The space environment is a volatile and challenging place for satellites to survive in. For Low Earth Orbiting (LEO) satellites, atomic oxygen (AO) is a constant corrosive effect that degrades the outer surface of satellites over long durations. Atomic oxygen exists in the atmosphere between 180 and 675 km and has a relatively high energy at 4.5 eV, which allows AO to break molecular bonds in materials on the surfaces of spacecraft. As the number and complexity of CubeSat missions increase, there is an increased risk that AO degradation on commercial off the shelf parts (COTS), such as antenna, could degrade the satellite’s ability to communicate with ground systems. This thesis looks at how AO erosion affects the performance of patch antennas for CubeSat applications. Patch antennas are small, cheap, low-profile antennas that can be used on CubeSats to communicate with the ground or other satellites. Patch antennas are semi-directional, providing higher gain and higher available frequencies than omnidirectional antennas. An AO chamber in the California Polytechnic State University San Luis Obispo’s (Cal Poly) Spacecraft Environments Testing Lab was used to expose the patch antennas for 24-hour and 48-hour tests. The 24-hour exposure saw an average AO fluence of 8.757 ± 0.807•1020 atoms/cm2 which corresponds to roughly 3.5 months of on-orbit AO exposure on the Ram side when in a 28.5° inclined orbit with an altitude of 400 km. The 48-hour exposure saw an average AO fluence of 1.595 ± 0.076•1021 atoms/cm2 which corresponds to approximately 6.4 months of on-orbit AO exposure on the Ram side when in a 28.5° inclined orbit with an altitude of 400 km. To test the performance of the patch antenna before and after AO exposure, an anechoic chamber in the Microwave Lab at Cal Poly was used to measure boresight gain and radiation pattern in the E-plane and H-plane. From the testing in the anechoic chamber it was determined that there was no apparent difference in the patch antenna’s gain and radiation pattern before and after AO exposure. By using a Fourier Transform Infrared Spectrometer (FTIR) it was discovered that the outer surface of the patch antennas were forming a silicon dioxide layer, which did not affect the performance of the patch antenna. Since silicon dioxide is resistant to AO erosion, it may be beneficial for CubeSats to include silica additives to their exposed antenna surfaces to prevent erosion.

The CubeSat Astronomy Network is a proposed system of multiple CubeSat spacecraft capable of performing follow-up observations of astronomical targets of interest. The system is intended to serve as a space-borne platform that can complement existing systems utilized for astronomical research by undergraduate and high school students. Much research and development work has been performed to develop model-based system engineering methodologies and products for CubeSat missions, including the Horizon Simulation Framework. The Horizon Simulation Framework enables the development of system models using the Extended Markup Language (XML), and its simulation program can generate system simulations over model-specified timespans. System requirements and constraints, as well as subsystem dependencies and functions, can also be directly specified in these models. Previous work using the framework has been performed to characterize “day-in-the-life” operations for Earth-observing spacecraft. A similar goal is intended for modeling the CubeSat Astronomy Network: simulating mission operations during nominal conditions to validate system and subsystem requirements. By developing this model, system and subsystem requirements derived in the course of preliminary design for the Network can be analyzed, modelled, and evaluated for feasibility. These results can then be used to inform design decisions related to system architecture and concept of operations at the early stages of design, while the models themselves can grow and mature alongside project development and be re-used for future design work.

The CUBES (CUbesat x-ray Background Explorer using Scintillators) is a payload on the KTH student satellite MIST (MIniature STudent satellite) to evaluate Silicon Photo-multiplier technology and new scintillators such as GAGG (Gadolinium Aluminium Gallium Garnet, Gd3Al2Ga3O12) for future use in hard X-ray polarisation studies of Gamma-Ray Bursts. CUBES itself is designed to study the MIST in-orbit radiation environment by using a detector which is comprised of a silicon photomultiplier coupled to different scintillator materials. Three of these detectors will be mounted on the payload platform and then coupled to inputs of an Application Specific Integrated Circuit (ASIC) and connected to a Field-programmable Gate Array (FPGA) which will store and send data through the downlink on the MIST satellite to ground. This thesis covers the software development for the FPGA, together with two radiation tests of components and the preparation of these. / CUBES / MIST

