The structure of direct-current microdischarges is investigated using a detailed two-dimensional multi-species continuum model. Microdischarges are directcurrent discharges that operate at a relatively high pressure of about 100 Torr and geometric dimensions in the 10-100 micrometer range. Our motivation for the study of microdischarges comes from a potential application of these devices in microthrusters for small satellite propulsion. The Micro Plasma Thruster (MPT) concept consists of a direct-current microdischarge in a geometry comprising a constant area flow section followed by a diverging exit nozzle. A detailed description of the plasma dynamics inside the MPT including power deposition, ionization, coupling of the plasma phenomena with high-speed flow, and propulsion system performance is reported in this study. A two-dimensional model is developed as part of this study. The model consists of a plasma module coupled to a flow module and is solved on a hybrid unstructured mesh framework. The plasma module provides a self-consistent, multispecies, multi-temperature description of the microdischarge phenomena while the flow module provides a description of the low Reynolds number compressible flow through the system. The plasma module solves conservation equations for plasma species continuity and electron energy, and Poisson’s equation for the self-consistent electric field. The flow module solves mass, bulk gas momentum and energy equations. The coupling of energy from the electrostatic field to the plasma species is modeled by the Joule heating term which appears in the electron and heavy species energy equations. Discretization of the Joule heating term on unstructured meshes requires special attention. We propose a new robust method for the numerical discretization of the Joule heating term on such meshes using a cell-centered, finite volume approach. A prototypical microhollow cathode discharge (MHCD) is studied to guide and validate the modeling effort for theMPT. Computational results for the impedance characteristics as well as electrodynamic and chemical features of the discharge are reported and compared to experimental results. At low current (< 0.1 mA), the plasma activity is localized inside the cylindrical hollow region of the discharge operating in the so-called “abnormal regime”. For larger currents, the discharge expands over the outer flat surface of the cathode and operates in the “normal regime”. Transient relaxation oscillations are predicted in the plasma properties for intermediate discharge currents ranging from 0.1 mA to 0.3 mA; a phenomenon that is reported in experiments. The MPT, in its present configuration, is found to operate as an electrothermal, rather than as an electrostatic thruster. A significant increase in specific impulse, compared to the cold gas micronozzle, is obtained from the power deposition into the expanding gas. For a discharge voltage of 750 V, a power input of 650 mW, and an argon mass flow rate of 5 sccm, the specific impulse of the device is increased by a factor of 1.5 to a value of 74 s. The microdischarge remains mostly confined inside the micronozzle and operates in an abnormal regime. Gas heating, primarily due to ion Joule heating, is found to have a strong influence on the overall discharge behavior. The study provides crucial understanding to aid in the design of direct-current microdischarge based thrusters. / text

Este trabalho teve como objetivo o desenvolvimento de materiais catalíticos à base de óxidos mistos, de baixo custo, para serem empregados como catalisadores mássicos, na decomposição do peróxido de hidrogênio 90%, em massa, possibilitando múltiplas partidas a frio em um micropropulsor de satélites a monopropelente. Foram utilizados diferentes métodos de síntese de óxidos mistos com altos teores de manganês e cobalto. Os materiais foram avaliados, preliminarmente, na decomposição espontânea do peróxido de hidrogênio em bancada (teste da gota). Em seguida, os catalisadores com melhores desempenhos foram selecionados e testados em um micropropulsor de 2N, onde foram monitorados o empuxo, a pressão e a temperatura da câmara do propulsor. Todos os catalisadores foram