Ingeniero Civil Eléctrico / El estándar de nanosatélites Cubesat fue pensado para facilitar el desarrollo de pequeños proyectos espaciales con fines científicos y educacionales, a un bajo costo y en cortos periodos de tiempo. Siguiendo esta línea, la Facultad de Ciencias Físicas y Matemáticas de la Uni- versidad de Chile ha impulsado el proyecto SUCHAI, que consiste en implementar, poner en órbita y operar el primer satélite desarrollado por una universidad del país. El computador a bordo de la aeronave, que consiste un sistema embebido de limitada capacidad de cómputo, escasa memoria y bajo consumo de energía, debe ejecutar el software de vuelo que controlará sus operaciones una vez en órbita. El objetivo de este trabajo es el diseño e implementación de este software para el satélite SUCHAI, como una solución confiable, flexible y extensible que sea la base para futuras misiones aeroespaciales. El diseño del software consiste en una estructura de tres capas, que consigue dividir el problema convenientemente. La de más bajo nivel considera los controladores de hardware, la capa intermedia alberga al sistema operativo, y la de nivel superior, contiene los detalles de la aplicación requerida específicamente para este sistema. Para la arquitectura de la capa de aplicación, se estudia y aplica el concepto de patrón de diseño, en específico, se realiza una adaptación de command pattern. De esta manera, el satélite se concibe como un ejecutor de comandos genéricos y se obtiene una solución mantenible, modificable y extensible en el tiempo, mediante la programación de los comandos concretos que sean requeridos. La implementación se realiza sobre un PIC24F y considera controladores para los periféricos I2C, RS232 y SPI, así como para los subsistemas de radiocomunicaciones y energía. Se decide utilizar el sistema operativo FreeRTOS, como capa intermedia, lo que permite contar con el procesamiento concurrente de tareas, herramientas de temporización y sincronización. Se ha puesto especial énfasis en la implementación de la arquitectura planteada para la capa de aplicación, consiguiendo un software capaz de ejecutar una serie de comandos, programados para cumplir los requerimientos operacionales del proyecto, lo cual representa el método principal para extender sus funcionalidades y adecuarse a futuras misiones. Para probar y verificar el sistema desarrollado, se ha utilizado la técnica denominada hardware on the loop simulation. Se han obteniendo datos de funcionamiento, bajo condiciones de operación hipotéticas, a través del registro generado por la consola serial. Con esto se verifican los requerimientos operacionales de la misión, con resultados exitosos, obteniendo el sistema base y funcional del satélite. Como trabajo futuro, se utilizará este software para integrar el resto de los sistemas del satélite SUCHAI, demostrando su capacidad de adaptación y extensión, en un paso previo a la prueba final: funcionar adecuadamente en el espacio exterior.

The CubeSat standard sprang from the desire to create a satellite standard that would open the doors for universities and other lower budget research institutions by making it more feasible to get their work into space. Since then, many other institutions and industries have been adopting variations on the standard for their own use. As more people are seeking out to use the CubeSat standard as their main bus, the standards and practices of the community have grown and expanded and with this growth, new challenges have been created. One such challenge is the bandwidth limitation in the RF-downlink. When carrying payloads requiring what might seem to be a relatively small (science data) bandwidth requirement (on the order of thousands of bps), the RF-link to ground is overloaded. Many approaches in the past have been put forth to help alleviate this issue, unfortunately, none have been fully adopted. This paper presents a solution that takes advantage of new technology yet to be fully exploited in space applications. The key to the solution lies in removing the bandwidth requirements by enabling onboard post-data processing and compression. In order to achieve the high computational needs, while minimizing power consumption, a Xilinx Zynq-7000 SoC is used, creating a highly-programmable, open integration device. This report outlines the design, fabrication and testing of this solution. The completion of the Zynq Processing System CubeSat Science Instrument Interface Electronics Board (or ZPS-Board), ultimately demonstrates the feasibility of this solution. Additionally, this research is funded by NASA’s JPL, with secondary motives for the creating of a space application Zynq-7000 SoC based product. Upon successful completion of the ZPS-Board, the product creates a platform for JPL to perform environmental testing in order to study the effects and performance characteristics of the Zynq in space applications.

Nanosatellites, such as Cubesats, are a rapidly growing sector of the space industry. Their popularity stems from their low development cost, short development cycle, and the widespread availability of COTS subsystems. Budget-conscious spacecraft designers are working to expand the range of missions that can be accomplished with nanosatellites, and a key area of development fueling this expansion is the creation of micropropulsion systems. One such system, originally developed at the Australian National University (ANU), is an electrothermal plasma thruster known as Pocket Rocket (PR). This device heats neutral propellant gas by exposing it to a Capacitively Coupled Plasma (CCP), then expels the heated gas to produce thrust. Significant work has gone towards understanding how PR creates and sustains a plasma and how this plasma heats the neutral gas. However, no research has been published on varying in the device's geometry. This thesis aims to observe how the size of the RF electrode affects PR operation, and to determine if it can be adjusted to improve performance. To this end, a thruster has been built which allows the geometry of the RF electrode to be easily varied. Measurements of the plasma density at the exit of this thruster with different sizes of electrode were then used to validate a Computational Fluid Dynamics (CFD) model capable of approximately reproducing experimental measurements from both this study and from the ANU team. From this CFD, the number of argon ions in the thruster was found for each geometry, since collisions between argon ions and neutrals are primarily responsible for the heating observed in the thruster. A geometry using a 10.5 mm electrode was observed to produce a 23% increase in the quantity of ions produced compared to the baseline 5 mm electrode size, and a 3.5 mm electrode appears to produce 88% more ions.

