Since the dawn of the space race, satellites have grown rapidly in complexity and shrunk equally rapidly in size. Most of them contain an Attitude Determination and Control System (ADCS) on board for pointing and detumbling manoeuvres. These intricate systems are designed for an outer space environment, hence, phenomenon otherwise abscent in space, such as gravity and aerodynamic drag present a challenge in validating these systems on Earth. The gimbal suspension testbed aims to provide a 3 Degree of Freedom (DoF) suspension where the mounted satellite under test can rotate about either axis. The suspension induces disturbance torques that must be modeled in order for the testbed to be characterized. This is accomplished by formulating the necessary gimbal dynamics, bearing friction, aerodynamic and Center of Mass (CoM) displacement torque model. This yields a relationship from which all torques present in the system can be expressed in terms of the angles, angular velocities and angular accelerations of the gimbal frames. By measuring the angles and obtaining the velocities and accelerations through numerical differentiation, the torques that correspond to a certain motion can be calculated. Furthermore, the thesis covers the iterative design of the gimbal suspension and all of its constituents, the angular measurement method and a Finite Element Method (FEM) simulation to estimate deformations. The result is presented in terms of a simulation that validates the models by predicting its behaviour for certain movement. The final result is a series of characterization plots that tells the user of the gimbal testbed how much torque must be produced by the CubeSat ADCS in order to operate it. / Sedan begynnelsen av rymdkapplöpningen har satelliter snabbt ökat i komplexitet och lika snabbt minskat i storlek. De flesta satelliter har ett attitydsbestänings- och kontrollsystem (ADCS) ombord för att kunna utföra vissa manövrar. Dessa system är designade för rymdmiljön, därför kan fenomen som annars är frånvarande i rymden, så som gravitation och luftmotstånd, innebära en utmaning då man önskar att validera systemet på jorden. Gimbalupphängningen förmedlar rotation med tre frihetsgrader där satelliten under test kan rotera kring alla tre axlar. Upphängningen inducerar störmoment som måste modelleras för att den ska bli ordentligt karaktäriserad. Detta åstadkoms genom att formulera gimbalens dynamiska förhållanden, kullagerfriktion, luftmotstånd och masscenterförflyttning. Dessa samband kopplar samman alla moment som är närvarande i systemet som funktion av gimbalramarnas vinklar, vinkelhastigheter och vinkelaccelerationer. Genom att mäta vinklarna och erhålla vinkelhastigheter och vinkelacceleration genom numerisk derivering kan momenten som motsvarar den uppmätta rörelsen beräknas. Dessutom presenteras den iterativa designen av gimbalupphängningen och alla dess beståndsdelar, vinkelmätningsmetoden och en finita elementmetodssimulering för att uppskaffa deformationer. Resultatet presenteras i form av simuleringar som validerar modellen genom att förutspå dess beteende för viss rörelse. Det slutgiltiga resultatet är en serie av karaktäriseringsgrafer som förmedlar till användaren just hur mycket moment dess CubeSats ADCS måste producera för att kunna använda gimbalupphängingen.

