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  • About
  • The Global ETD Search service is a free service for researchers to find electronic theses and dissertations. This service is provided by the Networked Digital Library of Theses and Dissertations.
    Our metadata is collected from universities around the world. If you manage a university/consortium/country archive and want to be added, details can be found on the NDLTD website.
11

Anode fall as relevant to plasma thrusters

Horner, Brigitte 06 1900 (has links)
The behavior of the electric field together with the electron and ion densities in the vicinity of a nonemitting, plane anode is investigated. The selected approach involves non-linear analysis techniques on the continuum equations for steady-state, isothermal conditions where both ionization and two-body recombination are included. Ions, created through electron bombardment of neutral atoms, are repelled toward two stagnation regions: within or near the sheath boundary and near the plasma interface. These equilibria form as a result of the chemistry present: recombination establishes the latter while ionization stipulates the former. As presented, the sheath is fundamentally unstable - ions are driven toward the negative electrode. Using nitrogen data for a numeric example, the following observations are made: a sufficiently strong applied electric field pushes the ion density toward that ofthe electrons through a well - a constrictive phenomenon. Both a transition region, dominated by density gradients, and a diffusion-driven zone are found to move the system toward the plasma interface. The characteristics of this process are influenced by the applied electric field, but the instability of the chemistry-induced stagnation regions precludes numeric convergence. Insufficient dissipation may prevent the stability of the anode fall model as presented. Suggested improvements to the model descriptions include considering the effects of temperature gradients, magnetic fields, three-body recombination, diffusion written in terms of the electric field, multi-dimensionality and/or timedependencies^
12

Very low earth orbit propellant collection feasibility assessment

Singh, Lake Austin 12 January 2015 (has links)
This work focuses on the concept of sustainable propellant collection. The concept consists of gathering ambient gas while on-orbit and using it as propellant. Propellant collection could potentially enable operation in very-low Earth orbits without compromising spacecraft lifetime. This work conducts a detailed analysis of propellant collection from a physics perspective in order to test the assertions of previous researchers that propellant collection can dramatically reduce the cost of propellant on-orbit. Major design factors for propellant collection are identified from the fundamental propellant collection equations, which are derived in this work from first principles. A sensitivity analysis on the parameters in these equations determines the relative importance of each parameter to the overall performance of a propellant-collecting vehicle. The propellant collection equations enable the study of where propellant collection is technically feasible as a function of orbit and vehicle performance parameters. Two case studies conducted for a very-low Earth orbit science mission and a propellant depot-type mission serve to demonstrate the application of the propellant collection equations derived in this work. The results of this work show where propellant collection is technically feasible for a wide range of orbit and vehicle performance parameters. Propellant collection can support very-low Earth operation with presently available technology, and a number of research developments can further extend propellant-collecting concepts' ability to operate at low altitudes. However, propellant collection is not presently suitable for propellant depot applications due to limitations in power.
13

Evaluation and comparison of electric propulsion motors for submarines /

Harbour, Joel P. January 2001 (has links)
Thesis (Naval Engineer and M.S. in Electrical Engineering and Computer Science)--Massachusetts Institute of Technology, 2001. / Includes bibliographical references (p. 100-106). Also available online.
14

Langmuir probe measurements in the plume of a pulsed plasma thruster

Byrne, Lawrence Thomas. January 2002 (has links)
Thesis (M.S.)--Worcester Polytechnic Institute. / Keywords: PPT; pulsed plasma thruster; Langmuir probe; plasma diagnostics; electric propulsion; electron temperature; electron density. Includes bibliographical references (p. 97-102).
15

Optimization of a magnetoplasmadynamic arc thruster

Krolak, Matthew Joseph. January 2007 (has links)
Thesis (M.S.) -- Worcester Polytechnic Institute. / Keywords: Electric Propulsion; Plasma thruster; MPD. Includes bibliographical references (leaves 6-13).
16

Characterization of a Low Current LaB6 Heaterless Hollow Cathode with Krypton Propellant

