position estimation and have developed a similar measurement vs. propagation weighting algorithm for attitude estimation. This paper provides an overview of Bandit and describes the control and estimation approach. Results using our 6DOF flight simulator are provided, demonstrating that these methods show promise for flight use

Coupled orbit-attitude mission design in the circular restricted three-body problem

Trajectory design increasingly leverages multi-body dynamical structures that are based on an understanding of various types of orbits in the Circular Restricted Three-Body Problem (CR3BP). Given the more complex dynamical environment, mission applications may also benefit from deeper insight into the attitude motion. In this investigation, the attitude dynamics are coupled with the trajectories in the CR3BP. In a highly sensitive dynamical model, such as the orbit-attitude CR3BP, periodic solutions allow delineation of the fundamental dynamical structures. Periodic solutions are also a subset of motions that are bounded over an infinite time-span (assuming no perturbing factors), without the necessity to integrate over an infinite time interval. Euler equations of motion and quaternion kinematics describe the rotational behavior of the spacecraft, whereas the translation of the center of mass is modeled in the CR3BP equations. A multiple shooting and continuation procedure are employed to target orbit-attitude periodic solutions in this model. Application of Floquet theory, Poincaré mapping, and grid search to identify initial guesses for the targeting algorithm is described. In the Earth-Moon system, representative scenarios are explored for axisymmetric vehicles with various inertia characteristics, assuming that the vehicles move along Lyapunov, halo as well as distant retrograde orbits. A rich structure of possible periodic behaviors appears to pervade the solution space in the coupled problem. The stability analysis of the attitude dynamics for the selected families is included. Among the computed solutions, marginally stable and slowly diverging rotational behaviors exist and may offer interesting mission applications. Additionally, the solar radiation pressure is included and a fully coupled orbit-attitude model is developed. With specific application to solar sails, various guidance algorithms are explored to direct the spacecraft along a desired path, when the mutual interaction between orbit and attitude dynamics is considered. Each strategy implements a different form of control input, ranging from instantaneous reorientation of the sail pointing direction to the application of control torques, and it is demonstrated within a simple station keeping scenario

Optimal Guidance and Control for Electromagnetic Formation Flying

学位の種別: 修士University of Tokyo(東京大学