The presented research discusses the design, analysis, and testing of a low current, LaB6 heaterless hollow cathode for space propulsion applications. A heaterless design using LaB6 is chosen to reduce complexity and increase electrical power efficiency and robustness. Argon propellant is used due to its more favorable breakdown voltage characteristics compared to xenon. An original model for the insert region plasma is derived by combining several analyses in literature. This model allows the simultaneous calculation of many plasma and thermal parameters in the cathode using only two completely unobtrusive measurements, and requires several assumptions which are common in hollow cathode research. The design of the cathode and its subsystems are presented in detail. No diagnostics were used in the cathode except direct voltage measurements in the power circuit. A discussion of emitter poisoning and ignition behavior is presented. The cathode is characterized by measuring anode and keeper voltages as a function of anode current and propellant flow rate, with the cathode discharging directly to a flat metal anode. Results are consistent with those obtained by previous investigations of argon hollow cathodes. This data is used with the derived plasma model to calculate the dependence of various parameters on current and flow rate. A discussion of the spot-plume transition behavior is presented. Finally, insights and design improvements are discussed based on the experimental results. / Master of Science / In recent years, the space industry has seen rapidly accelerating growth due to the continuing advancement of technology. A critical area of spacecraft technology is the spacecraft’s propulsion system, which allows the vehicle to achieve and maintain its desired orbit or trajectory through space. One class of propulsion systems known as “electric propulsion” uses electrical power to accelerate the fuel of the spacecraft. These types of propulsion systems are far more efficient than traditional propulsion systems, which use chemical reactions to create thrust. One of the main components of certain types of electric propulsion systems is the hollow cathode, which initiates and sustains the thruster operation. In this research, a hollow cathode with several non-conventional characteristics is developed and tested. First of all, standard hollow cathodes use a heater to bring the cathode up to operational temperature, but this design is heaterless which offers several benefits to the cathode and electrical power system designs. Secondly, the cathode uses a non-conventional choice of material for the “emitter”, which emits electrons when heated and allows the cathode to operate. Lastly, while typical electric propulsion systems use xenon for fuel, this cathode uses argon which has several benefits over xenon including cost. An overview of electric propulsion is presented, as well as a new physics-based model of this type of cathode that allows useful calculations based on simple measurements. The design and test results of the cathode are discussed in detail, with several interesting and insightful behaviors that were noted during testing. Heaterless cathodes have the potential to improve the efficiency, cost, and weight of electric propulsion systems, and this research therefore contributes to an important field for the future of space exploration.

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Colloid thrusters, and electrospray thrusters as a whole, have been around since the 1960s. When they were first developed, the high efficiency and fine thrust control was overshadowed by the high power requirement for such a low thrust that the system provides. This caused the technology to be put on hold for aerospace applications. Now, as small satellites are becoming more prevalent, there has been a resurgence in interest in electrospray thruster technology. The recent advancements in tech- nology allow electrospray thrusters to use significantly less power and occupy less volume than their predecessors. As electrospray technology continues to advance, these thrusters are meeting the demands of small satellite propulsion. As such, in an effort to keep the spacecraft propulsion curriculum current with today’s technology, a colloid thruster is designed, built, tested, and implemented as a laboratory activity at California Polytechnic State University, San Luis Obispo. Electrospray thrusters work by placing a voltage on an ionic liquid and extracting either beads of propellant or ions to generate thrust. By definition, colloid thrusters are a specific class of electrospray thrusters that use solvents, such as glycerol or formamide, to emit droplets or, in special cases, ions to generate thrust. To keep with the University’s “Learn by Doing” pedagogical philosophy, the thruster for this activity is designed to have a tactile and experiential impact on the students. The final design is a scaled up configuration of an existing electrospray design so that the students can easily see each component with the naked eye and can be correlated to a real world thruster that they might see in industry. As a laboratory experiment, the thruster needs to be able to utilize current equip- ment in the Space Environments and Testing Laboratory. One of the Student Vacuum Chambers (SVC) is utilized as well as two 1 kV power supplies and a 100V power supply. An indirect method of measuring performance metrics needs to be developed as there are no thrust balances sensitive enough in the lab designated for undergrad- uate use. As such, the students will be using the mass of the propellant, the time of operation, and knowledge of the propellant’s properties to estimate the performance of the thruster. To prove success of the thruster, a performance profile of the thruster is produced using an indirect method of measurement as well as visual observations of the thruster moving propellant byway of the electrospray theory. The tests show thrusts produced between 96-311 μN with an Isp ranging from 1270-1684 seconds. The visual evidence demonstrates propellant being collected as well as the operation of the thruster under the electrospray theory. The visual evidence also sheds light on which emission mode the thruster is operating at as well as a self-correcting failure mode that was occurring. The thruster is implemented as a lab for Cal Poly’s AERO 402 Spacecraft Propulsion Lab in Fall 2018, and it receives positive feedback from the students through an anonymous survey. While the colloid thruster demonstrates success in meeting performance and pedagog- ical goals, future work should be continued to improve the thruster. Further design and manufacturing work can be undertaken to improve the efficiency and decrease failure due to propellant impingement. Additionally, the procurement of power sup- plies capable of applying higher voltages can provide a greater range of operation which can enable a more dynamic student discovery of electrospray thrusters.

