"A study is presented detailing the simulation of a drag-free follow-on mission to NASA’s Gravity Recovery and Climate Experiment (GRACE). This work evaluates controller performance, as well as thrust, power, and propellant mass requirements for drag-free spacecraft operation at orbital altitudes of 160 - 225 kilometers. In addition, sensitivities to thermospheric wind, GPS signal accuracy and availability of ephemeris data are studied. Orbital dynamics were modeled in Matlab and take into account 2 body gravity effects, J2-J6 non-spherical Earth effects, atmospheric drag and control thrust. A drag model is used in which the drag acceleration is a function of the spacecraft’s relative velocity to the atmosphere, and a “drag parameter,” which includes the spacecraft’s drag coefficient and local mass density of the atmosphere. A MSISE-90 atmospheric model is used to provide local mass densities as well as free stream flow conditions for a Direct Simulation Monte Carlo drag analysis used to validate the spacecraft drag coefficient. The controller is designed around an onboard inertial sensor which uses a freely floating reference mass to measure deviations in the spacecraft position, resulting from non-gravitational forces, from a desired target orbit. Thruster (control actuator) models are based on two different Hall thrusters for providing the orbital along-track acceleration, colloid thrusters for the normal acceleration, and a miniature xenon ion thruster (MiXI) for the cross-track acceleration. The most demanding propulsion requirements correspond to the lowest altitude considered, 160 kilometers. At this altitude the maximum along-track thrust component is calculated to be 98 millinewtons with a required dynamic (throttling) response of 41 mN/s. The maximum position error at this altitude was shown to be in the along-track direction with a magnitude of 3314.9 nanometers and a peak spectral content of 1800 nm/sqrt(Hz) at about 0.1 Hz. At 225 kilometers, the maximum along-track thrust component reduces to 10.3 millinewtons. The maximum dynamic response at this altitude is 4.23 mN/s. The maximum along-track position error is reduced to 367.9 nanometers with a spectral content peak of 40 nm/sqrt(Hz) at 0.1 Hz. For all altitudes, the maximum state errors increase as the mission length increases, however, higher altitude missions show less of a maximum displacement error increase over time than those of lower orbits. The ability of a colloid thruster to control the normal drift is found to be dependent on how frequently the spacecraft state data is updated. Reducing the period between updates from 10 seconds to 1 second reduces the maximum normal state error component from 199 nanometers to less than 32 nanometers, suggesting that spacecraft state update frequency could be a major driver in keeping the spacecraft on the target trajectory. Sensitivity of maximum required thrust and accumulated sensor error to measurement uncertainty is found to be less of a driver than state update frequency. A ‘worst case” thermospheric wind gust was modeled to show the increase on propulsion requirements if such an event were to occur. At 200 kilometers, maximum winds have been measured to be in increase of 650 m/s in the westward direction in the southern pole region. Assuming the majority of the 650 m/s gust occurs over a 4 second time span, the maximum required cross-track thrust at 200 kilometers increases from 1.12 to 2.01 millinewtons. This large increase may drive the thruster choice for a drag-free mission at a similar altitude. For the spacecraft point design considered with a propellant mass fraction of 0.18, the mission lifetime for the 160 km case was calculated to be 0.76 years. This increases 2.27 years at an altitude of 225 km."

"The development of Retarding Potential Analyzers (RPAs) capable of measuring high-density stationary and flowing plasmas is presented. These new plasma diagnostics address the limitations of existing RPAs and can operate in plasmas with electron densities in excess of 1x1018 m-3. Such plasmas can be produced by high-powered Hall Thrusters, Pulsed Plasma Thrusters (PPTs), and other plasma sources. The Single-Channel micro-Retarding Potential Analyzer (SC-microRPA) developed has a minimum channel diameter of 200 microns, electrode spacing on the sub-millimeter scale and can operate in plasmas with densities of up to 1x1017 m-3. The electrode series consists of 100 micron thick molybdenum electrodes and Teflon insulating spacers. The alignment process of the channel, as well as the design and fabrication of the stainless steel outer housing, the Delrin insulating tube, and all other microRPA components are detailed. To expand the applicability of the SC-microRPA to densities above 1x1018 m-3 a low transparency Microchannel Plate (MCP) has been incorporated in the design of a Multi-Channel micro-Retarding Potential Analyzer (MC-microRPA). The current collection theory for the SC-microRPA and the MC-microRPA is also derived. The theory is applicable to microRPAs with arbitrary channel length to diameter ratios and accounts for the reduction of ion flux due to the microchannel plate in the case of the MC-microRPA, due to absorption of ions by channel walls, and due to the applied potential. Current-voltage curves are obtained for incoming plasma flows that range from near-stationary to hypersonic, with temperatures in the range of 0.1 to 10 eV, and densities in the range of 1x1015 m-3 to 1x1021 m-3. The SC-microRPA current collection theory is validated by comparisons with the classical RPA theory and particle-in-cell simulations. Determination of unknown plasma properties is based on a fuzzy-logic approach that uses the generated current-voltage curves as lookup tables."

