Master of Engineering, Case Western Reserve University, 2025, EMC - Aerospace Engineering

Analysis of three missions has been carried out with a set of three power and three propulsion systems to determine system synergy as well as to find an optimal system for each mission. The three missions are GTO to LLO, LMO, and Titan flyby. The three propulsion systems of analysis are Hall Effect Thrusters, Magnetoplasmadynamic Thrusters, and VASIMR Thrusters. The three power systems of analysis are Silicon photovoltaics, multi-junction photovoltaics, and nuclear reactors. The mission to Low Lunar Orbit has a maximum trip time of 8 weeks, and four systems are capable of achieving this result. Those systems are Hall-Nuclear, MPDNuclear, VASIMR-MJ, and VASIMR-Nuclear, the last of which achieves the target in just 26 days. The VASIMR-Nuclear system is also capable of bringing the most passengers with a total capacity of 255 people. The mission to LMO was limited to a maximum trip time of 9 months, and 1 system is capable of achieving this result. This system is again VASIMR-Nuclear, capable of bringing 31 people the LMO in 241 days, or about 8 months. The mission to Titan flyby using the VASIMR-Nuclear system is capable of bringing 54500 kg of dry mass to Titan flyby at 1000 km at 3.13 km/s relative to the planet.

Committee: Paul Barnhart (Advisor); Majid Rashidi (Committee Member); Richard Bachmann (Committee Member) Subjects: Aerospace Engineering

Master of Sciences (Engineering), Case Western Reserve University, 2024, EMC - Aerospace Engineering

As deep space missions expand in scope and distance, the efficiency of propulsion systems becomes paramount. This thesis analyzes the impact of small measurement errors in the thrust profiles of Hall-Effect Thrusters, a common type of electric propulsor, known for increased efficiency compared to traditional chemical propulsion despite lower thrust. Due to their prolonged operational times, these errors compound, affecting trajectory and mission success. Through analyzing the AEPS Hall-Effect Thruster prototype, designed for NASA's Artemis Program's Gateway space station, using curve fitting and a Monte Carlo simulation, we assess the effects of these errors on an example mission to Alpha Centauri. Results show plasma dynamics cause the majority of error, but cause minimal trajectory deviation and propellant loss. This reinforces electric propulsion's suitability for long-distance space travel. This work informs spacecraft mission design, providing valuable insights into fuel efficiency and system selection, as well as building upon NASA Glenn Research Center's Electric Propulsion and Power Laboratory's prior research.

Committee: Bryan Schmidt (Committee Chair); John Yim (Other); Richard Bachmann (Committee Member); Paul Barnhart (Committee Member) Subjects: Aerospace Engineering; Engineering; Mechanical Engineering

Master of Science in Mechanical Engineering, Cleveland State University, 2018, Washkewicz College of Engineering

Currently there is a large interest in the use of more efficient means of propulsion in long term missions due to the costs and difficulties associated with placing and maintaining the needed fuel for conventional chemical systems in orbit. Mass reduction of upper stages will return large returns due to the great reduction in required lower stage fuel. Due to these factors, alternatives are undergoing active research, though this paper is concerned with the area of electrical propulsion. Electric propulsion is broadly defined as propulsion where the energization of the exhaust occurs via application of electromagnetic fields as opposed to chemical reactions or thermal processes. Frequently plasmas are involved in such processes, and as such, diagnostics related to establishing the characteristics of plasmas are of great value to the field, especially any techniques which do not disrupt the plasma in testing. This study focused on the use of non-intrusive optical techniques to measure electron temperature in the near field of a Hall thruster, a type of electric propulsion. Results indicate that the pursued technique may be of utility to providing a simple to set up and execute diagnostic for determining electron temperatures in the near field of a hall thruster. This will allow for less-expensive rapid turnaround on obtaining of experimental data to test or provide as input for models of plasma behavior, wear in electric propulsion, thruster design, and operation confirmation.