As conventional chemical rockets reach the outer limits of their abilities, significant research is going into alternative thruster technologies, some of which decouple the maximum thrust and efficiency from the propellant's internal chemical energy by supplying energy to the propellant as needed. Of particular interest and potential is the electrically powered thruster, which promises very high specific thrust using relatively inexpensive and stable propellant gasses. Some such thrusters, specifically ion thrusters, have achieved significant popularity for various applications. However, there exist other classes of electrical thrusters which promise even higher levels of efficiency and performance. This thesis will focus on one such thruster type - the magnetoplasmadynamic thruster - which uses an ionized propellant flow and large currents to accelerate the propellant gas by electrical and magnetic force interactions. The necessary background will be presented in order to understand and characterize the operation of such devices, and a theoretical model will be developed in order to estimate the levels of performance which can be expected. Simulations will be performed and analyzed in order to better understand the principles on which these devices are designed. Finally, a thruster package will be designed and built in order to test the performance of the device and accuracy of the model. This will include a high-current power supply, ignition circuit, gas delivery system, and nozzle. Finally, the measured performance of this thruster package will be measured and compared to the theoretical predictions in order to validate the models constructed for this type of thruster.

The rectangular pulsed plasma thruster (PPT) is an electromagnetic thruster that ablates Teflon propellant to produce thrust in a discharge that lasts 5-20 microseconds. In order to integrate PPTs onto spacecraft, it is necessary to investigate possible thruster plume-spacecraft interactions. The PPT plume consists of neutral and charged particles from the ablation of the Teflon fuel bar as well as electrode materials. In this thesis a novel application of quadruple Langmuir probes is implemented in the PPT plume to obtain electron temperature, electron density, and ion speed ratio measurements (ion speed divided by most probable thermal speed). The pulsed plasma thruster used is a NASA Glenn laboratory model based on the LES 8/9 series of PPTs, and is similar in design to the Earth Observing-1 satellite PPT. At the 20 J discharge energy level, the thruster ablates 26.6 mg of Teflon, creating an impulse bit of 256 mN-s with a specific impulse of 986 s. The quadruple probes were operated in the so-called current mode, eliminating the need to make voltage measurements. The current collection to the parallel to the flow electrodes is based on Laframboise's theory for probe to Debye length ratios between 5 and 100, and on the thin-sheath theory for ratios above 100. The ion current to the perpendicular probe is based on a model by Kanal and is a function of the ion speed ratio, the applied non-dimensional potential and the collection area. A formal error analysis is performed using the complete set of nonlinear current collection equations. The quadruple Langmuir probes were mounted on a computer controlled motion system that allowed movement in the radial direction, and the thruster was mounted on a motion system that allowed angular variation. Measurements were taken at 10, 15 and 20 cm form the Teflon fuel bar face, at angles up to 40 degrees off of the centerline axis at discharge energy levels of 5, 20, and 40 J. All data points are based on an average of four PPT pulses. Data analysis shows the temporal and spatial variation in the plume. Electron temperatures show two peaks during the length of the pulse, a trend most evident during the 20 J and 40 J discharge energies at 10 cm from the surface of the Teflon fuel bar. The electron temperatures after the initial high temperature peak are below 2 eV. Electron densities are highest near the thruster exit plane. At 10 cm from the Teflon surface, maximum electron densities are 1.04e20 ± 2.8e19 m-3, 9.8e20 ± 2.3e20 m-3, and 1.38e21 ± 4.05e20 m-3 for the 5 J, 20 J and 40 J discharge energy, respectively. The electrons densities decrease to 2.8x1019 ± 8.9e18 m-3, 1.2e20 ± 4.2e19 m-3, and 4.5e20 ± 1.2e20 m-3 at 20 cm for the 5 J, 20 J, and 40 J cases, respectively. Electron temperature and density decrease with increasing angle away from the centerline, and with increasing downstream distance. The plume is more symmetric in the parallel plane than in the perpendicular plane. Ion speed ratios are lowest near the thruster exit, increase with increasing downstream distance, but do not show any consistent angular variation. Peak speed ratios at a radial distance of 10 cm are 5.9±3.6, 5.3±0.39, and 4.8±0.41 for the 5 J, 20 J and 40 J discharge energies, respectively. The ratios increase to 6.05±5.9, 7.5±1.6, and 6.09±0.72 at a radial distance of 20 cm. Estimates of ion velocities show peak values between 36 km/s to 40 km/s, 26 km/s to 30 km/s, and 26 km/s to 36 km/s for the % J, 20 J, and 40 J discharge energies, respectively.

