An autonomous mission manager is being developed for use on CubeSats to perform proximity operations with other vehicles. The mission manager software is designed to run in real-time on a microprocessor used on a CubeSat. A simulation tool was developed that provides orbital dynamics and sensor measurements to test the mission manager software. A scenario was developed to demonstrate the control of a spacecraft from 1 km to 1 m to a target vehicle. Two small satellites were simulated in near-circular orbits around Earth at an approximate 400 km altitude. Each satellite is incorporated with simulated sensors and a Kalman filter. The simulation tool includes models for accelerometers and Global Positioning System receivers. Noise corruption is added to the modeled sensors to simulate imperfect knowledge. The simulation environment is capable of modeling Earth as a spherical or non-spherical body with spherical gravitational harmonics. Simulation parameters, such as the vehicle's initial states, Earth gravity model, and sensor noise are easily changed without recompiling the program through a simulation input file. / text

The project is an orbital design study of a proposed CubeSat at Mississippi State University. The launch date is not specified. As for the mission, it is defined as forest fire detection. CubeSats are small satellites that are 10 x 10 x 10 cm in dimension and has a mass no more than 1 kg. They are currently used in different applications in many countries as an easy access to space. The analyses of this project have been carried out using a commercial software package, System Tool Kit (STK), developed by Analytical Graphics, Incorporated. This software provides a tool for performing simulations required for determining the orbit or the trajectory for satellites. In addition, a perturbation orbital study has been conducted and different propagators have been tested.

The entire contiguous United States experienced a solar eclipse on August 21st, 2017 which passed from the Pacific to the Atlantic Coasts. The path of totality crossed 14 states while other states had partial eclipse. Due to the rarity of this event, it was known as “The Great American Eclipse” and NASA collaborated with 52 universities across the United States to launch weather balloon payloads to record this impactful event. Although Montana State University designed a workshop for all universities involved in order to assist those not experienced in the area, Mississippi State University decided to design our own payload. Our system was designed in order to meet the standards of a 2U CubeSat. One key aspect of our payload is that it is entirely made from 3D printed parts with over 100 prototype parts made over the length of two years. Instead of buying an off the shelf flight computer, we designed and built a custom Hexa-Processor Computer Board which gave us flexibility with the computation needs. A turret was also developed that housed two cameras and could spin 360 degrees, allowing it to counter act the rotations of the payload in order to obtain a stabilized image. The payload was launched in Kentucky and was a successful flight without any damages to the payload.

This research demonstrates the potential of novel technology for space-based remote sensing of the topside ionosphere-plasmasphere, supported by ionospheric imaging, which can augment and enhance our current understanding of the Earth’s plasmasphere. The research was conducted in two phases. The first was the development of a technology demonstrator ‘TOPCAT’ that installed a dual-frequency GPS receiver dedicated for topside ionosphere-plasmasphere imaging into a Low Earth Orbit (LEO). The novelties of TOPCAT were that it was designed from commercial-off-the-shelf (COTS) components and was installed on-board the CubeSat ‘UKube-1’, greatly reducing development and launch costs of the instrument. The successful launch of TOPCAT for space-borne remote sensing of the topside ionosphere and plasmasphere could provide the necessary proof of concept for the installation of a constellation of CubeSats – a possible next phase that may be implemented in the future. Thus, in its first stage, the thesis discusses the development of TOPCAT, together with design challenges encountered from constraints imposed by CubeSat technology. The discussion also includes the series of qualification tests performed to successfully qualify TOPCAT as a space-worthy payload design that can remotely image regions beyond the ionosphere. The second phase of research was the validation of the Multi-Instrument Data Analysis System (MIDAS) for the topside ionosphere and plasmasphere. A tomography algorithm originally developed for the ionosphere, MIDAS uses total electron content (TEC) measurements from differential phase of GPS signals, and inverts them to derive the electron density of the region. The thesis investigates the extension of MIDAS to image regions beyond the ionosphere by validating the algorithm for the topside ionosphere and plasmasphere. The process was carried out by first reconstructing a simulation by Gallagher et al. [1988] to verify the quality of the images. This was followed by the use of real GPS phase data from the COSMIC constellation to reconstruct the topside ionosphere-plasmasphere, and the qualitative comparison of the images with previous independent observations obtained through COSMIC and Jason-1 missions. Results showed that MIDAS can successfully reconstruct the undisturbed (quiet) topside ionosphere-plasmasphere using COSMIC data. However, imaging the storm-time topside ionosphere-plasmasphere requires better data coverage (i.e. more receivers) as the resolution offered by COSMIC was not sufficient to reconstruct fast-evolving structures – thereby emphasising the need for more data sources providing high resolution global coverage, such as a constellation of CubeSats with LEO-based GPS receivers.

