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.

"Orbital flight of CubeSats in extremely Low Earth Orbit, defined here as an altitude between 150 – 250 km, has the potential to enable a wide range of missions in support of atmospheric measurements, national security, and natural resource monitoring. In this work, a mission study is presented to demonstrate the feasibility of using commercially available sensor and electric thruster technology to extend the orbital lifetime of a 3U CubeSat flying at an altitude of 210 km. The CubeSat consists of a 3U configuration and assumes the use of commercially available sensors, GPS, and electric power systems. The thruster is a de-rated version of a commercially available electrospray thruster operating at 2 W, 0.175 mN thrust, and an Isp of 500 s. The mission consists of two phases. In Phase I the CubeSat is deployed from the International Space Station orbit (414 km) and uses the thruster to de-orbit to the target altitude of 210 km. Phase II then begins during which the propulsion system is used to extend the mission lifetime until propellant is fully expended. A control algorithm based on maintaining a target orbital energy is presented in which simulated GPS updates are corrupted with measurement noise to simulate state data which would be available to the spacecraft computer. An Extended Kalman Filter is used to generate estimates of the orbital dynamic state between the 1 Hz GPS updates, allowing thruster control commands at a frequency of 10 Hz. For Phase I, operating at full thrust, the spacecraft requires 25.21 days to descend from 414 to 210 km, corresponding to a ΔV = 96.25 m/s and a propellant consumption of 77.8 g. Phase II, the primary mission phase, lasts for 57.83 days, corresponding to a ΔV = 119.15 m/s during which the remaining 94.2 g of propellant are consumed. "

Digital communications devices designed with application-specific integrated circuit (ASIC) technology suffer from one very significant limitation�the integrated circuits are not programmable. Therefore, deploying a new algorithm or an updated standard requires new hardware. Field-programmable gate arrays (FPGAs) solve this problem by introducing what is essentially reconfigurable hardware. Thus, digital communications devices designed on FPGAs are capable of accommodating multiple communications protocols without the need to deploy new hardware, and can support new protocols in a matter of seconds. In addition, FPGAs provide a means to update systems that are physical difficult to access. For these reasons, FPGAs provide us with an ideal platform for implementing adaptive communications algorithms. This thesis focuses on using FPGAs to implement an adaptive digital communications system. Using the Universal Software Radio Peripheral (USRP) as a base, this thesis aims to create a highly-adaptive, plug and play software-defined radio (SDR) that fits CubeSat form-factor satellites. Such a radio platform would enable CubeSat engineers to develop new satellites faster and with lower costs. This thesis presents a new system, the COSMIAC CubeSat SDR, that adapts the USRP platform to better suit the space and power limitations of a CubeSat.

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 guidance, navigation, and control (GN&C) module is being designed and fabricated as part of a series of CubeSats being built by the Satellite Design Laboratory at the University of Texas. A spacecraft attitude control simulation environment called StarBox was created in order to perform trade studies and conduct performance analysis for the GN&C module. Navigation and control algorithms were tested using StarBox and then implemented onto an embedded flight computer. These algorithms were then tested in a hardware-in-the-loop simulation. In addition, the feasibility of utilizing advanced constrained attitude control algorithms was investigated by focusing on implementation in flight software. A mechanical and electrical design for the GN&C module was completed. A prototype system was set up on a bench-top for integrated testing. The analysis indicates that the system will satisfy the requirements of several CubeSat missions, including the current missions at the University of Texas known as Bevo2 and ARMADILLO. / text

