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

Drag-free and attitude control for the GOCE satellite

The paper concerns Drag-Free and Attitude Control of the European satellite Gravity field and steady-state Ocean Circulation Explorer (GOCE) during the science phase. Design has followed Embedded Model Control, where a spacecraft/environment discrete-time model becomes the realtime control core and is interfaced to actuators and sensors via tuneable feedback laws. Drag-free control implies cancelling non-gravitational forces and all torques, leaving the satellite to free fall subject only to gravity. In addition, for reasons of science, the spacecraft must be carefully aligned to the local orbital frame, retrieved from range and rate of a Global Positioning System receiver. Accurate drag-free and attitude control requires proportional and low-noise thrusting, which in turn raises the problem of propellant saving. Six-axis drag-free control is driven by accurate acceleration measurements provided by the mission payload. Their angular components must be combined with the star-tracker attitude so as to compensate accelerometer drift. Simulated results are presented and discusse

Mass and power modeling of communication satellites

Analytic estimating relationships for the mass and power requirements for major satellite subsystems are described. The model for each subsystem is keyed to the performance drivers and system requirements that influence their selection and use. Guidelines are also given for choosing among alternative technologies which accounts for other significant variables such as cost, risk, schedule, operations, heritage, and life requirements. These models are intended for application to first order systems analyses, where resources do not warrant detailed development of a communications system scenario. Given this ground rule, the models are simplified to 'smoothed' representation of reality. Therefore, the user is cautioned that cost, schedule, and risk may be significantly impacted where interpolations are sufficiently different from existing hardware as to warrant development of new devices

Robust Control Allocation among Overactuated Spacecraft Thrusters under Ellipsoidal Uncertainty

Spacecrafts with overactuated and redundant thrusters can work normally once some of them are out of work, which improves the reliability of spacecraft in orbit. In this way, the desired command of controller needs to be dynamically allocated among thrusters. Considering that uncertain factors may appear in forms of dynamics, installation errors, thrust deviations, or failures, this paper proposes a robust control allocation under ellipsoidal uncertainty. This method uses the uncertainty ellipsoid set to describe the uncertainty of the thrusters firstly and establish the thrust allocation robust reference model and then transforms it into a cone optimization model which can be solved as an optimized problem. Finally, this paper adopts the interior-point method for solving the optimization problem. In this way, difficulties of solving the problem caused by parameter uncertainties are avoided effectively. Finally, we take satellite rendezvous and docking as simulation scenarios; it is verified that the cumulative distribution error and maximum error can be reduced by more than 15% when the random error of control efficiency matrix is 5%–20%; also, precision of thruster allocation is improved

Autonomous thruster failure recovery on underactuated spacecraft using model predictive control

Thruster failures historically account for a large percentage of failures that have occurred on orbit. These failures are typically handled through redundancy, however, with the push to using smaller, less expensive satellites in clusters or formations there is a need to perform thruster failure recovery without additional hardware. This means that a thruster failure may cause the spacecraft to become underactuated, requiring more advanced control techniques. A model of a thruster-controlled spacecraft is developed and analyzed with a nonlinear controllability test, highlighting several challenges including coupling, nonlinearities, severe control input saturation, and nonholonomicity. Model Predictive Control (MPC) is proposed as a control technique to solve these challenges. However, the real-time, online implementation of MPC brings about many issues. A method of performing MPC online is described, implemented and tested in simulation as well as in hardware on the Synchronized Position-Hold, Engage, Reorient Experimental Satellites (SPHERES) testbed at the Massachusetts Institute of Technology (MIT) and on the International Space Station (ISS). These results show that MPC provided improved performance over a simple path planning technique

A Comparison of Structurally Connected and Multiple Spacecraft Interferometers

Structurally connected and multiple spacecraft interferometers are compared in an attempt to establish the maximum baseline (referred to as the "cross-over baseline") for which it is preferable to operate a single-structure interferometer in space rather than an interferometer composed of numerous, smaller spacecraft. This comparison is made using the total launched mass of each configuration as the comparison metric. A framework of study within which structurally connected and multiple spacecraft interferometers can be compared is presented in block diagram form. This methodology is then applied to twenty-two different combinations of trade space parameters to investigate the effects of different orbits, orientations, truss materials, propellants, attitude control actuators, onboard disturbance sources, and performance requirements on the cross-over baseline. Rotating interferometers and the potential advantages of adding active structural control to the connected truss of the structurally connected interferometer are also examined. The minimum mass design of the structurally connected interferometer that meets all performance-requirements and satisfies all imposed constraints is determined as a function of baseline. This minimum mass design is then compared to the design of the multiple spacecraft interferometer. It is discovered that the design of the minimum mass structurally connected interferometer that meets all performance requirements and constraints in solar orbit is limited by the minimum allowable aspect ratio, areal density, and gage of the struts. In the formulation of the problem used in this study, there is no advantage to adding active structural control to the truss for interferometers in solar orbit. The cross-over baseline for missions of practical duration (ranging from one week to thirty years) in solar orbit is approximately 400 m for non-rotating interferometers and 650 m for rotating interferometers

Generic Control Allocation Toolbox for Preliminary Vehicle Design

This paper describes the development and application of a Generic Control Allocation Toolbox developed at NASA Langley Research Center (LGCAT) intended to aid engineers during the preliminary design phase of an aerospace vehicle. The static controllability space in the forms of a Theoretical Attainable Moment Set, , or Theoretical Attainable Force Set, is difficult to visualize for modern vehicles with multiple types of redundant control effectors. The objective of LGCAT is to provide system engineers and designers early in the vehicle design phase with quick insights on how control effector parameters such as quantity, sizing, location, orientation, redundancy, etc., affect the overall controllability and other performance metrics. Having such information in hand allows system engineers to make more informed decisions on overall mission objectives such as performance vs. reliability vs. cost, etc. early in a vehicle design phase and reduce the number of iterations necessary in the design and analysis cycles. LGCAT can accept a variety of control effector types including aerodynamic surfaces, rotors, thrust vector control (TVC) engines, and reaction control systems (RCS). LGCAT is MATLAB based, user friendly, and is capable of performing the analysis in the Graphical User Interface (GUI) or script mode. Current add-on features include interfacing with engineering level codes such as Vehicle Sketch Pad (VSP) and generating the corresponding and for an arbitrary vehicle design. These capabilities potentially make LGCAT an integral part of the preliminary design phase for any vehicle

Space applications of Automation, Robotics and Machine Intelligence Systems (ARAMIS). Volume 2: Space projects overview

Applications of automation, robotics, and machine intelligence systems (ARAMIS) to space activities, and their related ground support functions are studied so that informed decisions can be made on which aspects of ARAMIS to develop. The space project breakdowns, which are used to identify tasks ('functional elements'), are described. The study method concentrates on the production of a matrix relating space project tasks to pieces of ARAMIS