Committee: Antonie van den Bogert Dr. (Committee Chair); Wei Zhang Dr. (Committee Member); Asuquo Ebiana Dr. (Committee Member); George Williams Dr. (Advisor) Subjects: Aerospace Engineering; Engineering; Mechanical Engineering

Master of Science in Engineering (MSEgr), Wright State University, 2006, Mechanical Engineering

Enhancing the confinement of primary electrons within the plasma in a discharge chamber of an ion thruster improves plasma ionization and consequently the thruster's performance. This work computationally calculates the location, position, and orientation of the permanent magnets that provide a ring-cusp magnetic field that maximizes electron confinement in an axi-symmetric cylindrical aluminum-wall discharge chamber. Small samarium cobalt magnets are circumferentially arranged in a ring around the front, side, or back wall of the chamber. The generated ring-cusp magnetic field for any specified magnet configuration is calculated using MAXWELL2D, a two dimensional electromagnetic field simulation computer code. For various magnet configurations, PRIMA, a particle-in-cell computer code modified by Mahalingam and Menart, is used to model the trajectory of the primary electrons in the magnetic field. The confinement length, the length of time an electron is retained within the chamber, is output by PRIMA, and it is the parameter used to determine the performance of the magnet configurations surveyed. The performance of various magnet ring pairs are studied and guidelines on the location, position, and orientation of the magnet rings are obtained. These guidelines are then combined to give complex ring-cusp magnet ring arrangements on a fixed size discharge chamber. For three complex arrangements having three magnet rings, a decrease in the chambers confinement ability is seen when the applied guidelines are slightly violated. This observed decrease validates the guidelines deduced in this work.

Committee: James Menart (Advisor) Subjects:

Master of Science, The Ohio State University, 2012, Electrical and Computer Engineering

This thesis is concerned with a problem of balancing the humanoid robot after an external impact. Dynamic model of the humanoid robot is derived using Lagrangian dynamic formulation. Use of the maximum joint accelerations to reject disturbance is studied. In our approach, we propose the use of non-natural force like thruster on the torso of the humanoid robot for balance recovery. Mathematical simulation of derived dynamic model is performed using MATLAB. Plotted results prove the validity and usefulness of the proposed approach. We also show that, acceleration compensation and using thruster are complementary to each other. We prove that both techniques can be used together to reject large disturbances in minimum time.

Committee: Dr. Yuan Zheng (Advisor); Dr. Hooshang Hemami (Committee Member) Subjects: Electrical Engineering; Engineering; Mechanical Engineering; Robotics; Robots

Doctor of Philosophy, The Ohio State University, 2007, Aeronautical and Astronautical Engineering

Decay of ion thruster grids due to impact by charge exchange ions is the main life limiting factor in ion propulsion systems. Any system which can reduce the number or energy of ions impacting the grids will add to the life expectancy at current power levels. One possible technique for reducing damage from charge exchange ions would involve the incorporation in the grids of axial aligned embedded magnetic fields. These fields, generated by currents running around the grid apertures would form mini magnetic nozzles guiding beam ions through the aperture while diverting charge exchange ions from directly impacting the grids. A simple computational simulation of the environment within a single set of ion thruster grids has been created for use in evaluating the response of single charge exchange ions to many different grid geometries and field arrangements. This effort involved the development of the ionLite code which simulates the grid geometries, fields, and ion beam charge distribution while allowing individual charge exchange ions to be created at any point and their trajectories and eventual fates to be determined. Using this code the use of auxiliary magnetic fields was examined. This analysis shows that energy transfer to a simulated accelerator grid from charge exchange ions can be reduced by approximately 20%, but only at vary large magnitude magnetic field strengths (order of 100 T). It was found that for the configurations investigated the optimum performance resulted when the applied magnetic field was just enough to cause the particle Larmor radius to be approximately equal to the grid aperture radius. The use of lower mass propellants such as Neon or Helium allow for this benefit at fields on the order of 20 T. The potential impact of the embedded magnetic fields is shown to be very sensitive to grid geometry, and therefore it is probable that a different configuration could provide even greater reduction in kinetic energy transfer at moderate field leve (open full item for complete abstract)

Committee: Thomas York (Advisor) Subjects: Engineering, Aerospace

Doctor of Philosophy, The Ohio State University, 2003, Aeronautical and Astronautical Engineering