This thesis presents the design of a working and testable satellite with particular emphasis on the electrical, mechanical, and thermal modelling and performance issues for the ECOSat project in the framework of the Canadian Satellite Design Competition. In order of importance, based on the design challenges for the satellite structure were the dynamics modelling and analysis, thermal modeling and analysis, and assembly and integration modeling. Both the dynamics and thermal modeling of the satellite were completed using Finite Element Analysis (FEA) in NX with the NASTRAN solver. The dynamic analysis study was performed first since it has the primary design driver for the structure. These frequencies are of concern due to the 90 Hz or greater fundamental frequency requirement for each axis. The dynamic modes of the satellite structure had the largest influence not only on the design of the structure but also its interface to the electronic systems as these had to meet the required testing qualification levels. It was found that the first fundamental frequency appeared near 200 Hz in the XY plane of the structure. The second study performed was on the thermal modeling of the satellite both for extreme operating conditions in “Hot” and “Cold” cases. Operational limiting cases were identified for the batteries in the cold and hot case study, and the power amplifier for the transmitter was identified for the hot case study. For the batteries to perform satisfactorily for the cold and hot case problem, a metal bracket with an electric heater was added to the design. The heaters were added to the design as a resistive heating element, the additional thermal coupling from the bracket improved heat transfer during the hot case. A trade study analysis was conducted for the power amplifier. Here, a bi directional heat spreader made of pyrolytic graphite attached to a frame member with high thermal inertia was chosen as the optimal solution. Finally, the third study performed tested the interface and clearance requirements of the satellite. The synergistic integration of the electrical and mechanical systems required significant attention in order to ensure the successful assembly, integration, and testing of the two systems. The investigation focused on the cabling assemblies of the satellite. Several design iterations were required for the power regulation, transmitter, receiver, modem, and onboard computer systems. Detailed assembly drawings were created for the cabling assembly fabrication prior to the final integration of the electrical and mechanical systems. The performance simulations show that the satellite systems meet or exceed the required launch qualification tests as well as the thermal cycling requirements for all systems and their components to operate within the manufacturer specified values. Once completely assembled and launched into orbit, the satellite should be able to perform and within its operational and mission requirements in both a sun synchronous or polar orbit at a range of altitudes. / Graduate / 0538 / 0544 / 0548 / jtcurran@uvic.ca

Magíster en Ciencias de la Ingeniería, Mención Mecánica / El presente trabajo estudia el uso de algoritmos genéticos para automatizar el diseño del sistema de control térmico pasivo de un nanosatélite tipo CubeSat de 3U en una configuración HotCase. Una metodología de tres pasos es propuesta. En el primer paso, las soluciones candidatas se representan por medio un mosaico de dos materiales diferentes distribuidos sobre las caras externas de un satélite tipo CubeSat. Estas soluciones son evaluadas utilizando el método de los elementos finitos (FEM) y clasificadas de acuerdo a su capacidad de cumplir los requerimientos de temperatura considerados para la misión. En un segundo paso se construye un modelo a escala del mejor individuo. Se realiza un ensayo térmico del modelo creado, describiendo el montaje y la electrónica necesaria para obtener la información de emperatura sobre las caras del CubeSat. En un tercer paso la simulación es adaptada utilizando los datos obtenidos en el ensayo térmico con el propósito de validar el proceso de optimización. Para este procedimiento se actualizan los parámetros con los cuales se analiza la solución optimizada ajustándolos a los utilizados en el ensayo experimental. Las diferencias entre los resultados de simulación y los obtenidos a través del ensayo térmico se encuentran en el orden de los 1,45 (K) para las caras que apuntan directamente al sol y 2,4 (K) para las caras que están directamente apuntando a la superficie terrestre. Ambos resultados considerando un error cuadrático medio. También se observó que el diseño optimizado presenta una disminución de 5 (K) comparado con un diseño definido por un usuario que considera los mismos materiales. También se estudia el efecto de la pintura sobre la superficie del CubeSat, donde se observa que el diseño optimizado reduce la temperatura en 8 (K) comparado con un satélite sin pintar (paredes de aluminio expuestas al ambiente).