caracterizados por Adsorção de Nitrogênio, Termogravimetria, Espectroscopia Fotoeletrônica de Raios-X, Difratometria de Raios-X e Resistência Mecânica à Compressão Radial, visando correlacionar suas propriedades físico-químicas com suas atividades na decomposição catalítica do peróxido de hidrogênio concentrado. Os catalisadores denominados MnAl2 e Co4MnAl, sintetizados pelo método da co-precipitação em solução aquosa, foram aqueles que apresentaram os melhores resultados, sendo capazes de decompor espontaneamente o H2O2 sem sofrer desativação ou fragmentação após os testes. / This work aimed to develop mixed oxides, at low cost, to be used as catalysts in the decomposition of hydrogen peroxide, 90% by weight, allowing multiple starts in a microthruster of monopropellant satellites. Different synthesis methods of mixed oxides with high levels of manganese and cobalt oxides were employed. The materials were evaluated, preliminarily, in the spontaneous decomposition of hydrogen peroxide (drop test). Then, the best performing catalysts were selected and tested in a 2N microthruster, where the thrust, the pressure, and temperature in the chamber was monitored. All catalysts were characterized by nitrogen adsorption, thermogravimetry, X-ray photoelectron spectroscopy, X-ray Diffraction, and Mechanical Strength Radial Compression , aiming to correlate their physicochemical properties with their activity in the catalytic decomposition of concentrated hydrogen peroxide.The catalysts called MnAl2 and Co4MnAl, synthesized by co-precipitation in aqueous solution, were those that showed the best results, being able to spontaneously decompose H2O2 without undergoing deactivation or fragmentation after testing.

Este trabalho teve como objetivo o desenvolvimento de materiais catalíticos à base de óxidos mistos, de baixo custo, para serem empregados como catalisadores mássicos, na decomposição do peróxido de hidrogênio 90%, em massa, possibilitando múltiplas partidas a frio em um micropropulsor de satélites a monopropelente. Foram utilizados diferentes métodos de síntese de óxidos mistos com altos teores de manganês e cobalto. Os materiais foram avaliados, preliminarmente, na decomposição espontânea do peróxido de hidrogênio em bancada (teste da gota). Em seguida, os catalisadores com melhores desempenhos foram selecionados e testados em um micropropulsor de 2N, onde foram monitorados o empuxo, a pressão e a temperatura da câmara do propulsor. Todos os catalisadores foram caracterizados por Adsorção de Nitrogênio, Termogravimetria, Espectroscopia Fotoeletrônica de Raios-X, Difratometria de Raios-X e Resistência Mecânica à Compressão Radial, visando correlacionar suas propriedades físico-químicas com suas atividades na decomposição catalítica do peróxido de hidrogênio concentrado. Os catalisadores denominados MnAl2 e Co4MnAl, sintetizados pelo método da co-precipitação em solução aquosa, foram aqueles que apresentaram os melhores resultados, sendo capazes de decompor espontaneamente o H2O2 sem sofrer desativação ou fragmentação após os testes. / This work aimed to develop mixed oxides, at low cost, to be used as catalysts in the decomposition of hydrogen peroxide, 90% by weight, allowing multiple starts in a microthruster of monopropellant satellites. Different synthesis methods of mixed oxides with high levels of manganese and cobalt oxides were employed. The materials were evaluated, preliminarily, in the spontaneous decomposition of hydrogen peroxide (drop test). Then, the best performing catalysts were selected and tested in a 2N microthruster, where the thrust, the pressure, and temperature in the chamber was monitored. All catalysts were characterized by nitrogen adsorption, thermogravimetry, X-ray photoelectron spectroscopy, X-ray Diffraction, and Mechanical Strength Radial Compression , aiming to correlate their physicochemical properties with their activity in the catalytic decomposition of concentrated hydrogen peroxide.The catalysts called MnAl2 and Co4MnAl, synthesized by co-precipitation in aqueous solution, were those that showed the best results, being able to spontaneously decompose H2O2 without undergoing deactivation or fragmentation after testing.