Jain, Prachi Lalit 25 June 2020 (has links)
A first-generation LaB6 heaterless hollow cathode with a flat-plate anode is experimentally investigated. The cathode is characterized using krypton as propellant at varying flow rates, discharge currents and cathode-anode distances. Voltage probes, used to make direct voltage measurements in the ignition circuit, are the only diagnostic tool used experimentally. A plasma model is used to infer plasma parameters in the cathode emitter region. The cathode characterization results are consistent with those obtained during previous investigations of 1 A-class LaB6 hollow cathode with krypton. A peak-to-peak anode voltage criterion is used to identify the discharge modes and the occurrence of mode transition. Fourier analysis of the keeper and anode voltage waveforms carried out to study the discharge mode behavior reveals resonant frequencies ranging from 40 to 150 kHz. Lastly, post-test visual observations of the cathode components show signs of emitter poisoning and keeper erosion. / Master of Science / Recent years have seen rapid growth in the development of both stand-alone satellites and satellite constellations. A critical component of these satellites is the on-board propulsion system, which is responsible for controlling their orientation with respect to the object of interest and keeping the spacecraft in the assigned orbit. Generally, electric propulsion systems are used for this purpose. These types of propulsion systems use electrical power to change the velocity of satellite, providing a small thrust for a long duration of time as compared to chemical propulsion systems. Certain types of electric thrusters utilize a hollow cathode device as an electron source to start-off and support the thruster operation. In this research, a non-conventional hollow cathode for low power applications is developed and tested. The main characteristic of the developed cathode is the heaterless configuration, which eliminates the heater module used in conventional cathodes to enable the cathode to reach its operational temperature. The absence of a heater reduces the complexity of the cathode and the electrical power system. The cathode utilizes an electron emitter material which is insensitive to impurities and air exposure. Additionally, unlike typical electric thrusters which use xenon as the fuel, this cathode uses krypton which is similar to xenon but is less expensive. The presented work includes an overview of electric propulsion and the hollow cathode operation, followed by a detailed discussion of the heaterless hollow cathode design, the experimental setup and the test results. Several noteworthy findings regarding cathode operation are included as well. This research shows that the non-conventional heaterless hollow cathode and its operation with krypton have the potential to improve the overall thruster performance by reducing the weight and the cost, thus contributing to an integral aspect of satellite on-board propulsion.
17

Lunar Robotic Precursor Missions Using Electric Propulsion

Winski, Richard G. 05 January 2007 (has links)
A trade study is carried out for the design of electric propulsion based lunar robotic precursor missions. The focus is to understand the relationships between payload mass delivered, electric propulsion power, and trip time. The results are compared against a baseline system using chemical propulsion with LOX/H2. The major differences between the chemical propulsion based and electric propulsion based systems are presented in terms of the payload mass and trip time. It is shown that solar electric propulsion offers significant advantage over chemical propulsion in delivering non-time critical payloads to lunar orbit. / Master of Science
18

Modeling Differential Charging of Composite Spacecraft Bodies Using the Coliseum Framework

Barrie, Alexander 10 October 2006 (has links)
The COLISEUM framework is a tool designed for electric propulsion plume interactions. Virginia Tech has been developing a module for COLISEUM called DRACO, a particle-in-cell based code capable of plume modeling for geometrically complex spacecraft. This work integrates a charging module into DRACO. Charge is collected via particle impingement on the spacecraft surface and converted to potential. Charge can be stored in the surface, or added to a local ground potential. Current can flow through the surface and is governed by the internal electric field in the spacecraft. Several test cases were run to demonstrate the code's capabilities. The first was a plume impingement of a composite spherical probe by a xenon thruster. It was shown that the majority of current conducted will reach the interior of the spacecraft, not other surface elements. A conductive interior will therefore result in a uniform surface potential, even for low surface conductivities. A second test case showed a composite spacecraft exposed to a solar wind. This test showed that when a potential gradient is applied to a conductive body, the ground potential of the spacecraft will lower significantly to compensate and maintain a zero net current. The case that used the semiconductive material showed that the effect of the potential gradient can be lowered in cases such as this, where some charge will always be stuck in the surface. If a dielectric material is used, the gradient will disappear altogether. The final test case showed the effect of charge exchange ion backflow on the potential of a spacecraft similar to the DAWN spacecraft. This case showed that CEX ion distribution is not even along the spacecraft and will result in a transverse potential gradient along the panel. / Master of Science
19