A control system is designed and validated for a 3-unit CubeSat science mission. Utilizing three magnetorquers and one reaction wheel the system achieves a 6 rotation per minute spin rate and orbit normal pointing vector of the long axis of a 3-unit CubeSat. The design is validated across an evolution of scenarios, from idealized to flight-like with expected bias and noise terms added in. Estimated mass imbalances and the limitations of the power system driving the magnetorquers force the final system design to use only magnetorquers. Considerations are also taken for the science instrument to limit the interference of magnetorquers. The overall satellite design and software implementation are also briefly discussed. / text

A control system is designed and validated for a 3-unit CubeSat science mission. Utilizing three magnetorquers and one reaction wheel the system achieves a 6 rotation per minute spin rate and orbit normal pointing vector of the long axis of a 3-unit CubeSat. The design is validated across an evolution of scenarios, from idealized to flight-like with expected bias and noise terms added in. Estimated mass imbalances and the limitations of the power system driving the magnetorquers force the final system design to use only magnetorquers. Considerations are also taken for the science instrument to limit the interference of magnetorquers. The overall satellite design and software implementation are also briefly discussed. / text

Ingeniero Civil Eléctrico / Esta memoria se enmarca dentro del proyecto SUCHAI, el que busca diseñar, construir, integrar, programar, lanzar y operar un nano-satélite de estándar Cubesat con estudiantes de pregrado. En este contexto el objetivo de la memoria es formalizar, analizar y presentar: El proceso de diseño seguido hasta obtener los requerimientos. La arquitectura y operación de este. Y finalmente la verificación a las que fue sometido para comprobar su funcionamiento. En la primera etapa se muestra cómo a partir de la misión y restricciones del proyecto se extraen los requerimientos y se reúnen en la llamada Matriz de Trazabilidad de Requerimientos (Requirements Traceability Verification Matrix - RTVM en inglés). En la segunda parte se detalla la arquitectura y funcionamiento del bus-SUCHAI, haciendo para ello una analogía con el modelo OSI. En la tercera y final se somete al satélite a pruebas parciales en el laboratorio, y a una prueba total en una cámara de alto vacío, buscando comprobar si se cumplen cada uno de los requisitos, usando para ello nuevamente la mencionada matriz. Como conclusión de las tres etapas, se concluye que el SUCHAI aunque no satisface aún todos los requerimientos propuestos originalmente si lo hace con los más importantes. Lo que permite afirmar que se ha desarrollado un bus lo suficientemente modular, fiable y capaz para ser lanzado y operado en el espacio. Por último, tres aportes del trabajo de memoria son destacados, las propuestas de mejoras para el bus-SUCHAI, la aplicación de herramientas de Ingeniería de Sistemas como la RTVM en un proyecto Cubesat y la propuesta de un proceso de diseño para desarrollos similares.