Η παρούσα διπλωματική εργασία πραγματεύεται την μελέτη και κατασκευή του συστήματος τροφοδοσίας του UPSat, ενός δορυφόρου τύπου Cubesat. Η εργασία αυτή εκπονήθηκε στο Εργαστήριο Ηλεκτρομηχανικής Μετατροπής Ενέργειας του Τμήματος Ηλεκτρολόγων Μηχανικών και Τεχνολογίας Υπολογιστών της Πολυτεχνικής Σχολής του Πανεπιστημίου Πατρών. Σκοπός είναι η μελέτη και η κατασκευή των βασικών διατάξεων του συστήματος τροφοδοσίας ώστε να καλύπτονται οι ενεργειακές απαιτήσεις όλων των υποσυστημάτων του δορυφόρου. Αρχικά μελετώνται οι συνθήκες που θα πρέπει να αντιμετωπίσει ο δορυφόρος κατά την αποστολή του. Παρουσιάζονται τα βασικότερα υποσυστήματα του δορυφόρου και οι διαθέσιμες επιλογές τροχιάς, ενώ με βάση στοιχεία προηγούμενων αποστολών επιλέγεται η τοπολογία του συστήματος και γίνεται ο ισολογισμός ισχύος. Στη συνέχεια γίνεται η ανάλυση των τροφοδοτικών που θα κατασκευαστούν, η επιλογή των κατάλληλων τοπολογιών και ο υπολογισμός των απαραίτητων στοιχείων των μετατροπέων. Ακολουθεί η προσομοίωση των κυκλωμάτων με κατάλληλο λογισμικό, το πρόγραμμα PSpice, ώστε τα αποτελέσματά τους να συγκριθούν με τα θεωρητικά αλλά και με αυτά της τελικής κατασκευής. Το επόμενο βήμα είναι η κατασκευή των κυκλωμάτων. Παρουσιάζονται τα τεχνικά χαρακτηριστικά των βασικών εξαρτημάτων και γίνεται η ανάλυση του προγράμματος του μικροελεγκτή για την παραγωγή παλμών. Επιπρόσθετα γίνεται αναφορά και στις διατάξεις μετρήσεων και ελέγχου του συστήματος. Τέλος δίνονται οι μετρήσεις και τα παλμογραφήματα που προέκυψαν από τις διατάξεις για την εξαγωγή συμπερασμάτων. / This diploma thesis deals with the design and construction of the Power Supply System of UPSat, a satellite based on the Cubesat standard. The work was conducted at the Laboratory of Electromechanical Energy Conversion of the Department of Electrical and Computer Engineering of Engineering School of University of Patras. The aim is to design and construct the main circuits of the Power Supply Unit, in order to cover all the energy required by the subsystems of the satellite. Initially, the conditions to be faced by the satellite during the mission are taken into consideration. The basic subsystems of the satellite and the orbit options are presented and based on data from previous missions, the topology of the whole system is selected. Then follows the theoretical analysis of the power supply unit to be built, choosing the appropriate topologies followed by the calculation of all the necessary components. The next step is to simulate all the circuits with the appropriate software, in this case the PSpice program, in order to compare the results with the theoretical but also with those of the final construction. Following, begins the construction of the circuits. Technical characteristics of the basic components are presented and also the analysis of the program of the microcontroller that generates a pulse for a mosfet. In addition, there is the analysis of the measurements and control circuitry. Finally oscilloscope graphs and measurements, occurred from the experiments transacted after the finalization of the construction, are adduced.

The CubeSat Program has provided access to space for many universities, private companies, and government institutions primarily due to the low cost of CubeSat satellite development. While these costs are orders of magnitude lower than similarly capable nano-satellite missions, they are still outside of the budgetary constraints of many potential developers including university and high school clubs. Using 3D printed plastics in the production of CubeSat structures and mechanisms presents a large cost savings opportunity that will allow these institutions to participate in the development of these satellites, expanding the educational and scientific impact of the CubeSat Program. Five rapid prototype plastics manufactured with four different 3D printing technologies were studied to determine their survivability when subjected to the required vibration testing and thermal bakeout that all CubeSats are must pass through before integration and launch. ASTM D638 Type V tensile bar samples of each plastic were procured and subjected to a thermal bakeout and tensile testing to determine the thermal and outgassing effects on their mechanical properties. This information was used to design a concept structure for use in a low budget CubeSat mission. Finite Element Analysis in Abaqus was then utilized to test the integrity of this structure under a worst case load condition derived from the ELaNa 6 launch vibration profile. Results from the analysis show that Objet FullCure720 photopolymer resin, DSM Somos Prototherm 12120 photopolymer resin, and Windform XT carbon fiber filled nylon all provide adequate strength to survive the environmental testing conditions required for this system to proceed through flight integration and launch.