Memoria para optar al título de Ingeniero Civil Eléctrico / En el vuelo espacial, una maniobra orbital corresponde al uso de sistemas de propulsión para cambiar la órbita de un vehículo espacial. Es actualmente casi la única manera de desplazarse a través del espacio exterior y por lo tanto su aplicación resulta de gran importancia para el diseño físico como para el diseño de misiones de satélites. En el caso particular de nanosatélites, los sistemas de propulsión presentan grandes restricciones tanto de capacidad como tamaño, por lo cual normalmente se hace necesario utilizar sistemas de propulsión eléctrica, los cuales poseen un nivel de empuje bajo, resultando comúnmente en tiempos de propulsión de larga duración, del orden de cientos de órbitas. El presente trabajo pretende calcular los propulsivos, energéticos y temporales de llevar a cabo maniobras orbitales de bajo empuje predeterminadas. Es decir, la cantidad de propelente, potencia, energía y tiempo necesarios para ejecutar una maniobra sub-óptima definida manualmente en base a resultados de la literatura. En la primera parte del presente trabajo se presenta el marco teórico donde se describen los conceptos necesarios para poder comprender y analizar el modelo realizado. Se mencionan principalmente conceptos relacionados con la astrodinámica, los principios de propulsión y la ejecución de maniobras orbitales. Se construyó un modelo en python basado en las ecuaciones de variación de parámetros incorporando perturbaciones gravitacionales de la Tierra, el Sol, la Luna, el arrastre atmósférico y la presión solar. Este modelo además incorporó la capacidad de perfilado del empuje a lo largo de su órbita y finalmente la fijación de órbitas objetivo basado en leyes de control derivadas analíticamente. Además se validó la dinámica básica y perturbada del modelo mediante comparaciones con el software comercial de simulación de satélites Systems Toolkit STK. Una vez completado el modelo se procedió a realizar las simulaciones de intéres, incluyendo desorbitación, mantenimiento orbital y movimiento relativo. A partir de los escenarios estudiados se estima, en primer lugar, que el satélite SUCHAI tendrá un tiempo de desorbitación de 7 años, cayendo entre 2024 y 2025. Éste tiempo puede ser reducido entre un 20% y 30% utilizando propulsión basada en componentes comerciales. Las maniobras probadas, a nivel general, no poseen mayor problema energético. En el caso de mantenimiento orbital en órbita baja, la perturbación que genera mayor efecto es el arrastre atmosférico, por lo cual sólo resulta conveniente modificar el semieje-mayor. Finalmente se observa para el escenario de movimiento relativo, que en ausencia de perturbaciones es imposible que un chipsat expulsado de un cubesat en órbita quede orbitándolo.

Tesis para optar al grado de Magíster en Ciencias de la Ingeniería, Mención Eléctrica / Memoria para optar al título de Ingeniero Civil Eléctrico / Históricamente, la mecánica estadística ha creado herramientas para describir la evolución de sistemas y procesos en equilibrio termodinámico. Sin embargo, los procesos del mundo real no siempre ocurren en condiciones de equilibrio. La turbulencia en fluidos, la materia granular y las máquinas moleculares son sistemas que tienen que lidiar constantemente con esta condición. En base a esto, se han desarrollado herramientas ampliamente utilizadas por la comunidad científica, conocidas como los Teoremas de Fluctuación. No obstante, se ha demostrado -mediante experimentos y simulaciones- que dichos teoremas no son válidos incluso en sistemas de primer orden. Especificamente en [1], se demostró que para un circuito RC fuera del equilibrio, las fluctuaciones de potencia inyectada se atañen a los teoremas de fluctuación solamente si la magnitud de las fluctuaciones son acotadas a un rango específico, lo cual rápidamente deja de ser cierto al aumentar la magnitud del forzante. En vista de esta problemática, este trabajo de tesis busca ampliar la investigación anterior mediante la exposición de un circuito RC a un ambiente espacial. El objetivo principal es desarrollar un experimento que se inserta como carga útil o payload para el nanosatélite SUCHAI. Y además se busca medir los cambios en las fluctuaciones de potencia inyectada con respecto al ambiente espacial. Este payload forma parte de la misión de SUCHAI y conforma la primera iteración de una familia de experimentos electrónicos que permiten acceder al espacio a tiempo real y a costos accesibles. Los resultados obtenidos muestran que es posible forzar un circuito RC a un estado fuera del equilibrio bajo las restricciones del Cubesat. Sin embargo, los datos satelitales no muestran diferencias sustanciales con respecto a las fluctuaciones en tierra. Con respecto al escenario descrito, se realizaron pruebas en ambientes controlados de presión (5 · 10 −6 y 760 [Torr]) y temperatura (−30 ◦ C a 45 ◦ C); donde simultáneamente se comparó la decisión de utilizar un generador de señales y un osciloscopio para excitar y medir el circuito. Estos datos tampoco muestran una diferencia en las fluctuaciones generada por los cambios de presión y tempe- ratura. En una prueba final, se propuso medir un RC equivalente independiente al satélite y además filtrar la respuesta del generador de señales desde 20 MHz a 1.8 KHz, donde se logró percibir cambios considerables en las fluctuaciones debido al cambio de presión atmosférica. En conclusión, se establece la posibilidad de forzar un circuito RC a un estado fuera del equilibrio de forma controlada dentro de un Cubesat. Además, se demuestra la resilencia de los componentes RC comerciales de tecnología SMD a los cambios de presión y temperatura. Por otra parte, la elección de instrumentos de excitación (generador de números aleatorios y DAC), junto los instrumentos de medición (ADC) y el espectro del forzante para el ex- perimento deben ser probados anteriormente en ambientes controlados como una cámara de termovacío, para así validar la factibilidad de medir el ambiente mediante este enfoque.