Die Raumfahrt ist eine der konservativsten Industriebranchen. Neue Entwicklungen von Komponenten und Systemen beruhen auf existierenden Standards und eigene Erfahrungen der Entwickler. Die Systeme sollen in einem vorgegebenen engen Zeitrahmen projektiert, in sehr kleiner Stückzahl gefertigt und schließlich aufwendig qualifiziert werden. Erfahrungsgemäß reicht die Zeit für Entwicklungsiterationen und weitgehende Perfektionierung des Systems oft nicht aus. Fertige Sensoren, Subsysteme und Systeme sind Unikate, die nur für eine bestimme Funktion und in manchen Fällen sogar nur für bestimmte Missionen konzipiert sind. Eine Neuentwicklung solcher Komponenten ist extrem teuer und risikobehaftet. Deswegen werden flugerprobte Systeme ohne Änderungen und Optimierung mehrere Jahre eingesetzt, ohne Technologiefortschritte zu berücksichtigen. Aufgrund des enormen finanziellen Aufwandes und der Trägheit ist die konventionelle Vorgehensweise in der Entwicklung nicht direkt auf Kleinsatelliten übertragbar. Eine dynamische Entwicklung im Low Cost Bereich benötigt eine universale und für unterschiedliche Anwendungsbereiche leicht modifizierbare Strategie. Diese Strategie soll nicht nur flexibel sein, sondern auch zu einer möglichst optimalen und effizienten Hardwarelösung führen. Diese Arbeit stellt ein Software-Tool für eine zeit- und kosteneffiziente Entwicklung von Sternsensoren für Kleinsatelliten vor. Um eine maximale Leistung des Komplettsystems zu erreichen, soll der Sensor die Anforderungen und Randbedingungen vorgegebener Anwendungen erfüllen und darüber hinaus für diese Anwendungen optimiert sein. Wegen der komplexen Zusammenhänge zwischen den Parametern optischer Sensorsysteme ist keine „straightforward" Lösung des Problems möglich. Nur durch den Einsatz computerbasierter Optimierungsverfahren kann schnell und effizient ein bestmögliches Systemkonzept für die gegebenen Randbedingungen ausgearbeitet werden. / Aerospace is one of the most conservative industries. New developments of components and systems are based on existing standards and experience of developers. The systems should be projected in a given tight time frame, manufactured in very small quantities and finally qualified in a costly way. Experience shows that there is often insufficient time for development iterations and extensive perfection of the system. Finished sensors, subsystems and systems are unique, designed only for a specific function and in some cases even only for specific missions. New development of such components is extremely expensive and risky. For this reason, flight-proven systems are used for several years without modifications or optimization, and without taking technological advances into account. Due to the enormous financial effort and lethargy, the common approach to development is not directly applicable to small satellites. Dynamic development in the low-cost sector requires a universal strategy that can be easily modified for different applications. This strategy should not only be flexible, but also lead to the most optimal and efficient hardware solution. This work presents a software tool for a time and cost efficient development of star sensors for small satellites. In order to achieve maximal performance of the complete system, the sensor should fulfil the requirements and constraints of specified applications and, moreover, be optimized for these applications. Due to the complex interrelationships between the parameters of optical sensor systems, no straight forward solution of the problem is possible. Only by using computer based optimization methods, a best possible system concept for the given boundary conditions can be worked out quickly and efficiently.

We present a preliminary design for a neutron detection system that is compact, lightweight, and low power consuming, utilizing the CubeSat platform making it suitable for space-based applications. This is made possible using the scintillating crystal lithium indium diselenide ((LiInSe2)-Li-6), the first crystal to include Li-6 in the crystalline structure, and a silicon avalanche photodiode. The schematics of this instrument are presented as well as the response of the instrument to initial testing under alpha radiation. A principal aim of this work is to demonstrate the feasibility of such a neutron detection system within a CubeSat platform. The entire end-to-end system presented here is 10 x 10 x 15 cm(3), weighs 670 g, and requires 5 V direct current at 3 W. (C) 2016 Society of Photo-Optical Instrumentation Engineers (SPIE)

In the present paper, which was written in collaboration with the Observatory of Paris, thermal models of two CubeSat missions were created. The first goal of this work was to create a nodal simulation of the PicSat satellite to verify the survivability of the system within the extreme space environment. In a second step suitable countermeasures were suggested, if parts of the satellite exceeded a critical temperature limit. Additionally, the impacts of three failure modes were investigated. The second goal was to perform thermal pre-studies of different satellite configurations of the CIRCUS satellite based on the model of PicSat. The simulation conducted in this work showed that PicSat fulfils its thermal requirements, except for three components. Out of the failure modes considered in this study, only an error of the attitude determination and control systems (ADCS) showed a critical impact on the thermal state of PicSat. As the outcome of the pre-studies of CIRCUS, two of the considered configurations were recommended for further analysis. The results of this work will contribute to the future development of the PicSat and CIRCUS missions. / Dans le cadre de ce mémoire, qui a été rédigé en collaboration avec l'Observatoire de Paris, les modèles thermiques de deux missions CubeSat ont été créés. Le premier objectif de ce projet était de d\'velopper une simulation nodale du satellite PicSat afin de vérifier la capacité de survie du système dans les conditions hostiles de l'espace. Dans un second temps, des mesures correctives ont été suggérées si la température de certaines sections du satellite venait à excéder une limite critique. En outre, les impacts de trois défaillances ont été analysés. Le second objectif était de réaliser des pré-études thermiques pour différentes configurations du satellite CIRCUS basées sur le modèle de PicSat. Les simulations réalisées dans le cadre de ce projet ont montré que PicSat remplit ses exigences thermiques, à l'exception de trois composants. Des différentes défaillances étudiées, seule une erreur du système de détermination et de contrôle d'attitude (Attitude Determination and Control System ADCS) a montré un impact critique sur l'état thermique de PicSat. A l'issue des pré-études de CIRCUS, deux des configurations étudiées ont été recommandées pour être l'objet d'une analyse plus avancée. Les résultats de ce projet contribueront au développement à venir des missions des satellites PicSat et CIRCUS.