Characteristic measurements were made of a hundred megawatt modified helium inverse pinch switch and compared against numerical modeling and theoretically expected behavior. Thruster voltage was measured for currents between three and three hundred kilo amps and for mass flow rates between 0.96 and 40 grams per second. From that, characteristic voltage, power, and resistance curves were generated. Electron temperature measurements made inside the plasma flow were found to be between three and thirty electron volts. General expected behavior, such as decreasing resistance with increasing mass flow rate, were confirmed. The quasi steady assumption was studied between 1.5 and 1.7 milliseconds and found to be appropriate. A theoretical model, based on normal MPD thrust behavior, was used to estimate fall voltages and pumping coefficients. An empirical model for thruster voltage was then created to estimate the behavior of voltage as a function of the similarity parameter. The two models were then put together and found to be self consistent with the experimental data. Total temperatures, specific impulses, and efficiencies for assumed isentropic nozzle expansion were then calculated.

Committee: Gerald Gregorek (Advisor) Subjects: Engineering, Aerospace

Doctor of Philosophy, The Ohio State University, 2003, Aeronautical and Astronautical Engineering

Analytic models predict the possibility of extending the range of performance parameters of Pulsed Plasma Thrusters (PPT) by using propellants other than the traditionally used Teflon. A theoretical and experimental effort was initiated at The Ohio State University to investigate the use of alternative propellants for PPT. Analytical and numerical calculations (MACH2) indeed indicate a significant broadening of the obtainable range of specific impulse and thrust-to-power ratios when alternative propellants such as lithium or water are utilized. Consequently, in an effort to investigate changes in physical phenomena and thruster performance experimentally, a hybrid thruster was designed and built, facilitating the use of alternatively water or Teflon. The thruster design includes a unique water propellant feed system, allowing the supply of the water propellant without detrimentally affecting the inherent simplicity of the PPT system. ii The thruster operation and performance was investigated by several different diagnostic methods, including current and voltage measurements, Langmuir probes, and magnetic field probes. Furthermore, impact pressure measurements in the plume of the thruster allowed new insight into the plume structure and the accurate evaluation of impulse bits. Employment of the diagnostic methods for Teflon and water propellant enabled the unambigous identification of propellant related effects such as reduced electron temperature and higher exhaust velocites in the case of water propellant. The electromagnetic nature of the water thruster was clearly identified. For 30 J discharge energy, the water thruster requires only 5% of the mass bit of a Teflon thruster to produce an impulse bit 30% of the magnitude of the Teflon thruster, suggesting greatly increased propellant efficiencies. In agreement with the plasma diagnostic results, a specific impulse for the water thruster of up to 8000 s and efficiencies of up to 16% were evaluated.

Committee: Thomas York (Advisor) Subjects: Engineering, Aerospace

Master of Science in Engineering (MSEgr), Wright State University, 2013, Mechanical Engineering

The design and development of ion engines is a difficult and expensive process. In order to alleviate these costs and speed ion engine development, it is proposed to further develop a particle-in-cell (PIC), Monte-Carlo collision (MCC) model of an ion engine discharge chamber, which has previously been worked on by the Wright State Ion Engine Modeling Group. Performing detailed and accurate simulations of ion engines can lead to millions of dollars in savings in development costs. In order to recognize these savings more work must be done on the present day models used to simulate ion engine performance. The work presented in this thesis is an effort to do this with a computer model of the plasma in the discharge chamber of an ion engine. In particular, this thesis presents a few steps in the process of moving a Wright State developed PIC-MCC computer code, developed specifically for the plasma in the discharge chamber, to include detailed electric field calculations. This is a rather difficult process in that the electric fields present in the discharge chamber are strongly dependent on the location of the charged particles in the plasma. This means there is a strong and unstable connection between the particle position calculation and the electric field calculation. Other difficulties are the relatively large computational domain and the relatively large plasma density present. Because of the computational times involved,PIC-MCC techniques are generally not applied to large computational domains with high particle number densities, but this is the precise physical model that is required to obtain accurate results for the plasma in the discharge chamber of an ion engine. This thesis presents a few steps taken to get such a program to converge and to run in a stable fashion. Not only is getting the program to converge an issue, but getting convergence times that are less than one week is difficult. By no means is the work in this thesis a complete solutio (open full item for complete abstract)

Committee: James Menart Ph.D. (Advisor); Allen Jackson Ph.D. (Committee Member); Scott Thomas Ph.D. (Committee Member) Subjects: Aerospace Engineering; Electrical Engineering; Engineering; Plasma Physics