Three-axis attitude determination via Kalman filtering of magnetometer data

A three-axis Magnetometer/Kalman Filter attitude determination system for a spacecraft in low-altitude Earth orbit is developed, analyzed, and simulation tested. The motivation for developing this system is to achieve light weight and low cost for an attitude determination system. The extended Kalman filter estimates the attitude, attitude rates, and constant disturbance torques. Accuracy near that of the International Geomagnetic Reference Field model is achieved. Covariance computation and simulation testing demonstrate the filter's accuracy. One test case, a gravity-gradient stabilized spacecraft with a pitch momentum wheel and a magnetically-anchored damper, is a real satellite on which this attitude determination system will be used. The application to a nadir pointing satellite and the estimation of disturbance torques represent the significant extensions contributed by this paper. Beyond its usefulness purely for attitude determination, this system could be used as part of a low-cost three-axis attitude stabilization system

Design and implementation of an attitude determination and control system for the AntelSat

This thesis describes the design, analysis and construction of the Attitude Determination and Control System (ADCS) for the first Uruguayan nanosatellite, the AntelSat. The AntelSat project is a joint venture between the Electrical Engineering Institute (IIE) of Faculty of Engineering, Universidad de la República (UdelaR University) and Antel, the Uruguayan national telecommunications company. The satellite consists of a two-unit (2U) CubeSat, which implies that the ADCS is designed under tight mass, size, and energy constraints. In addition, these kind of satellites usually have limited sensing, computational and communication capabilities, motivating the need for autonomous and computationally eficient algorithms. Under these strict restraints, developing an effective attitude control system poses a significant challenge. As presented in this thesis, for the attitude determination section of the ADCS, data available from sensors is taken as inputs for the computation of an optimal quaternion estimator. The use of a quaternion implementation of an unscented Kalman filter is also discussed. Additionally, attitude control is based on magnetic actuation with magnetorquers being commanded by pulse width modulation. It is shown that the control system is able to achieve the detumbling of the satellite after separation from the launch interface using the reliable B-dot control law. Nadirpointing control is achieved with the use of a simple Linear Quadratic Regulator. Also pertinent is the simulation environment that was implemented to develop the attitude determination and control algorithms and also to validate their performance. ADCS hardware prototypes and flight versions that were designed and constructed are introduced.Este documento de tesis describe el diseño, análisis y construcción de el Sistema de Determinación y Control de Actitud (ADCS por sus siglas en inglés) del primer satélite uruguayo, el AntelSat. El proyecto AntelSat es una actividad conjunta entre el Instituto de Ingeniería Eléctrica (IIE) de la Facultad de Ingeniería de la Universidad de la República y Antel, la empresa de telecomunicaciones nacional de Uruguay. El satélite consiste en un CubeSat de dos unidades (2U), lo que implica que el ADCS es diseñado bajo estrictas restricciones de masa, tamaño y energía. Además, este tipo de satélites posee una capacidad computacional, de comunicaciones y de medición limitada, lo que motiva la necesidad de lograr algoritmos computacionalmente eficientes. Bajo estas estrictas limitaciones, el desarrollo de un sistema de control de actitud efectivo se traduce en un reto importante. Como se presenta en esta tesis, para el segmento de determinación de actitud del ADCS, la información proveniente de los sensores es tomada como entrada para el cálculo de un estimador de cuaternión óptimo. Se discute también el uso de una implementación con cuaterniones de un filtro de Kalman "unscented". Por otro lado, el control de actitud está basado en actuación magnética con magnetorquers comandados con modulación de ancho de pulso. Se demuestra que el sistema de control es capaz de reducir el valor de velocidad angular del satélite en la fase previa a la separación con la interfaz de lanzamiento, mediante la utilización del algoritmo B-dot. La estabilización de la actitud en modo de apunte al nadir se logra con el uso de un simple regulador lineal cuadrático. Por otra parte, se presenta el entorno de simulación que fue implementado para el desarrollo de algoritmos de determinación y control de actitud, y también para validar el desempeño de los mismos. A su vez, se exhiben el hardware del ADCS que fue diseñado y construido, tanto prototipos como versiones de vuelo