Asteroid missions are gaining interest from the scientific community and many new missions are planned. The Didymos binary asteroid is a Near-Earth Object and the target of the Asteroid Impact and Deflection Assessment (AIDA). This joint mission, developed by NASA and ESA, brings the possibility to build one of the first CubeSats for deep space missions: the ASPECT satellite. Navigation systems of a deep space satellite di er greatly from the common planetary missions. Orbital environment close to an asteroid requires a case-by-case analysis. In order to develop the Attitude Determination Control System (ADCS) for the mission, one needs detailed information about orbital disturbances in the vicinity of the asteroid. This work focuses on the development of a simulator that characterises the orbital disturbances a ecting the ASPECT satellite in the space environment near the Didymos asteroid. In this work, a model of orbital conditions and disturbances near the Didymos system was defined. The model integrates several classical and modern models of spacecraft motion and disturbance. An existing Low Earth Orbit (LEO) simulator was modified and updated accordingly to the ASPECT mission scenario. The developed simulator can be used to analyse the disturbances to be counteracted by the ADCS of the ASPECT satellite. The objective of the study was to quantify the e ect of both non-gravitational and gravitational disturbances. The simulator was used to analyse di erent orbit scenarios related to the period of the mission and to the relative distance between the spacecraft and the asteroid system. In every scenario, the solar radiation pressure was found to be the strongest of the disturbance forces. With the developed simulator, suitable spacecraft configurations and control systems can be chosen to mitigate the e ect of the disturbances on the attitude and orbit of the ASPECT satellite.

Attitude control is an essential subsystem of most spacecraft buses, therefore attitude determination plays a very important role as it is the feedback system for any closed-loop attitude control system. Of all attitude determination sensors star trackers are usually the most accurate ones. Unfortunately, the star trackers usually used on classical, large spacecrafts are too big, heavy and power hungry. For pico-satellites, which can only carry a limited amount of volume and mass and provide only limited power, these sensors obviously cannot be used. Consequently, miniaturized star trackers have been developed in recent years, but so far the available star trackers are not sufficiently miniaturized to be feasible for the use on pico-satellites, including STELLA, a miniaturized star tracker developed at the University of Würzburg. Therefore, further miniaturization is necessary, which is why the University of Würzburg is active on the research of star trackers for small satellites. A first prototype for a new star tracker for pico-satellites, called PicoStar, has been developed in the scope of this thesis. Using a simpler system design and new image sensors, its volume could be reduced by two-thirds and the mass by about half compared to STELLA. The expected performance is kept reasonably. There is still room for further reduction of the power consumption, as it is currently up to 30% higher than required. As this Master thesis focuses on the implementation of the embedded system and the optimization of the software of the star tracker, the prototype is not finalized. So far, the star tracker algorithm has been implemented and the attitude determination is running. First test results have shown that the next steps in the PicoStar development, among other things, have to be further calibration and testing. / Validerat; 20151109 (global_studentproject_submitter)

Ingeniero Civil Eléctrico / Esta memoria se enmarca en el análisis de algoritmos de Súper Resolución (SR). Estos algoritmos recibirían imágenes tomadas por un futuro satélite Cubesat desde la órbita terrestre, obteniéndose una imagen con mayor resolución. El principal objetivo de esta memoria consiste en estudiar el coste en recursos, rendimiento y otras consideraciones de estos algoritmos. Así se podrá determinar cuál algoritmo de SR de los considerados pueda ser utilizado con ventajas en una plataforma Cubesat. Para lograr este objetivo se utiliza las experiencias de operación del primer Cubesat Chileno (SUCHAI 1). Usando esta información se ha podido determinarse cuáles serían los requerimientos que se deben cumplir en los futuros SUCHAI de manera de disponer de la capacidad de obtener imágenes en SR. Se realizaron evaluaciones de los algoritmos de SR implementándose una interfaz que permitió comparar visualmente, correr reiteradamente (cambiando sus parámetros) los algoritmos en una plataforma Matlab, de manera de evaluar sus prestaciones y obtener una estimación de la mejora en resolución resultante de la imagen final, el tiempo empleado en su ejecución y la memoria utiliza. Como complemento se utilizó un método llamado Mapas de Calor que permite construir comparaciones de un dataset (se dispusieron de 100 imágenes y se usó una métrica llamada Peak Signal to Noise Ratio -PSNR-) construido Ad hoc. Con los datos obtenidos se sugieren escenarios sobre los cuales utilizar los algoritmos a ser ejecutado a bordo del satélite (considerando sus características) y se postula que para obtener mejores prestaciones de SR lo práctico es bajar las imágenes y ejecutar en tierra (lo cual probablemente requerirá mucha más capacidad de cálculo). Se ven algunos pro y contra de la SR usando Deep Learning y se menciona como posible ampliación del campo de la SR a la combinación con Mapas Digitales de Elevación (DEM). / Este trabajo ha sido parcialmente financiado por Proyecto ANILLO ACT 1505