"As new, smaller satellites are built, the need for improved on-board propulsion systems has grown. The pulsed plasma thruster has received attention due to its low power requirements, its simple propellant management, and the success of initial flight tests. Successful integration of PPTs on spacecraft requires the comprehensive evaluation of possible plume-spacecraft interactions. The PPT plume consists of neutrals and ions from the decomposition of the Teflon propellant, material from electrode erosion, as well as electromagnetic fields and optical emissions. To investigate the PPT plume, an on-going program is underway at WPI that combines experimental and computational investigations. Experimental investigation of the PPT plume is challenging due to the unsteady, pulsed as well as the partially ionized character of the plume. In this thesis, a triple Langmuir probe apparatus was designed and used to obtain electron temperature and density measurements in the plume of a PPT. This experimental investigation provides further characterization of the plume, much needed validation data for computational models, and is useful in thruster optimization studies. The pulsed plasma thruster used in this study is a rectangular geometry laboratory model built at NASA Lewis Research Center for component lifetime tests and plume studies. It is almost identical in size and performance to the LES 8/9 thruster, ablating 26.6 ug of Teflon, producing an impulse bit of 256 uN-s and a specific impulse of 986 s at 20 J. All experiments were carried out at NASA LeRC Electric Propulsion Laboratory. The experimental setup included triple Langmuir probes mounted on a moveable probe stand, to collect data over a wide range of locations and operating conditions. Triple probes have the ability to instantaneously measure electron temperature and density, and have the benefit of being relatively simple to use, compared to other methods used to measure these same properties. The implementation of this measuring technique is discussed in detail, to aid future work that utilizes these devices. Electron temperature and density was measured from up to 45 degrees from the centerline on planes parallel and perpendicular to the thruster electrodes, for thruster energy levels of 5, 20 and 40 J. Radial distances extend from 6 to 20 cm downstream from the Teflon surface. These locations cover the core of the PPT plume, over a range of energy levels that corresponds to proposed mission operating conditions. Data analysis shows the spatial and temporal variation of the plume. Maximum electron density near the exit of the thruster is 1.6 x 1020, 1.6 x 1021, and 1.8 x 1021 m-3 for the 5, 20 and 40 J discharges, respectively. At 20 cm downstream from the Teflon surface, densities are 1 x 1019, 1.5 x 1020 and 4.2 x 1020 for the 5, 20 and 40 J discharges, respectively. The average electron temperature at maximum density was found to vary between 3.75 and 4.0 eV for the above density measurements at the thruster exit, and 20 cm from the Teflon surface the temperatures are 0.5, 2.5, and 3 eV for the 5, 20 and 40 J discharges. Plume properties show a great degree of angular variation in the perpendicular plane and very little in the parallel plane, most likely due to the rectangular geometry of the PPT electrodes. Simultaneous electron temperature and density traces for a single thruster discharge show that the hottest electrons populate the leading edge of the plume. Analysis between pulses shows a 50% variation in density and a 25% variation in electron temperature. Error analysis estimates that maximum uncertainty in the temperature measurements to be approximately +/- 0.75 eV due to noise smoothing, and the maximum uncertainty in electron density to be +/- 60%, due to assumptions related to the triple probe theory. In addition, analysis of previously observed slow and fast ion components in the PPT plume was performed. The analysis shows that there is approximately a 3 us difference in creation time between the fast and slow ions, and that this correlates almost exactly with the half period of the oscillations in the thruster discharge current."

Automatic altitude compensation has been a holy grail of rocket propulsion for decades. Current state-of-the-art bell nozzles see large performance decreases at low altitudes, limiting rocket designs, shrinking payloads, and overall increasing costs. Aerospike nozzles are an old idea from the 1960’s that provide superior altitude-compensating performance and enhanced performance in vacuum, but have survivability issues that have stopped their application in satellite propulsion systems. A growing need for CubeSat propulsion systems provides the impetus to study aerospike nozzles in this application. This study built two aerospike nozzles using modern 3D metal printing techniques to test aerospikes at a size small enough to be potentially used on a CubeSat. Results indicated promising in-space performance, but further testing to determine thermal limits is deemed necessary.