published_or_final_version / Electrical and Electronic Engineering / Doctoral / Doctor of Philosophy

In this thesis, the analysis of the dynamic response of a Linear Induction Motor as an electromechanical system is done, accounting for all the governing equations implied in the process which are used to develop the corresponding simulation models. Once this model is presented, a feedback control system is implemented in order to analyze the controlled response of the motor, considering the applications and conditions analogue to aircraft launcher systems. Also a comparison between the Linear and Rotary induction motors describing the differences, similarities and equivalences will be developed. / by Armando Josâe Sinisterra. / Thesis (M.S.C.S.)--Florida Atlantic University, 2011. / Includes bibliography. / Electronic reproduction. Boca Raton, Fla., 2011. Mode of access: World Wide Web.

Electric propulsion for spacecraft has become increasingly commonplace in recent decades as designers take advantage of the significant propellant savings it can provide over traditional chemical propulsion. As electric propulsion systems are designed for very low thrust, the operational time required over the course of an entire mission is often quite long. The two most common types of electric thrusters both use hollow cathodes as electron emitters in the process of ionizing the propellant gas. These cathodes are one of the main life-limiting components of both ion and Hall thrusters designed to operate for tens of thousands of hours. Failure often occurs as a result of erosion by sputtering from high-energy ions generated in the plasma. The mechanism that is responsible for creating these high-energy ions is not well understood, and significant efforts have gone into characterizing the plasma produced by hollow cathodes. This work uses both a Langmuir probe and an emissive probe to characterize the variation of the plasma potential and density, the electron temperature, and the electron energy distribution function in the near plume region of a hollow cathode. The cathode used in this experiment is typical of one used in a 200-W class Hall thruster. Measurements were made to determine the variation of these parameters with radial position from the cathode orifice. Changes associated with varying the propellant and flow rate were also investigated. Results obtained from the cathode while running on both argon and xenon are shown. Two different methods for calculating the plasma density and electron temperature were used and are compared. The density and temperature were not strongly affected by reductions in the propellant flow rate. The electron energy distribution functions showed distinct shifts toward higher energies when the cathode was operated at lower flow rates. The plasma potential also displayed an abrupt change in magnitude near the cathode centerline. Significant increases in the magnitude of plasma potential oscillations at lower propellant flow rates were observed. Ions formed at the highest instantaneous plasma potentials may be responsible for the life-limiting erosion that is observed during long-duration operation of hollow cathodes.

Thesis (M.S.)--Worcester Polytechnic Institute. / Keywords: Langmuir probes; spacecraft; electric propulsion. Includes bibliographical references (p.69-71).