Beyond Atlas is a Swedish private company with the goal of exploring the solar system with cheap and reliable spacecraft. Part of their maiden mission, Hugin, aims to demonstrate navigation, propulsion, and communication technology on a 3U CubeSat. This thesis aims to investigate the feasibility of using the Enpulsion NANO electric propulsion (EP) system for deep-space applications and how to best demonstrate its capabilities in low-Earth orbit. Literature reviews of scientific papers and software simulations were conducted to gain an understanding of the underlying processes involved in EP in-orbit operations. Analyses were made on orbital maneuvers, momentum unloading, power and thermal restrictions. The results suggest that the EP system's capabilities is mainly limited by the saturation time of the reaction wheels restricting longer duration orbital maneuvers. Orbital maneuvers for demonstrating the capabilities are proposed based on the limitations imposed on the EP system by the rest of the spacecraft. On the basis of the results of this research, it can be concluded that the Enpulsion NANO thruster's operational range can be utilized both as a low thrust efficient main drive and as a high thrust maneuvering thruster for deep-space applications but is limited by the high power consumption and low thrust-to-power ratio.

"The ablative Teflon pulsed plasma thruster (PPT) is an onboard electromagnetic propulsion enabling technology for small spacecraft missions. The integration of PPTs onboard spacecraft requires the understanding and evaluation of possible thruster/spacecraft interactions. To aid in this effort the work presented in this thesis is directed towards the development and application of Langmuir probe techniques for use in the plume of PPTs. Double and triple Langmuir probes were developed and used to measure electron temperature and density of the PPT plume. The PPT used in this thesis was a laboratory model parallel plate ablative Teflon® PPT similar in size to the Earth Observing (EO-1) PPT operating in discharge energies between 5 and 40 Joules. The triple Langmuir probe was operated in the current-mode technique that requires biasing all three electrodes and measuring the resulting probe currents. This new implementation differs from the traditional voltage-mode technique that keeps one probe floating and requires a voltage measurement that is often susceptible to noise in the fluctuating PPT plume environment. The triple Langmuir probe theory developed in this work incorporates Laframboise’s current collection model for Debye length to probe radius ratios less than 100 in order to account for sheath expansion effects on ion collection, and incorporates the thin-sheath current collection model for Debye length to probe radius ratios greater than 100. Error analysis of the non-linear system of current collection equations that describe the operation of the current-mode triple Langmuir probe is performed as well. Measurements were taken at three radial locations, 5, 10, and 15 cm from the Teflon® surface of the PPT and at angles of 20 and 40 degrees to either side of the thruster centerline as well as at the centerline. These measurements were taken on two orthogonal planes, parallel and perpendicular to the PPT electrodes. A data-processing software was developed and implements the current-mode triple Langmuir probe theory and associated error analysis. Results show the time evolution of the electron temperature and density. Characteristic to all the data is the presence of hot electrons of approximately 5 to 10 eV at the beginning of the pulse, occurring near the peak of the discharge current. The electron temperature quickly drops off from its peak values to 1-2 eV for the remainder of the pulse. Peak electron densities occur after the peak temperatures. The maximum electron density values on the centerline of the plume of a laboratory PPT 10 cm from the Teflon® surface are 6.6x10^19 +/- 1.3x10^19 m^-3 for the 5 J PPT, 7.2x10^20 +/- 1.4x10^20 m^-3 for the 20 J PPT, and 1.2x10^21 +/- 2.7x10^20 m^-3 for the 40 J PPT. Results from the double Langmuir probe taken at r=10 cm, theta perpendicular=70 degrees and 90 degrees of a laboratory PPT showed good agreement with the triple probe method."

