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

The bulk motion of the neutral gas at altitudes between about 200 and 600 km is an important factor in predicting the onset of plasma instabilities that are known to distort and/or disrupt high frequency radio communications. A ram wind sensor is a space science instrument that, when mounted on a satellite in low-Earth orbit, makes in-situ measurements of the component of the neutral gas velocity that lies along the orbit track of the satellite. The instrument works by changing the voltage on one of a set of grids and measuring a corresponding electron current generated by ions flowing through the grid stack and detected by the microchannel plate, generating a function of current vs. voltage called an I-V curve. Traditionally, the size and power requirements of ram wind sensors has limited their use to larger satellites. In this thesis, the electrical design and basic testing of a cubesat compatible RWS known as the ram energy distribution detector (REDD) are described. The mechanical design of the REDD sensor 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 microcontroller design is reviewed as well, including the instrument's operating mode, 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 electron 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 REDD sensor to meet or exceed its design specifications. / 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

Nüvü Camēras, a Canadian company that designs a range of CCD and EMCCD cameras and controllers, recently began development on a miniaturized EMCCD controller for a CubeSat form factor. The detector for this payload requires near-cryogenic temperatures, approximately 188K, for performance operation. A temperature requirement of that magnitude for a CubeSat form factor is challenging given the low thermal mass, volume, surface area, and power availability for heat storage, dissipation and control systems that would typically be available for larger form factor spacecraft. The goal of this project is to design and per- form thermal analysis for the Nüvü Camēras CubeSat EMCCD Controller that allows for cold-biased active temperature control of both the controller electronics and detector. The EMCCD controller had an operational temperature range of −35◦C to +60◦C while the detector had a performance range of −110◦C to −85◦C with a desire to maintain a resolu- tion of ±0.25◦C. To meet these requirements, a system was designed within 3D modeling software Autodesk Inventor and imported into Thermal Desktop for thermal analysis and iteration. Models were updated based on thermal analysis results, adjusted by hand, and then tested again until a passive cooling and active heating system that met the require- ments was achieved. The final control system was shown to be capable of cooling from 20◦C (293.15K) to −85◦C (188.15K) and beyond given a Sun Synchronous orbit at 600km with attitude control and operational requirements. It was also shown to be capable of heating up, using resistive heaters on key components, beyond the thermal inertia of the system and environment, indicating viable control on orbit. In the future a PID control method can be implemented, and its use is being investigated by Nüvü Camēras for achieving the desired resolution of ±0.25◦C in the future. / Master of Science / Nüvü Camēras, a Canadian company that designs a range of Charge-Coupled Device (CCD) and Electon Multiplication Charge-Coupled Device (EMCCD) cameras and controllers, re- cently began development on a miniaturized EMCCD controller for a CubeSat form factor. CubeSats are a standard for nanosatellites, a classification of spacecraft with dimensions on the order of centimeters instead of meters. The Nüvü Camēras CubeSat EMCCD controller will operate a detector that enables the collection of extremely low light images, or even the counting of individual photons, that other CCD detectors are not capable of in such a form factor. This kind of quality comes with operating requirements such as a very low tempera- ture for the detector, which can be challenging to achieve in a CubeSat. The challenges are the low thermal mass for heat storage, low volume and surface area for fitting components and dissipating heat via radiation, and low power availability to run active systems such as cryocoolers. The goal of this project was to develop a temperature management system that tackles these challenges and allows the operation of the Nüvü Camēras CubeSat EMCCD controller on a variety of Low Earth Orbit (LEO) missions. To achieve this, a passive cool- ing system was designed to be cooled below the necessary temperature such that it can be effectively warmed back up as a method of active control with resistive heaters. This system was designed in a Computer Aided Design (CAD) program called Autodesk Inventor and analyzed in a thermal analysis program called Thermal Desktop. After a final analysis of the full setup, it was determined that the system would allow for thermal management at LEO for the Nüvü Camēras CubeSat EMCCD controller and detector.