<p>The commercial CubeSat industry has experienced explosive growth recently, and with falling  costs  and  growing  numbers  of  launch  providers,  the  trend  is  likely  to  continue.  The scientific missions CubeSats could complete are expanding, and this has resulted in a demand for reliable  high  specific  impulse  nanosatellite  propulsion  systems.  Interest  in  liquid-fed  pulsed plasma thrusters (LF-PPTs) to fulfill this role has grown lately. Prior work on a nanosatellite LF-PPT was done in the Purdue Electric Propulsion and Plasma Laboratory, but its high operational voltage and electrode size would be disadvantageous for integration on a CubeSat, which have strict volume limitations and provide only tens of Watts in power at low voltages. This work aims to address those disadvantages and further advance the development of a nanosatellite LF-PPT by reducing the operating voltage and removing long plate electrodes to prevent energy losses on components other than the expelled plasma sheet. Two major objectives are pursued: to construct a  coaxial  pulsed  plasma  thruster  operating  with  10s  to  100s  of  volts  and  to  characterize  the temporal evolution of the discharge parameters in this low-voltage operation scenario. </p> <p>It  took  three  experimental  design  iterations,  all  of  which  used  a  260  <em>uF</em> ,  400 <em>V</em> film capacitor, to arrive at a functional coaxial pulsed plasma thruster. First, a button gun was tested. It produced  a  peak  current  of ~16<em> kA</em>,  which  serves  as  the  expected  maximum  for  the  later experiments. Due to the presence of parasitic arcing, it revealed that electrical lines needed to be removed from vacuum chamber to enable testing at a wide range of pressures. Second, a coaxial PPT was designed, built, and tested. This design confirmed operation at discharge voltages <100 <em>V</em> across the plasma, achieving one of the project’s aims, and produced a peak current of 7.4 <em>kA</em>. However,  necessity  to  better  align  the  cathode and  provide  an  unobstructed  camera  view  for observation of the discharge column attachment to the cathode surface forced additional system redesign. Third, a revised coaxial PPT was built and tested. Using air as a propellant, the discharge generated a peak current of 10.4 <em>kA</em> at a mass flow rate of 2 mgs. The PPT cathode was imaged with an ICCD camera over a wide range of pressures, and the photos indicated “spotless” diffuse arc attachment to the cathode, which serves as evidence to expect low erosion rates. The direct measurements of the cathode erosion rate are planned for future. </p>