Hall Effect Thrusters (HETs) are a form of electric propulsion device which uses external electrical energy to produce thrust. When compared to various other electric propulsion devices, HETs are excellent candidates for future orbit transfer and interplanetary missions due to their relatively simple configuration, moderate thrust capability, higher thrust to power ratio, and lower thruster mass to power ratio. Due to the short history of HETs, the current design process of a new HET is a largely empirical and experimental science, and this has resulted in previous designs being developed in a narrow design space based on experimental data without systematic investigations of parameter correlations. In addition, current preliminary low-thrust trajectory optimizations, due to inherent difficulties in solution procedure, often assume constant or linear performances with available power in their applications of electric thrusters. The main obstacles come from the complex physics involved in HET technology and relatively small amounts of experimental data. Although physical theories and numerical simulations can provide a valuable tool for design space exploration at the inception of a new HET design and preliminary low-thrust trajectory optimization, the complex physics makes theoretical and numerical solutions difficult to obtain. Numerical implementations have been quite extensively conducted in the last two decades. An investigation of current methodologies reveals that to date, none provide a proper methodology for a new HET design at the conceptual design stage and the coupled low-thrust trajectory optimization. Thus, in the first half of this work, an efficient, robust, and self-consistent numerical method for the analysis of HETs is developed with a new approach. The key idea is to divide the analysis region into two regions in terms of electron dynamics based on physical intuition. Intensive validations are conducted for existing HETs from 1 kW to 50 kW classes. The second half of this work aims to construct a simultaneous design optimization environment though collaboration with experts in low-thrust trajectory optimization where a new HET and associated optimal low-thrust trajectory can be designed simultaneously. A demonstration for an orbit raising mission shows that the constructed simultaneous design optimization environment can be used effectively and synergistically for space missions involving HETs. It is expected that the present work will aid and ease the current expensive experimental HET design process and reduce preliminary space mission design cycles involving HETs.

This thesis aims to uncover preliminary design relationships for an inlet of a split-wing electric distributed propulsion regional airliner. Several aspects of the inlet design were investigated, including: the overall thickness of the airfoil section with respect to the chord, inlet throat area, and lip radius. These parameters were investigated using several angles of attack and mass flow rates through the fan. Computational fluid dynamics, with a 2nd Order turbulence model was used and validated against World War II era data from NACA, as those studies were the most pertinent wind tunnel data available. Additionally, other works by Boeing, Empirical Systems Aerospace (ESAero), Rolling Hills Research, and the Air Force Research Laboratories (AFRL) were considered as part of this design tool tradespace. Future work considerations include utilizing an airfoil section designed for M = 0.6 or 0.65 cruise conditions as opposed to a symmetrical airfoil section, extruding the 2-D airfoil section discussed in this thesis for 3-D effects, and incorporating fan rotational physics into the simulations to better account for inlet Mach number effects.

In order to enable comparison of Helicon ion thruster performance across different vacuum test facilities, an understanding of the effect of operating pressure on plasma plume properties is required. Plasma property measurements are compared for thruster operation at two separate vacuum facility operating pressures to determine the effect of neutral ingestion on Helicon ion thruster operation. The ion energy distribution function (IEDF), electron temperature, ion number density, and plasma potential are measured along the thruster main axis for a replica of the Madison Helicon eXperiment. Plasma property values recorded at the ‘high-pressure condition’ (3.0×10^(-4) Torr corrected for argon) are compared to values recorded at the ‘low-pressure condition’ (1.2×10^(-5) Torr corrected for argon) for thruster operation at 100 - 500 watts radio frequency forward power, 340 – 700 gauss source region magnetic field strength, and 1.3 - 60 sccm argon volumetric flow rate (0.039-1.782 mg/s). Differences in plasma behavior at the ‘high-pressure condition’ result from two primary neutral-plume interactions: collisions between accelerated beam ions and ingested neutrals leading to a reduction of ion energy and neutral ionization downstream of the thruster exit due to electron-neutral collisions. Electron temperature at higher operating pressures is lowered due to an electron cooling effect resulting from repeated collisions with neutral atoms. Results suggest that Helicon ion thruster plasma properties are greatly influenced when subjected to neutral ingestion.