ALGORITHMS AND OPTIMAL CONTROL FOR SPACECRAFT MAGNETIC ATTITUDE MANEUVERS

This study focused on providing applicable control solutions for spacecraft magnetic attitude control system. Basically, two main lines are pursued; first, developing detumbling control laws and second, an improvement in the three-axis attitude control schemes by extending magnetic rods activation time. Spacecraft, after separation from the launching mechanism, experiences a tumbling phase due to an undesired angular momentum. In this study, we present a new efficient variant of the B-dot detumbling law by introducing a substitute of the spacecraft angular velocity, based on the ambient magnetic field data. This B-dot law preserves the orthogonality, among the applied torque, dipole moment and magnetic field vectors. Most of the existing variants of the B-dot law in the literature don\u27t preserve this orthogonality. Furthermore, the problem of minimum-time spacecraft magnetic detumbling is revisited within the context of optimal control theory. Two formulations are presented; the first one assumes the availability of the angular velocity measurements for feedback. The second formulation assumes the availability of only the ambient magnetic field measurements in the feedback; the latter is considered another optimal-based B-dot law. A reduction in detumbling time is fulfilled by the proposed laws along with less power consumption for the proposed B-dot laws. In magnetic attitude maneuvers, magnetic rods and magnetometers usually operate alternatively, to avoid the magnetic rods\u27 noise effect on magnetometers measurements. Because of that, there will be no control authority over the spacecraft during the magnetometer measurement period. Hence longer maneuver times are usually experienced. In this study, a control scheme that enables the extension of the magnetic rods’ activation time is developed, regardless of the attitude control law. The key concept is replacing the real magnetic field measurement by a pseudo measurement, which is computed based on other sensors measurements. By applying a known command to the spacecraft and measuring the spacecraft response, it is possible to compute the ambient magnetic field around the spacecraft. The system mathematical singularity is solved using the Tikhonov regularization approach. Another developed approach estimates the magnetic field, using a relatively simple and fast dynamic model inside a Multiplicative Extended Kalman Filter. A less maneuver time with less power consumption are fulfilled. These control approaches are further validated using real telemetry data from CASSIOPE mission. This dissertation develops a stability analysis for the spacecraft magnetic attitude control, taking into consideration the alternate operation between the magnetic rods and the magnetometers. It is shown that the system stability degrades because of this alternate operation, supporting the proposed approach of extending the operation time of the magnetic rods

AN ATTITUDE DETERMINATION SYSTEM WITH MEMS GYROSCOPE DRIFT COMPENSATION FOR SMALL SATELLITES

This thesis presents the design of an attitude determination system for small satellites that automatically corrects for attitude drift. Existing attitude determination systems suffer from attitude drift due to the integration of noisy rate gyro sensors used to measure the change in attitude. This attitude drift leads to a gradual loss in attitude knowledge, as error between the estimated attitude and the actual attitude increases. In this thesis a Kalman filter is used to complete sensor fusion which combines sensor observations with a projected attitude based on the dynamics of the satellite. The system proposed in this thesis also utilizes a novel sensor called the stellar gyro to correct for the drift. The stellar gyro compares star field images taken at different times to determine orientation, and works in the presence of the sun and during eclipse. This device provides a relative attitude fix that can be used to update the attitude estimate provided by the Kalman filter, effectively compensating for drift. Simulink models are developed of the hardware and algorithms to model the effectiveness of the system. The Simulink models show that the attitude determination system is highly accurate, with steady state errors of less than 1 degree

ACS Without an Attitude

The book (ACS without an Attitude) is an introduction to spacecraft attitude control systems. It is based on a series of lectures that Dr. Hallock presented in the early 2000s to members of the GSFC flight software branch, the target audience being flight software engineers (developers and testers), fairly new to the field that desire an introductory understanding of spacecraft attitude determination and control

Implementation and validation of the attitude and determination control system of a pico-satellite

This end-of-degree project focuses on the software development of a pico-satellite Attitude Determination and Control System, using MATLAB, as well as its subsequent integration in C. The software carried out will be implemented in a 5 cm sided pico-satellite, called PocketQube. The research project will form part of the IEEE Open PocketQube Kit mission developed in the NanoSat laboratory belonging to the UPC. The pico-satellite attitude system can be broken down into two distinct branches, attitude determination and attitude control. Attitude determination involves the process of determining the orientation of the satellite using the measurements acquired by its sensors (they include photodiodes, gyroscopes, and magnetometers). On the other hand, attitude control is the process by which the orientation of the pico-satellite is controlled using specific control algorithms (such as detumbling and nadir pointing) and actuators, (such as magnetorquers). To analyse and understand the attitude of the pico-satellite, a model is developed that allows simulating the environment in which the satellite is located. This model encompasses pico-satellite dynamics, a model that describes orbital dynamics, and another that represents the various perturbation forces that affect the pico-satellite. In this way, it is possible to describe the orbit followed by the pico-satellite, including its position, speed and the external forces that influence and affect its behaviour during the orbital flight. Once the modelling of the simulation environment is complete, the determination and control algorithms are implemented, as well as the mathematical models that describe the behaviour of the sensors and actuators. These algorithms and models are designed with the aim of meeting the requirements established for the Determination and Control System. Subsequently, an exhaustive analysis of the results obtained during the simulation is carried out. The purpose of this analysis is to verify if the previously established requirements for the Attitude Determination and Control System are met. The conclusion of this work allows to start the test campaigns of this system in the PocketQube hardware, as well as to check if the results obtained in the simulations correspond to reality