Après le lancement du premier satellite artificiel en 1957, l'évolution de diverses technologies a favorisé la miniaturisation des satellites. En 1999, le développement des nano-satellites modulaires appelés CubeSats, qui ont la forme d'un cube d'un décimètre de côté et une masse de 1 kg à 10 kg, a été initié par un effort commun de l'Université polytechnique de Californie et de l'Université de Stanford. Depuis lors, grâce à l’utilisation de composants électroniques standards à faible coût, les CubeSats se sont largement répandus.Au cours des dernières années, le nombre de CubeSats lancés a régulièrement augmenté, mais moins de la moitié des missions ont atteint leurs objectifs. L'analyse des défaillances des CubeSats montre que la cause la plus évidente est le manque d’essais adéquats des composants du système ou du système au complet. Parmi les tâches particulièrement difficiles, on compte les essais « hardware-in-the-loop » (HIL) du système de contrôle d'attitude et d'orbite (SCAO) d’un CubeSat. Un système dédié à ces essais doit permettre des simulations fiables de l'environnement spatial et des mouvements réalistes des CubeSats. La façon la plus appropriée d’obtenir de telles conditions d’essai repose sur l’utilisation d’un coussin d'air. Toutefois, les mouvements du satellite sont alors contraints par les limites géométriques, qui sont inhérentes aux coussins d'air. De plus, après 15 années de développements de CubeSats, la liste des systèmes proposés pour tester leur SCAO reste très limitée.Aussi, cette thèse est consacrée à l’étude et à la conception d’un système robotique innovant pour des essais HIL du SCAO d’un CubeSat. La nouveauté principale du système d'essai proposé est l’usage de quatre coussins d'air au lieu d'un seul et l’emploi d’un robot manipulateur. Ce système doit permettre des mouvements non contraints du CubeSat. Outre la conception du système d'essai, cette thèse porte sur les questions liées: (i) à la détermination de l'orientation d’un CubeSat au moyen de mesures sans contact; (ii) au comportement de l’assemblage des coussins d'air; (iii) à l'équilibrage des masses du système.Afin de vérifier la faisabilité de la conception proposée, un prototype du système d'essai a été développé et testé. Plusieurs modifications destinées à en simplifier la structure et à réduire le temps de fabrication ont été effectuées. Un robot Adept Viper s650 est notamment utilisé à la place d'un mécanisme sphérique spécifiquement conçu. Une stratégie de commande est proposée dans le but d’assurer un mouvement adéquat du robot qui doit suivre les rotations du CubeSat. Finalement, les résultats obtenus sont présentés et une évaluation globale du système d'essai est discutée. / After the launch of the first artificial Earth satellite in 1957, the evolution of various technologies has fostered the miniaturization of satellites. In 1999, the development of standardized modular satellites with masses limited to a few kilograms, called CubeSats, was initiated by a joint effort of California Polytechnic State University and Stanford University. Since then, CubeSats became a widespread and significant trend, due to a number of available off-the-shelf low cost components.In last years, the number of launched CubeSats constantly grows, but less than half of all CubeSat missions achieved their goals (either partly or completely). The analysis of these failures shows that the most evident cause is a lack of proper component-level and system-level CubeSat testing. An especially challenging task is Hardware-In-the-Loop (HIL) tests of the Attitude Determination and Control System (ADCS). A system devoted to these tests shall offer reliable simulations of the space environment and allow realistic CubeSat motions. The most relevant approach to provide a satellite with such test conditions consists in using air bearing platforms. However, the possible satellite motions are strictly constrained because of geometrical limitations, which are inherent in the air bearing platforms. Despite 15 years of CubeSat history, the list of the air bearing platforms suitable for CubeSat ADCS test is very limited.This thesis is devoted to the design and development of an air bearing testbed for CubeSat ADCS HIL testing. The main novelty of the proposed testbed design consists in using four air bearings instead of one and in utilizing a robotic arm, which allows potentially unconstrained CubeSat motions. Besides the testbed design principle, this thesis deals with the related issues of the determination of the CubeSat orientation by means of contactless measurements, and of the behavior of the air bearings, as well as with the need of a mass balancing method.In order to verify the feasibility of the proposed design, a prototype of the testbed is developed and tested. Several modifications aimed at simplifying the structure and at shortening the fabrication timeline have been made. For this reason, the Adept Viper s650 robot is involved in place of a custom-designed 4DoF robotic arm. A control strategy is proposed in order to provide the robot with a proper motion to follow the CubeSat orientation. Finally, the obtained results are presented and the overall assessment of the proposed testbed is put into perspective.