Dévoiler les mystères du système solaire pour comprendre les mécanismes de la formation de la Terre, pour rechercher des signes de vie ou pour développer des colonies sur d’autres planètes, dépend de notre capacité à repousser les limites de l'ingénierie et de la science. Pour cela, il est important de développer des technologies de pointe pour permettre aux véhicules spatiaux de survivre la phase d'entrée ou de rentrée atmosphérique. Lors de l’entrée ou de la rentrée, l’engin spatial peut être exposé à flux radiatifs intenses qui ne peuvent pas encore être prédits avec précision, imposant ainsi des marges de sécurité sur la conception des systèmes de protection thermique. Ces incertitudes augmentent lorsque le bouclier thermique est constitué d'un matériau ablatif car sa dégradation introduit de nouvelles espèces chimiques réagissant avec le plasma produit devant le véhicule, ce qui affecte le rayonnement. Le but de cette thèse est d’étudier les flux de chaleur radiatifs sur un véhicule de rentrée de petite taille en présence d’un bouclier ablatif (Thermal Protection System, ou TPS), en utilisant des simulations numériques et des expériences pour développer un instrument de vol qui sera embarqué à bord du CubeSat QARMAN.Une évaluation de la trajectoire de rentrée du véhicule QARMAN (masse : 5 kg) a été réalisée en utilisant un code maison à 6 degrés de liberté. Un ensemble de simulations Monte Carlo ont permis de quantifier les incertitudes et ont montré un maximum de ± 15% écart par rapport à la trajectoire nominale. Les spectres sans ablation ont alors été déterminés en utilisant une approche découplée avec deux codes : Stagline (VKI) et SPECAIR (EM2C, CentraleSupélec). Ces simulations ont été effectuées pour la trajectoire nominale ainsi que pour la gamme des incertitudes. Elles ont permis de mettre en évidence un comportement non-linéaire des caractéristiques spectrales par rapport aux valeurs nominales, avec une augmentation drastique vers la fin de la mission.Les effets de l'ablation ont été étudiés avec une nouvelle technique de mesure développée au cours de cette thèse. Basée sur deux sondes de mesure de rayonnement, l’une refroidie et l’autre recouverte d’un matériau ablatif, cette méthode permet de quantifier l'émission et l'absorption induite par tout type de TPS ayant des interactions gaz-surface avec l'écoulement, dans l’hypothèse que les raies d’émission et d’absorption des espèces ablatives ne soient pas superposées. La méthode a été validée sur un échantillon de graphite TPS. Elle a ensuite été appliquée à la prédiction du rayonnement attendu lors de la mission QARMAN (Cork P50 TPS). Cette étude a également permis de sélectionner un spectromètre d’émission adapté à la mission QARMAN et aux objectifs de la thèse (plage de 350 à 800 nm pour une masse de 68 g).Un instrument de mesure de rayonnement standard a été testé et les limites de cet appareil ont été établies. Deux nouvelles technologies ont été développées et la charge utile (spectromètre d’émission INES) a été construite et intégrée au véhicule QARMAN. Un étalonnage spectral et thermique dédié a également été développé pour maximiser la qualité du retour scientifique en prenant en compte les variations de température dans la baie de charge utile de QARMAN.L’instrument proposé est, à ce jour, la seule charge utile non intrusive capable d’effectuer des mesures radiatives sans limitations liées à la contamination par les poussières et gaz d'ablation. L’instrument peut aussi fournir des mesures de la récession, de la sublimation et du gonflement du TPS avec une précision d'au moins 0,2 mm. Le fonctionnement de l'appareil a été démontré pour une grande variété de conditions de test, y compris différents profils d'enthalpie, mélanges de gaz et matériaux de TPS. / Unveiling the mysteries of the solar system to understand the mechanisms of Earth’s formation, to search for signs of life, or to develop settlements on other planets, depends on our abilities to push the limits of engineering and science. One of the key aspects of space exploration is the development of advanced technologies to sustain the entry/reentry phase. During entry or reentry, the spacecraft may be exposed to intense radiative fluxes that cannot be accurately predicted yet, thus imposing high safety margins on the design of thermal protection systems. These uncertainties rise when the heat shield is made of an ablative material as its degradation introduces new chemical species reacting with the flow affecting radiation processes. The goal of this thesis is to study the radiative heat fluxes onto a small size reentry vehicle in the presence of an ablative TPS, using numerical simulations and experiments to develop a flight instrument that will be carried onboard the QARMAN CubeSat.An assessment of the reentry trajectory of the 5-kg QARMAN vehicle was performed using a custom 6-degree of freedom code. An extensive set of Monte Carlo simulations allowed to quantify uncertainties and showed a maximum of ±15% deviation from the nominal trajectory. The spectra without ablation were then computed using a decoupled approach with two codes: Stagline (VKI) and SPECAIR (EM2C, CentraleSupélec). These simulations were performed for the nominal trajectory as well as for the range of uncertainties. They showed a nonlinear behavior of the spectral features deviations from nominal with a drastic increase toward the end of the mission.The effects of ablation were studied with a new measurement technique developed during this thesis. Based on two radiation measurement probes, one cooled and the other with an ablative surface, it allows to quantify the emission and absorption induced by any kind of TPS having gas-surface interactions with the flow, provided that the radiative emission or absorption features of the ablative species do not fully overlap. The method was validated on a graphite TPS sample. It was then applied to determine the radiation expected during the QARMAN mission (Cork P50 TPS). This study also allowed to select an emission spectrometer (350-800 nm range for a 68-g mass).A standard radiation instrument was tested and the limits of this device shown. On those lessons learned, two new technologies were developed and an emission spectrometer payload (INES) was built and integrated into the QARMAN reentry CubeSat. A dedicated spectral and thermal calibration was also developed to maximize the quality of the scientific return by tackling the non-standard internal temperature variations of QARMAN’s payload bay.Relying on two inventions made during this study, the apparatus is at the time of writing, the only non-intrusive payload capable of making radiative measurements without limitations due to ablation dust contamination. The instrument can also provide measurements of recession, sublimation and swelling of the TPS with a precision of at least 0.2 mm. Operation of the apparatus was demonstrated for a wide variety of test conditions, including different enthalpy profiles, gas mixtures and TPS materials.