Ever since the revolutionary CubeSat form factor took hold in the Aerospace industry, there has been a desire to send them further and further into space. This thesis introduces an optimization approach to deployment that explores new possibilities of interplanetary CubeSats. In this approach there are three categories of objective functions that are defined by the type of trajectory of a “primary” spacecraft, which carries the CubeSat deployer. These categories are flyby, orbiter, and lander. For each category the objective function starts with four design variables. These are the ΔV of the deployer broken up into three component directions and the true anomaly at the time of deployment. The method then calculates the mission specific objective to be minimized and uses Matlab®’s built in gradient-based optimizer, fmincon. The results show that in the flyby category, the CubeSat has a significantly different turning angle than the primary. The CubeSat can even flyby on the opposite side of the planet. In the orbiter case it is shown that the method works by testing it with two objective functions, the difference in inclination and the difference in eccentricity between the primary and the CubeSat. It is shown that the inclination can be changed by 0.1314° and the eccentricity can be changed by 0.0033. These values, although low in magnitude, are an order of magnitude greater than non-optimal deployment scenarios. Still, another optimization method is introduced to find out how much extra ΔV the CubeSat would need to reach a desired change. This shows that with just an extra 75 m/s of ΔV, the CubeSat can change its orbit by 5°. This could come from either a propulsion system or a modified deployer. The final category, lander, used the flight path angle when entering the atmosphere as an objective. The method shows that flight path angle can be changed by 2.6°. Overall, these examples have proven that the method can find optimal solutions to CubeSat deployment scenarios at other planets.

Cílem práce je návrh univerzální řídicí jednotky pro CubeSat založené na obvodu FPGA. Taková jednotka doposud nebyla komerčně dostupná a navržená jednotka má tak dobrý potenciál zaplnit příslušné místo na trhu komponent pro CubeSat. Celá jednotka je navržena z komerčně dostupných komponent. Návrh jednotky je proveden tak, aby umožnil její funkci ve vesmírném prostředí. Stav konfigurace FPGA je pravidelně kontrolován a v případě zjištěné chyby dochází automaticky k rekonfiguraci FPGA a návratu jednotky do výchozího stavu. Jednotka obsahuje sadu senzorů, které monitorují její stav a v případě potřeby je možné na základě jejich výstupů provést opatření z hlediska ochrany funkce jednotky. Dvě paměti MRAM umožňují uložení tovární a uživatelské konfigurace FPGA, mezi kterými dochází k automatickému přepnutí na základě korektnosti uživatelské konfigurace.

Current CubeSat communication technology limits the amount of time, and number of accesses with ground stations. It has been proposed to use a constellation of CubeSats to improve relay performance and increase the number of accesses between a CubeSat and ground stations. By using the spatial and temporal analysis features of STK, coupled with the STK/Matlab interface a robust tool was created to analyze the performance of CubeSat constellations based on a store-and-forward communications model which is not currently supported by the STK Engine. Utilizing the Connect messaging format through a socket connection on the local machine, a Matlab graphical user interface, called SATCAT, was constructed in order to provide a user with the ability to control many aspects of the STK Engine externally. A function was created to use three Time Ordered Access (TOA) reports from STK to determine how long it would take for data to be relayed from a target to a ground station through a constellation of CubeSats. Three sample scenarios were created to demonstrate the use and performance analysis capabilities of SATCAT. The performance of a single CubeSat was analyzed and compared to the performance of a three CubeSat constellation and a thirty-seven CubeSat constellation. It was shown that a constellation of three CubeSats decreased the average relay time from 328 minutes to 149 minutes and a constellation of thirty-seven CubeSats further reduced the average relay time to only 3 minutes. While decreasing the average relay time, the constellation of three CubeSats also increased the number of accesses over a twenty-four hour period from 6 to 36, and the constellation of thirty-seven CubeSats allowed for 564 accesses.

The Poly-Picosatellite Orbital Deployer (P-POD) has undergone a series of revisions over the years. The latest revision, described in this Master’s Thesis, incorporates new capabilities like EMI shielding, an inert gas purge system, and an electrical interface to the CubeSats after they are integrated into the P-POD. Additionally, some mass reduction modifications are made to the P-POD, while its overall strength is increased. The P-POD inert gas purge system successfully flew, on a previous revision P-POD. The P-POD components are analyzed to a set of dynamic loads for qualification, and successfully undergoes random vibration qualification testing. The P-POD encounters some problems in thermal vacuum cycling qualification and EMI testing, but there is evidence that the issues can be mitigated. A path forward is laid out to complete both sets of testing.