Le propulseur PEGASES (Plasma Propulsion with Electronegative Gases) est un nouveau type de propulseur électrique pour la propulsion spatiale. Il utilise des ions négatifs et positifs créés par une décharge radiofréquence à couplage inductif pour générer la poussée. L’accélération électrostatique des ions est assurée par un ensemble de grilles polarisées. Un filtre magnétique est utilisé pour augmenter la quantité d'ions négatifs dans la cavité du propulseur. Le propulseur PEGASES est non seulement une source qui permet de créer un plasma d'ions négatifs à forte densité, et même un plasma d'ion-ion, mais il peut également être utilisé comme un propulseur ionique classique. Cela signifie qu'un plasma est créé dans un gaz électropositif (e.g. Xe) et que les ions positifs sont extraits et accélérés. Dans ce cas, il est nécessaire de neutraliser le plasma derrière la zone d'accélération, comme dans d'autres propulseurs ioniques. Les performances du propulseur PEGASES ont été étudiées principalement dans du xénon afin de comparer les résultats obtenus avec les propulseurs ioniques de type RIT. Le propulseur a été étudié à l'aide d'une série de sondes telles qu’une sonde de Langmuir, une sonde plane, une sonde capacitive et un RPA (pour Analyseur à Champ Retardateur). De plus, une sonde en champs croisés ExB a été développée pour mesurer la vitesse des ions quittant le propulseur ainsi que la fraction des différentes espèces ioniques présentes dans le plasma. / The PEGASES thruster (Plasma Propulsion with Electronegative Gases) is a novel type of electric thruster for space propulsion. It uses negative and positive ions produced by an inductively coupled radio frequency discharge to create the thrust by electrostatically accelerating the ions through a set of grids. A magnetic filter is used to increase the amount of negative ions in the cavity of the thruster. The PEGASES thruster is not only a source to create a strongly negative ion plasma or even an ion-ion plasma but it can also be used as a classical ion thruster. This means that a plasma is created and only the positive ions are extracted and accelerated making it necessary to neutralize the plasma behind the acceleration stage like in other ion thrusters. The performances of the PEGASES thruster have been investigated mainly in xenon in order to compare the obtained results with RIT-type ion thrusters. The thruster has been investigated with the help of a variety of probes such as a Langmuir probe, a planar probe, a capacitive probe and a RPA (Retarding Potential Analyzer). In addition, an ExB probe has been developed to measure the velocity of the ions leaving the thruster and to differentiate between the ion species present in the plasma.

The need for micropropulsion solutions for spacecraft has been steadily increasing as scientific payloads require higher accuracy maneuvers and as the use of small form-factor spacecraft such as CubeSats becomes more common. Of the technologies used for this purpose, electrospray thrusters offer performance that make them an ideal choice. Electrosprays offer high accuracy impulse bits at low power and high efficiency, and have low volume requirements. Design choice reasoning and preliminary testing results are presented for two electrospray thruster designs. The first thruster, named the Demonstration thruster, is operated in atmospheric conditions and serves as a highly visible example of the basic concepts of electrospray technology applied to micropropulsion. It features a single capillary needle emitter and the acetone propellant flow is driven actively by a syringe pump. The second thruster, named the Research thruster, is operated in the vacuum environment and is designed for modularity for its expected use in future research efforts. Propellant flow is also driven actively using a syringe pump. Initial configuration of the Research thruster is a linear array of five capillary needle emitters, though testing is conducted with only one emitter in this thesis. Tests using un-doped glycerol and sodium iodide doped glycerol (20% by weight) are conducted for the Research thruster. Both thruster designs use stainless steel 18 gauge blunt dispensing needles (0.038 in / 0.965 mm ID) as their emitters. Applied voltage to the emitter(s) relative to the grounded extractor is swept from 2100 V to 3700 V for the Demonstration thruster testing and from 4000 V to 4500 V for the Research thruster. Currents incident on a collection plate downstream of the emission plume and on the extractors of the thrusters were measured directly with a pico-ammeter. Measurements made during testing of the Demonstration thruster are inconsistent due to charge loss as propellant travels through the air, though currents as high as 5.1x10-9 A on the collection plate and 2x10-7 A on the extractor are recorded. Currents for Research thruster testing using un-doped glycerol were measured as high as 4.9x10-8 A on the collection plate and 5x10-9 A on the extractor, showing an interception rate as high as 17%. Currents using sodium iodide doped glycerol were measured as high as 7x10-7 A on the collection plate. Discussion is given for the visual qualities of cone-jet emission for all testing. Keywords:

With an increased demand in Cube Satellite (CubeSat) development for low cost science and exploration missions, a push for the development of micro-propulsion technology has emerged, which seeks to increase CubeSat capabilities for novel mission concepts. One type of micro-propulsion system currently under development, known as Pocket Rocket, is an electrothermal plasma micro-thruster. Pocket Rocket uses a capacitively coupled plasma, generated by radio-frequency, in order to provide neutral gas heating via ion-neutral collisions within a gas discharge tube. When compared to a cold-gas thruster of similar size, this gas heating mechanism allows Pocket Rocket to increase the exit thermal velocity of its gaseous propellant for increased thrust. Previous experimental work has only investigated use of the gas discharge tube's orifice for propellant expansion into vacuum. This thesis aims to answer if Pocket Rocket may see an increase in thrust with the addition of a micro-nozzle, placed at the end of the gas discharge tube. With the addition of a conical ε = 10, α = 30° micro-nozzle, performance increases of up to 6% during plasma operation, and 25% during cold gas operation, have been observed. Propellant heating has also been observed to increase by up to 60 K within the gas discharge tube.

No description available.