This dissertation addresses rigid body attitude control with piecewise sinusoidal signals. We consider rigid-body attitude kinematics on SO(3) with a class of sinusoidal inputs. We present a new closed-form solution of the rotation matrix kinematics. The solution is analyzed and used to prove controllability. We then present kinematic-level orientation-feedback controllers for setpoint tracking and command following. Next, we extend the sinusoidal kinematic-level control to the dynamic level. As a representative dynamic system, we consider a CubeSat with vibrating momentum actuators that are driven by small $\epsilon$-amplitude piecewise sinusoidal internal torques. The CubeSat kinetics are derived using Newton-Euler's equations of motion. We assume there is no external forcing and the system conserves zero angular momentum. A second-order approximation of the CubeSat rotational motion on SO(3) is derived and used to derive a setpoint tracking controller that yields order O(ε2) closed-loop error. Numerical simulations are presented to demonstrate the performance of the controls. We also examine the effect of the external damping on the CubeSat kinetics. In addition, we investigate the feasibility of the piecewise sinusoidal control techniques using an experimental CubeSat system. We present the design of the CubeSat mechanical system, the control system hardware, and the attitude control software. Then, we present and discuss the experiment results of yaw motion control. Furthermore, we experimentally validate the analysis of the external damping effect on the CubeSat kinetics.

Testing is an integral part of a satellite’s development, requirements verification and risk mitigation efforts. A robust test program serves to verify construction, integration and assembly workmanship, ensures component, subsystem and system level functionality and reduces risk of mission or capability loss on orbit. The objective of this thesis was to develop a detailed test program for pico-satellites with a focus on the Cal Poly CubeSat architecture. The test program established a testing baseline from which other programs or users could tailor to meet their needs. Inclusive of the test program was a detailed decomposition of discrete and derived test requirements compiled from the CubeSat and Launch Vehicle communities, military guidelines, and industry standards. The test requirements were integrated into a methodical, efficient and risk adverse test flow for verification.

The Naval Postgraduate School is currently developing their first CubeSat, the Solar Cell Array Tester CubeSat, or NPS-SCAT. Launching a CubeSat, such as NPS-SCAT, requires environmental testing to ensure not only the success of the mission, but also the safety of other CubeSats housed in the same deployer. This thesis will address the development of CubeSat vibration testing methodology at NPS, including subsystem testing, engineering unit qualification, and flight unit testing. In addition, the new Cal Poly CubeSat Test POD Mk III will be introduced and evaluated based upon comparison with the Poly Picosatellite Orbital Deployer (P-POD). Using examples from the development of NPS-SCAT and test data from Cal Poly’s Test POD Mk III and P-POD, the current CubeSat testing methodology will be verified and an improved method for NPS CubeSat subsystem testing will be presented.