This dissertation research presents a comprehensive study to answer the question of “Can it be possible to integrate a high gain optically transparent antenna array directly on top of solar cells?”. The answer to such question is extremely important in space exploration where very small satellites have been extensively employed. Due to their small mass and size, those small satellites create challenges for one to mount the antennas, and the challenge is further increased when a high gain antenna is need for more communication capacity. Based on feasibility studies, the dissertation concludes that it is possible to do such an integration, and then proceeds to present the approaches for design and integration. On the element level, the thesis presents research in assessing the effects between a planar antenna integrated on the solar cell and the photovoltaic cell. A series of experiments were designed to perform assessments for antennas operating from C to X bands. It is concluded that a commercial triple junction space–certified solar cell normally would decrease the gain of the antenna to 2–3 dB and is not affected by the working states of solar cells. The shadow of the antenna casts on solar cells, however, is not significant (less than 2%). The thesis also provides a model of a common space solar cell that helps to explain the gain loss. The model was validated by experimental data, and it was utilized to predict iv a possible custom design of solar cell where with a minimal design modification, it would facilitate less gain loss of the antenna integrated on top. On the array level, the research surveys different high gain antenna array design and then focus on an optimal sub–wavelength reflectarray design. The final antenna array design is a 30 cm by 20 cm, X band (8.475 GHz) reflectarray that shows 94% transparency, 24 dB gain, and higher than 40% aperture efficiency. The design is then prototyped and tested on actual solar panel. The measurement of the reflectarray placed on the solar panel showed a gain of 22.46 dB and an aperture efficiency of 29.3%. While those results are considered excellent, the thesis continues to address the reasons for reduction of the antenna’s performance due to the solar panel, through both theoretical analysis and experiments.

This thesis outlines an orbit determination and 3-axis attitude determination system for use on orbit as applicable to 1U CubeSats and other picosatellites. The constraints imposed by the CubeSat form factor led to the need for a simple configuration and relaxed accuracy requirements. To design a system within the tight mass, volume, and power constraints inherent to CubeSats, a balance between hardware complexity, software complexity and accuracy is sought. The proposed solution consists of a simple orbit propagator, magnetometers with a magnetic field look-up table, Sun sensors with an analytic Sun direction model, and the TRIAD method to combine vector observations into attitude information. The orbit propagator is a simple model of a circular trajectory with several frequently updated parameters and can provide orbital position data with average and maximum errors—when compared to SGP4—of less than 3.7km and 10.7km for 14 days. The magnetic field look up table provides useful information from a small memory footprint; only 480 data points provide a mean error of approximately 0.2° and a maximum error of approximately 2°—when compared to the IGRF model. The Sun’s direction is modeled, and as expected, can be modeled simply and accurately. Combining the magnetic field and Sun direction models with inaccurate sensors and the TRIAD method results in useful attitude information from a very simple system. A system with Sun sensor error standard deviation of 1° and magnetometer error standard deviation of 5° yields results with average error of only 2.74°, and 99% of the errors in this case are less than approximately 13°. The system outlined provides crude attitude determination with software and hardware requirements that are well within the capabilities of current 1U CubeSats—something that many other systems, such as Kalman filters or star trackers, cannot do. It also provides an excellent starting point for future ADCS systems, which will significantly increase the ability of CubeSats.

Cal Poly San Luis Obispo’s PolySat team is designing the Multipurpose Orbital Spring Ejection System (MOSES) in order to record acceleration data during the launch of CubeSats as well as to provide GPS coordinates to locate the position of CubeSats once they are injected into orbit. This work focuses on the design and development of the acceleration data acquisition (DAQ) subsystem of MOSES. This subsystem is designed around the need for a high-speed sampling system of at least 200 kHz across four channels of data, plus low-area limitations in the MOSES form factor which is roughly half the size of a standard CubeSat. To address these specifications, the design explores system implementation around a Xilinx Artix-7 FPGA with a built-in analog-to-digital converter and a custom hardware solution.

CubeSat designers are increasingly looking to incorporate delicate structures and optics into their payloads. These delicate payloads, however, may not survive the required absolute-worst-case launch vibration testing needed for flight certification. To help address this problem, and to better match testing conditions to real-world launch environments, this thesis introduces Sync, a compact 1/4U CubeSat payload designed to collect data on the vibrations and thermal environments CubeSats experience inside a deployer on the way to orbit. This data can be used to better understand the launch environment for different vehicles, and help develop new, more realistic testing guidelines that could enable more delicate payloads to be launched.