CubeSat Design and Attitude Control with Micro Pulsed Plasma Thrusters

This study presents the overall design of a 3U CubeSat equipped with commercial-off-the shelf hardware, Teflon-fueled micro-Pulsed Plasma Thrusters (µPPT) and an attitude determination and control system. The µPPT is sized by the impulse bit and pulse frequency required for continuous compensation of expected maximum disturbance torques at altitudes between 400 and 1000 km, and to perform stabilization of up to 20 deg/s and slew maneuvers of up to 180 degrees. The study involves realistic power constraints anticipated on the 3U CubeSat. Attitude estimation is implemented using the q-method for static attitude determination of the quaternion using pairs of the spacecraft-sun and magnetic field vectors. The quaternion estimate and the gyroscope measurements are used with an extended Kalman filter to obtain the attitude estimates. Proportional and derivative control algorithms use the static attitude estimation in order to calculate the angular momentum required to compensate for the disturbance torques and to achieve specified stabilization and slewing maneuvers or combinations. Two control methods are developed: paired firing method, and separate control algorithm and thruster allocation methods which determines the optimal utilization of the available thrusters and introduces redundancy. Simulations results are presented for a 3U CubeSat under stabilization, pointing, and pointing and spinning scenarios

A Micromechanical INS/GPS System for Small Satellites

The cost and complexity of large satellite space missions continue to escalate. To reduce costs, more attention is being directed toward small lightweight satellites where future demand is expected to grow dramatically. Specifically, micromechanical inertial systems and microstrip global positioning system (GPS) antennas incorporating flip-chip bonding, application specific integrated circuits (ASIC) and MCM technologies will be required. Traditional microsatellite pointing systems do not employ active control. Many systems allow the satellite to point coarsely using gravity gradient, then attempt to maintain the image on the focal plane with fast-steering mirrors. Draper's approach is to actively control the line of sight pointing by utilizing on-board attitude determination with micromechanical inertial sensors and reaction wheel control actuators. Draper has developed commercial and tactical-grade micromechanical inertial sensors, The small size, low weight, and low cost of these gyroscopes and accelerometers enable systems previously impractical because of size and cost. Evolving micromechanical inertial sensors can be applied to closed-loop, active control of small satellites for micro-radian precision-pointing missions. An inertial reference feedback control loop can be used to determine attitude and line of sight jitter to provide error information to the controller for correction. At low frequencies, the error signal is provided by GPS. At higher frequencies, feedback is provided by the micromechanical gyros. This blending of sensors provides wide-band sensing from dc to operational frequencies. First order simulation has shown that the performance of existing micromechanical gyros, with integrated GPS, is feasible for a pointing mission of 10 micro-radians of jitter stability and approximately 1 milli-radian absolute error, for a satellite with 1 meter antenna separation. Improved performance micromechanical sensors currently under development will be suitable for a range of micro-nano-satellite applications

COBE attitude as seen from the FDF

The goal of the Flight Dynamics Facility (FDF) attitude support is twofold: to determine spacecraft attitude and to explain deviations from nominal attitude behavior. Attitude determination often requires resolving contradictions in the sensor observations. This may be accomplished by applying calibration corrections or by revising the observation models. After accounting for all known sources of error, solution accuracy should be limited only by observation and propagation noise. The second half of the goal is to explain why the attitude may not be as originally intended. Reasons for such deviations include sensor or actuator misalignments and control system performance. In these cases, the ability to explain the behavior should, in principle, be limited only by knowledge of the sensor and actuator data and external torques. Documented here are some results obtained to date in support of the Cosmic Background Explorer (COBE). Advantages and shortcomings of the integrated attitude determination/sensor calibration software are discussed. Some preliminary attitude solutions using data from the Diffuse Infrared Background Experiment (DIRBE) instrument are presented and compared to solutions using Sun and Earth sensors. A dynamical model is constructed to illustrate the relative importance of the various sensor imprefections. This model also shows the connection between the high- and low-frequency attitude oscillations

Advances in Spacecraft Attitude Control

Spacecraft attitude maneuvers comply with Euler's moment equations, a set of three nonlinear, coupled differential equations. Nonlinearities complicate the mathematical treatment of the seemingly simple action of rotating, and these complications lead to a robust lineage of research. This book is meant for basic scientifically inclined readers, and commences with a chapter on the basics of spaceflight and leverages this remediation to reveal very advanced topics to new spaceflight enthusiasts. The topics learned from reading this text will prepare students and faculties to investigate interesting spaceflight problems in an era where cube satellites have made such investigations attainable by even small universities. It is the fondest hope of the editor and authors that readers enjoy this book