Spacecraft health monitoring is essential to ensure that a spacecraft is operating properly and has no anomalies that could jeopardize its mission. Many of the current methods of monitoring system health are difficult to use as the complexity of spacecraft increase, and are in many cases impractical on CubeSat satellites which have strict size and resource limitations. To overcome these problems, new data-driven techniques such as Inductive Monitoring System (IMS), use data mining and machine learning on archived system telemetry to create models that characterize nominal system behavior. The models that IMS creates are in the form of clusters that capture the relationship between a set of sensors in time series data. Each of these clusters define a nominal operating state of the satellite and the range of sensor values that represent it. These characterizations can then be autonomously compared against real-time telemetry on-board the spacecraft to determine if the spacecraft is operating nominally. This thesis presents an adaption of IMS to create a spacecraft health monitoring system for CubeSat missions developed by the PolySat lab. This system is integrated into PolySat's flight software and provides real time health monitoring of the spacecraft during its mission. Any anomalies detected are reported and further analysis can be done to determine the cause. The system can also be used for the analysis of archived events. The IMS algorithms used by the system were validated, and ground testing was done to determine the performance, reliability, and accuracy of the system. The system was successful in the detection and identification of known anomalies in archived flight telemetry from the IPEX mission. In addition, real-time monitoring performed on the satellite yielded great results that give us confidence in the use of this system in all future missions.

With an increased demand in Cube Satellite (CubeSat) development for low cost science and exploration missions, a push for the development of micro-propulsion technology has emerged, which seeks to increase CubeSat capabilities for novel mission concepts. One type of micro-propulsion system currently under development, known as Pocket Rocket, is an electrothermal plasma micro-thruster. Pocket Rocket uses a capacitively coupled plasma, generated by radio-frequency, in order to provide neutral gas heating via ion-neutral collisions within a gas discharge tube. When compared to a cold-gas thruster of similar size, this gas heating mechanism allows Pocket Rocket to increase the exit thermal velocity of its gaseous propellant for increased thrust. Previous experimental work has only investigated use of the gas discharge tube's orifice for propellant expansion into vacuum. This thesis aims to answer if Pocket Rocket may see an increase in thrust with the addition of a micro-nozzle, placed at the end of the gas discharge tube. With the addition of a conical ε = 10, α = 30° micro-nozzle, performance increases of up to 6% during plasma operation, and 25% during cold gas operation, have been observed. Propellant heating has also been observed to increase by up to 60 K within the gas discharge tube.

Ingeniero Civil Eléctrico / Las misiones de satélites tipo CubeSat actualmente no tienen planificado el periodo posterior a la vida operacional del satélite. Al ser lanzados cada vez en mayor cantidad, estas naves tienen el potencial de congestionar el espacio en órbita terrestre baja y presentar un peligro creciente de colisiones. Dentro de las estrategias para remover a estas naves de la órbita es la de descenso controlado o deorbitación realizado con propulsores eléctricos. Dentro de las tecnologías existentes de propulsores eléctricos se encuentra el electrospray que presenta una alternativa atractiva para su uso en satélites tipo CubeSat. En el presente trabajo se evalúa esta tecnología y su uso en la deorbitación de CubeSats de tres unidades, para considerar su desarrollo para misiones futuras. En la primera parte del presente trabajo se presenta un marco teórico que permite entender las problemáticas que presenta el diseño y construcción de sistemas para CubeSats. Luego se explora la propulsión espacial eléctrica para estudiar su impacto los cambios orbitales de una órbita terrestre baja. Finalmente se presenta el fenómeno basal de la tecnología electrospray para luego comprenderlo en términos de su desempeño como sistema de propulsión espacial eléctrica. Se construyó un experimento en base a un capilar para estudiar las variables de importancia expuestas en la primera parte de este trabajo. En primer lugar se determinó la región de operación del experimento donde se producía de forma estable el fenómeno basal del electrospray, cuantificando las variables de importancia. De esta región de estabilidad se calcularon los parámetros de desempeño propulsivo de empuje e impulso específico en base un modelo físico. Luego se realizan pruebas para estudiar el comportamiento del experimento ante el vacío presente fuera de la atmósfera terrestre. Una vez determinados los parámetros de desempeño propulsivo, se realizan cálculos para determinar el desempeño total de un propulsor práctico utilizando esta tecnología en los puntos de operación que maximizan cada parámetro respectivamente. Junto al desempeño, se estima el volumen utilizado y la cantidad de propelente factible. Finalmente se evaluó el desempeño del propulsor propuesto en términos de vida orbital mediante la simulación en el software STK. Estas simulaciones incluyen dos maniobras cuyos resultados son comparados a los resultados de una caída sin la asistencia de propulsión. Mediante las simulaciones se obtuvieron reducciones significativas a la vida orbital del satélite lo que el propulsor posee el desempeño suficiente para su uso en deorbitación de CubeSats de tres unidades.