CubeSat development has seen a rise since the first launch in 2003 due to faster design process and low launch costs. It has played a vital role in providing access to space to small start-ups and academic organizations with low budgets. It has also enabled the testing of different upcoming technologies in space and has helped in providing hands-on experience to students taking part in design of such platforms. University of Pisa, in collaboration with SITAEL, has also taken an initiative to design and develop a CubeSat to test the FEEP thruster, design of which is presented in the thesis. A FEEP system was designed to fit within 1U dimensions and with a dry mass of approximately 820 grams. The system is based on slit emitter which provides an advantage over already available technologies in the market which uses needle emitters. Slit emitter scan achieve multiple Taylor cones without the need of clustering as used in needle emitters and also have a higher Thrust to Power Ratio. A propellant comparison was done considering all the properties required for an ideal propellant for a FEEP system. This comparison led to the selection of indium as working propellant which has an atomic mass of 114.8 u and a melting point of 156.6 °C. The FEEP system was designed keeping in mind easy assembling and modularity of thruster for ease in changing parts. The design consists of three different modules that are assembled separately and then joined together to complete the assembling of the system. The propellant tank, which also houses the emitter, has an internal volume of 32.75 cm3 and can hold approximately 240 grams of indium, which has a density of 7.31 g/cm3. During mission analysis, a 600km altitude orbit was proposed by analyzing the amount of propellant required for drag compensation and de-orbit maneuver at different altitudes with worst case values for ballistic coefficient and Thrust to Weight Ratio. At this altitude, the propellant requirement is 254.4 grams, 14.4 grams more than that of what can fit in the propellant tank of the designed thruster. However, both design of the system and mission analysis are ongoing processes and changes would be made in the future to either one or both to meet the requirements.

Cube Satellite (CubeSat) launches have been on the rise since its first launch in2003. This popularity is mainly due to faster design process and lower launch costs.However, most CubeSats are launched into Low Earth Orbits (LEOs), with nomissions to Geostationary Transfer Orbits (GTOs). However, many mission areplanned for the next half-decade.A major challenge to launch a CubeSat into a GTO is the thermal environmentof the higher altitude orbits. These orbits are significantly colder due to reducedheating from Earth’s planetary and albedo radiations, and a possibility for longereclipses due to the eccentricities of GTOs.A thermal investigation of the thermal environment was done using the Foresailmissions as examples, as the missions currently are set to fly the first missions toPolar LEO. The trajectories for the second Foresail mission are being evaluated,with the GTO being a strong contender. This thermal investigation is done througha comparative study of the two missions. The thermal effects of a few missionspecific scenarios were also evaluated.This provided a holistic thermal design of the first Foresail mission. A region specificthermal solution for the battery was analyzed. The various mission scenarios andtheir comparisons with the LEO mission, also formed a basis of the feasibility ofvarious situations on the second mission. Moreover, the results, in part assessedthe thermally feasibility to launch a 3U CubeSat into a GTO.The results showed GTOs show larger magnitude of variation of thermal loads ascompared to LEOs. However, these variations are more gradual due to the largerorbital periods. A 3U CubeSat can be launched into both, the LEO and GTOenvironments with passive thermal control. The properties of the thermal coatsvary slightly. However, it is not possible to passively control the CubeSat if theeclipse occurred at the aphelion of the orbit.

Small satellites are satellites that weight less than 500 kg. Compared to larger satellites, a small satellite, especially a cube satellite, has limited surface area. The limited surface area casts challenges for allocating essential parts, such as antennas, for the satellite. Therefore, antennas that are conformal to the satellite surface have distinct advantages over other types of antennas that need significant mounting area. One of the very effective conformal antennas is cavity-backed slot antennas that can be integrated around solar cells and do not compete for extra surface area. The previous study performed on cavity-backed slot antennas was mainly a feasibility study and did not address realistic concerns such as effective feeding methods for the antennas. This thesis work is aimed at providing more detailed study on achieving high quality circular polarization (CP) and simplified feed design to initiate effective integration of the antenna with solar panel. In order to accurately characterize an antenna, an effective antenna range in an anechoic chamber is important. Utah State University had an effective near-field range; however, there was not an fully shielded anechoic chamber. As another objective of this thesis work, a state-of-the-art anechoic chamber has been constructed, calibrated, and utilized to measure different antenna parameters. This thesis also shows correct methods to measure important antenna properties such as CP and antenna efficiency.

The current technological age is embodied by a constant push for increased performance and efficiency of electronic devices. This push is particularly observable for technologies that comprise free electron sources, which are used in various technologies including electronic displays, x-ray sources, telecommunication equipment, and spacecraft propulsion. Performance of these systems can be increased by reducing weight and power consumption, but is often limited by a bulky electron source with a high energy demand. Carbon nanotubes (CNTs) show favorable properties for field electron emission (FE) and performance as electron sources. This dissertation details the developments of a uniquely designed Spindt type CNT field emission array (CFEA), from initial concept to working prototype, to specifically prevent electrical shorting of the gate. The CFEA is patent pending in the United States. Process development enabled fabrication of a CFEA with a yield of up to 82%. Furthermore, a novel oxygen plasma etch process was developed to reverse shorting after CNT synthesis. CFEA testing demonstrates FE with a current density of up to 293 μA/cm² at the anode and 1.68 mA/cm² at the gate, with lifetimes in excess of 100 hours. A detailed analysis of eighty tested CFEAs revealed three distinct types of damage. Surprisingly, about half of the damaged chips are not electrically shorted, indicating that the CFEAs are very robust. Potential applications of this technology as cathodes for spacecraft electric propulsion were explored. Exposure to an operating electric propulsion thruster showed no significant effect or damage to the CFEAs, marking the first experimental study of CNT field emitters in an electric propulsion environment. A second effort in spacecraft propulsion is a collaboration with the Air Force Institute of Technology (AFIT). CFEAs are the payload on an AFIT developed Cube Satellite, called ALICE, to test electron emission in the space environment. ALICE has passed flight tests and is awaiting launch scheduled for 5 December 2013.