Validation of the DRACO Particle-in-Cell Code using Busek 200W Hall Thruster Experimental Data

Spicer, Randy Lee 30 August 2007 (has links)
This thesis discusses the recent developments to the electric propulsion plume code DRACO as well as a validation and sensitivity analysis of the code using data from an AFRL experiment using a Busek 200 W Hall Thruster. DRACO is a PIC code that models particles kinematically while using finite differences schemes to solve the electric potential and field. The DRACO code has been recently modified to improve simulation results, functionality and performance. A particle source has been added that uses the Hall Thruster device code HPHall as input for a source to model Hall Thrusters. The code is now also capable of using a non-uniform mesh that uses any combination of uniform, linear and exponential stretching schemes in any of the three directions. A stretched mesh can be used to refine simulation results in certain areas, such as the exit of a thruster, or improve performance by reducing the number of cells in a mesh. Finally, DRACO now has the capability of using a DSMC collision scheme as well as performing recombination collisions. A sensitivity analysis of the newly upgraded DRACO code was performed to test the new functionalities of the code as well as validate the code using experimental data gathered at AFRL using a Busek 200 W Hall Thruster. A simulation was created that attempts to numerically recreate the AFRL experiment and the validation is performed by comparing the plasma potential, polytropic temperature, ion number density of the thruster plume as well as Faraday and ExB probe results. The study compares the newly developed HPHall source with older source models and also compares the variations of the HPHall source. The field solver and collision model used are also compared to determine how to achieve the best results using the DRACO code. Finally, both uniform and non-uniform meshes are tested to determine if a non-uniform mesh can be properly implemented to improve simulation results and performance. The results from the validation and sensitivity study show that the DRACO code can be used to recreate a vacuum chamber simulation using a Hall Thruster. The best results occur when the newly developed HPHall source is used with a MCC collision scheme using a projected background neutral density and CEX collision tracking. A stretched mesh was tested and proved results that are as accurate as a uniform mesh, if not more accurate in locations of high mesh refinement. / Master of Science
20

Thermal-Fluid Analysis of a Lithium Vaporizer for a High Power Magnetoplasmadynamic Thruster

St. Rock, Brian Eric 09 January 2007 (has links)
A lithium vaporizer for a high-power magnetoplasmadynamic (MPD) thruster is modeled using a parallel approach. A one-dimensional, thermal-resistive network is developed and used to calculate the required vaporizer length and power as a function of various parameters. After comparing results predicted by this network model with preheat power data for a 200 kW MPD thruster, we investigate performance over a parameter space of interest for the Advanced Lithium-Fed, Applied-field, Lorentz Force Accelerator (ALFA2) thruster. Heater power sensitivity to cathode tube emissivity, mass flow rate, and vapor superheat are presented. The cold-start heater power for 80 mg/s is found to range from 3.38 to 3.60 kW, corresponding to a vaporizer (axial) length of 18 to 26 cm. The strongest drivers of vaporizer performance are cathode tube emissivity and a conduction heat sink through the mounting flange. Also, for the baseline ALFA2 case, it is shown that increasing the vapor superheat from 100 K to 300 K has the effect of lowering the vaporizer thermal efficiency from 57% to 49%. Also, a finite-volume computational fluid dynamic (CFD) is implemented in FLUENT 6.2 which includes conjugated heat transfer to the solid components of the cathode. This model uses a single-fluid mixture model to simulate the effects of the two-phase vaporizer flow with source terms that model the vaporization. This model provides a solution of higher fidelity by including three-dimensional fluid dynamics such as thermal and momentum boundary layers, as well as calculating a higher resolution temperature distribution throughout the cathode assembly. Results from this model are presented for three mass flow rates of interest (40 mg/s, 80 mg/s, and 120 mg/s). Using a fixed power and length taken from the conceptual ALFA2 design, the dryout point ranges from 12.3-17.6 cm from the base of the cathode assembly for 40 mg/s and 80 mg/s, respectively. For the 120 mg/s case, the two-phase flow never reaches dryout. Finally, results two modeling approaches are compare favorably, with a maximum disagreement of 13.0 percent in prediction of the dryout point and 4.2 percent in predictions of the exit temperature.

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