CubeSats provide an inexpensive means for space-based research. However, optimal mission design depends on minimizing payload size and power. This thesis investigates the GLO (GFCR (Gas Filter Correlation Radiometry) Limb Occultation) prototype, a new small-form-factor design that enables sub-kilometer resolution of the vertical profile of atmospheric trace species to determine radiative influences. This technology improves SWAP (Size, Weight, And Power) over heritage SOFIE and HALOE instruments and provides a cost-effective alternative for solar occultation limb monitoring. A python script was developed to analyze solar intensity through GLO telescope channels. Non-uniform aerosol images used a peak intensity algorithm compared to the edge detection function designed for GFCR channels. Scaling corrections were made for beam splitter inaccuracy and SNR was characterized for frame collection. Different cameras were tested to weigh accuracy versus cost of a camera baffle. Using the Langley plot method, solar intensity versus changes in the solar zenith angle were measured for extrapolation of optical depths. AERONET, a network of ground-based sun photometers measuring atmospheric aerosols, was used for aerosol optical depth validation. Spectral Calculator transmission data allowed for GFCR vacuum channel comparison, gas cell spectral analysis, and gas cell to vacuum channel optical depth examination. Ground testing provided promising results with the low-cost prototype. It will be further evaluated through a balloon flight demonstration using a flight-ready GLO instrument. Additionally, analysis for the DUSTIE mission is planned and simulated using STK and Matlab. This includes CubeSat bus selection, orbit analysis for occultation occurrences, power budgeting, and communication capabilities. / Master of Science

In the last decade the world has seen a steadily increasing number of Cube Satellites deployed to Low Earth Orbit. Traditionally, these cubesats rely on Amateur Radio communications technology that are proven to work from space. However, as data volumes increase, the existing Amateur Radio protocols, combined with the restrictions of use for the Amateur Radio Spectrum, as well as the trend to build one control station per cubesat, result in a bottle neck effect whereby existing communications methods are no longer sufficient to support the increasing data volumes of the spacecraft. This Masters Thesis work explores the concept of deploying a network of distributed ground station receiver nodes for the purposes of increasing access time to the spacecraft, and thereby increasing the potential amount of data that can be transferred from orbit to the ground. The current trends in cubesat communications will be analyzed and an argument will be made in favor of transitioning to more modern digital communications approaches for on orbit missions. Finally, a candidate ground station receiver node design is presented a possible design that could be used to deploy such a network. / Master of Science

A retarding potential analyzer (RPA) is an instrument that, when mounted on a satellite in low-Earth orbit, makes in-situ measurements of ion density, temperature and speed relative to the satellite frame. The instrument works by changing the voltage on one of a set of grids and measuring a corresponding current generated by ions flowing through the grid, generating a function of current vs. voltage called an I-V curve. Traditionally, the size and power requirements of retarding potential analyzers has limited their use to larger satellites. In this thesis, the electrical design and basic testing of a retarding potential analyzer for use on resource- limited cubesat platforms are described. The mechanical design of the retarding potential analyzer is first described, and the requirements of the electrical design are presented. The electrical requirements are based on both the characteristics of the ionosphereic flight environment, and on the size and power requirements typical of the small cubesat platforms for which the instrument is intended. The electrical hardware is then described in detail. The digital design is reviewed as well, including the instrument's operating modes, command and data structure, and timing scheme. Test data showing the basic functionality of the instrument are then presented. Bench tests validate the design by confirming its ability to control voltages and measure small currents. End-to-end tests were also performed in a vacuum chamber to mimic the ionospheric environment. These data are presented to show the ability of the RPA to meet or exceed its design specifications. / Master of Science