The focus of this thesis is the design, manufacturing and testing of a deployable radio antenna for the ORCASat nanosatellite. First, the context, motivation, requirements, as well as constraints for this project are introduced. Next, a brief overview of theoretical concepts relevant to the contents of this thesis are presented. After the introduction of the relevant background and theory, a literature review is undertaken, and an experiment-based methodology is established. Prior to conceptualizing a new design, detailed consideration is also given to previous attempts at designing a dipole for ORCASat. The root cause of the problems with these attempts is determined experimentally as the presence of ground planes on the circuit board supporting the antenna. After this preliminary investigation, the blocks required for the ORCASat antenna are introduced as the transmission line feeder, the balun, the impedance matching block, and the antenna arm feed. For each of these components, competing design concepts are developed, and the advantages and disadvantages of each of these concepts are presented. After this, the winning design concept is selected and developed into a manufacturable design. This design is identified as a tunable tape spring half wave dipole antenna featuring a specialized feed with electrically and mechanically optimal characteristics, no impedance matching, and a lossy choke balun wound from the coaxial cable feeder, all mounted on a circuit board in a pre-existing Delrin antenna deployer. Next, the manufacturing and assembly of this design is undertaken, followed by the consideration of an informal commissioning procedure. As part of this, a test article consisting of an incomplete prototype of the dipole is tested, and it is shown to have desirable voltage standing wave ratio, input impedance, and return loss characteristics, as well as excellent tunability. Having established that this test article is a good candidate to meet project requirements, it is updated to include as many of the final components of the antenna as possible. Then, formal test procedures for the verification of the tunability, return loss, VSWR, input impedance, antenna pattern, and absolute gain are established, and executed. Based on the results of this formal verification test campaign, it is concluded that the test article meets the requirements presented at the beginning of this thesis, and it is suitable as a radio antenna for the ORCASat mission. After this, the work is concluded by a set of recommendations for future work to prepare the antenna developed in this thesis for flight. / Graduate

As development of CubeSat based architectures increase, methods of deploying constellations of CubeSats are required to increase functionality of future systems. Given their low cost and quickly increasing launch opportunities, large numbers of CubeSats can easily be developed and deployed in orbit. However, as secondary payloads, CubeSats are severely limited in their options for deployment into appropriate constellation geometries. This thesis examines the current methods for deploying cubes and proposes new and efficient geometries using secondary launch opportunities. Due to the current deployment hardware architecture, only the use of different launch opportunities, deployment direction, and deployment timing for individual cubes in a single launch are explored. The deployed constellations are examined for equal separation of Cubes in a single plane and effectiveness of ground coverage of two regions. The regions examined are a large near-equatorial zone and a medium sized high latitude, high population density zone. Results indicate that simple deployment strategies can be utilized to provide significant CubeSat dispersion to create efficient constellation geometries. The same deployment strategies can be used to develop a multitude of differently dispersed constellations. Different launch opportunities can be utilized to tailor a constellation for a specific region or mission objective. Constellations can also be augmented using multiple launch opportunities to optimize a constellation towards a specific mission or region. The tools developed to obtain these results can also be used to perform specific analysis on any region in order to optimize future constellations for other applications.

International Telemetering Conference Proceedings / October 21, 2002 / Town & Country Hotel and Conference Center, San Diego, California / A telemetry system has been developed at the University of Arizona to serve as a baseline for future CubeSat designs. Two satellites are scheduled for launch in November of 2002. One features a beacon that operates autonomously of all but the power system and can independently deploy the antennas. The other will test the performance of new semiconductor devices in low earth orbit. Sensors will monitor voltages, currents (from which attitude and tumble rate can be derived), received signal strength and a distribution of temperatures. The CubeSat’s architecture, operating system, sensors, telemetry format and link budget are discussed.

ITC/USA 2015 Conference Proceedings / The Fifty-First Annual International Telemetering Conference and Technical Exhibition / October 26-29, 2015 / Bally's Hotel & Convention Center, Las Vegas, NV / The goal of this project is maximize the performance of the telemetering system for the INCA cubesat mission using what we are learning in class to develop tests to figure out the optimal selection of frame scheme, data rate, and modulation technique based on the requirements of the mission. This project will help me learn about different modulation techniques and give me real world experience testing a telemetry system. I will evaluate my results through a comparison of the error rates for the different modulation schemes and do statistical analysis to show the reliability of the data. The results will be useful to any future mission that implements the same satellite communication system including future missions at NMSU.

This dissertation introduces novel techniques for exploiting the environmental aerodynamic forces to actively control the attitude of the spacecraft operating in the lower and middle thermosphere. It includes both simulations and real spacecraft attitude determination and control subsystem development, which provide a complete picture of the application of the aerodynamic forces to benefit space missions that are operating very close to Earth, as well as contribute to the knowledge of rarefied gas aerodynamics in the lower and middle part of the thermosphere. The research starts by reviewing the current progress of thermosphere science and rarefied gas aerodynamics to construct a high fidelity aerodynamic model for spacecraft operating in the rarefied gas (mainly atomic oxygen) environment in very low Earth orbits (below 450 km) and following by a brief system level analysis of the benefits and challenges for the spacecraft flying lower to Earth. A real spacecraft is also developed to validate of the application of the aerodynamic forces for attitude control. The aspect of the design included in this dissertation focuses mainly on the attitude determination and control system development of satellite. The CubeSat has a generic design with deployable solar panels that can be rotated to control the aerodynamic torques. Based on the common attitude control requirements of spacecraft operating in very low Earth orbits, and the hardware capability of the satellite three novel orbital aerodynamic attitude control strategies are proposed: Energy Optimized B-dot Detumbling into an Aerostable State; Active Orbital Aerodynamic Coarse Pitch/Yaw Control; a 3-axis Orbital Aerodynamic Torques Adaptive Sliding Mode Control. The control performance for each control algorithm is validated numerically in high-fidelity attitude propagators. Knowledge of the thermospheric winds is important as they influence the control performance and the dynamic response of the spacecraft, aerostable designs steering into the thermosphere wind vector. Two novel computational methods to measure the thermospheric wind from the dynamic response of the spacecraft due to aerodynamic forces are proposed. The in-situ measured wind vector benefits the attitude observation in the feedback control systems, which helps to improve the adapting performance and to increase the control accuracy. The proposed novel aerodynamic attitude control algorithms can be adapted for similar spacecraft operating in the very low Earth orbits with modifications to the deployable solar panels or adding movable aerodynamic control surfaces. In addition, this proposed orbital aerodynamic attitude control system works not only in the very low Earth orbits but can also be potentially implemented for spacecraft operating in the rarefied gas region of the atmospheres of other planets.