abstract: Two methods of improving the life and efficiency of the Pulsed Inductive Thruster (PIT) have been investigated. The first is a trade study of available switches to determine the best device to implement in the PIT design. The second is the design of a coil to improve coupling between the accelerator coil and the plasma. Experiments were done with both permanent and electromagnets to investigate the feasibility of implementing a modified Halbach array within the PIT to promote better plasma coupling and decrease the unused space within the thruster. This array proved to promote more complete coupling on the edges of the coil where it had been weak in previous studies. Numerical analysis was done to predict the performance of a PIT that utilized each suggested switch type. This model utilized the Alfven velocity to determine the critical mass and energy of these theoretical thrusters. / Dissertation/Thesis / Masters Thesis Aerospace Engineering 2018

Le xénon est utilisé par la plupart des systèmes de propulsion électrique à grilles. Cependant sa rareté, son coût de production important ainsi que son usage dans de nombreuses applications industrielles font apparaître la nécessité de trouver une alternative à ce propergol. Il est apparu que l'iode était un candidat potentiel pour cela, étant beaucoup moins cher à produire et beaucoup moins rare. Il se présente sous forme de cristaux violacés dans les conditions standards de pression et de température et possède une pression de vaporisation peu élevée ainsi qu'un potentiel d'ionisation plus bas que celui du xénon. Un modèle global d'un plasma d'iode dans un propulseur électrique à grilles a donc été développé afin d'étudier le comportement et les performances d'un tel dispositif. Ces résultats de l'iode sont comparés à ceux du même dispositif obtenus par l'utilisation du xénon, les conditions d'opération étant bien évidemment similaires. Le modèle prédit une efficacité globale du propulseur 15% plus grande pour l'iode. Les résultats du modèle global en iode sont également comparés avec des résultats expérimentaux obtenus dans un propulseur électrique à grilles, sous des conditions d'opération et paramètres d'entrée similaires. Un band d'essai expérimental entièrement dédié à l'étude de l'iode comme nouveau propergol pour la propulsion à grilles a en effet été assemblé avec toutes les précautions nécessaires, l'iode étant un élément corrosif et chimiquement actif avec certains matériaux. Le banc d'essai en iode fut également utilisé pour effectuer la preuve de concept en iode du propulseur PEGASES. / Most state-of-the-art electric space propulsion systems such as gridded and Hall thrusters use xenon as the propellant gas. However, xenon is rare, expensive to produce and used in a number of competing industrial applications. Alternatives to xenon are currently being investigated, and iodine has emerged as a potential candidate. Its lower cost, larger availability, its solid state at standard temperature and pressure, its low vapour pressure and its low ionization potential makes it an attractive option. A global model of iodine plasma inside an ion gridded thruster has therefore been developed to study behaviour and performances of this propellant. We compared the iodine results with ones obtained in xenon under otherwise similar conditions. The model predicts a thruster efficiency 15% higher for iodine compared to xenon. Results of the iodine global model were compared with experimental data obtained under similar operating conditions and input parameters in a gridded ion thruster. An experimental test bench dedicated to iodine plasma study, inside a classic ion gridded thruster and PEGASES thruster, has been assembled with all precautions needed. Iodine is a corrosive gas and chemically active with certain metals and the right choice of materials is therefore important. The positive ion and electron densities obtained by the model and in experiments appeared to show close values, indicating that the iodine chemistry and reaction set used in the global model seem relevant to a first order approximation.

Transportation is a primary pollution source contributing to 14 percent of global greenhouse gas emissions, and 12 percent of transportation emissions came from maritime activities. Emissions from the ferry industry, which carries roughly 2.1 billion passengers and 250 million vehicles annually, is a major concern for the general public due to their near-shore operations. Compared to the rapidly advancing clean automotive propulsion, fuel efficiency and emissions improvements for marine vessels are more urgent and beneficial due to the significantly higher petroleum fuel consumption and heavy pollutants and the relatively slow adoption of clean propulsion technology by the marine industry. Hybrid electric propulsion, proven to be effective for ground vehicles, presents a promising solution for more efficient clean marine transportation. Due to the diversified hull/propulsor design and operation cycle, the development of a hybrid electric marine propulsion system demands model-based design and control optimization for each unique and small batch production vessel. The integrated design and control optimization further require accurate and computation efficient hull resistance and propulsor thrust calculation methods that can be used to predict needed propulsion power and gauge vessel performance, energy efficiency, and emissions. This research focuses on improving the low-order empirical hull resistance and propulsor thrust models in the longitudinal direction by extracting model parameters from one-pass computational fluid dynamics (CFD) simulation and testing the acquired models in integrated design optimization of the marine propulsion system. The model is implemented in MATLAB/Simulink and ANSYS Aqwa and validated using operation data from BC Ferries’ ship Tachek. The modified low-order model (M-LOM) is then used in the integrated optimizations of propulsion system component sizes and operation control strategies for another BC Ferries’ ship, Skeena Queen. The performance, energy efficiency, and emissions of various propulsion options, including nature gas-mechanical and natural gas-electric benchmarks, and hybrid electric alternatives of series hybrid, parallel hybrid, and battery/pure electric are compared to demonstrate the benefits of the new method in completing these complex tasks and hybrid electric marine propulsion. The research forms the foundation for further studies to achieve more accurate propulsion demand prediction and a more comprehensive lifecycle cost assessment of clean marine propulsion solutions. / Graduate

High Efficiency Multistage Plasma Thrusters (HEMPTs) are a relatively new form of electric propulsion that show promise for use on a variety of missions and have several advantages over their older EP competitors. One such advantage is their long predicted lifetime and minimal wall erosion due to a unique periodic permanent magnet system. A laboratory HEMPT was built and donated by JPL for testing at Cal Poly. Previous work was done to characterize the performance of this thruster and it was found to exhibit a large plume divergence, resulting in decreased thrust and specific impulse. This thesis explores the design and application of a magnetic shield to modify the thruster’s magnetic field to force more ion current towards the centerline. A previous Cal Poly thesis explored the same concept, and that work is continued and furthered here. The previous thesis tested a shield which increased centerline current but decreased performance. A new shield design which should avoid this performance decrease is studied here. Magnetic modelling of the thruster was performed using COMSOL. This model was verified using guassmeters to measure the field strength at many discrete points within and near the HEMPT, with a focus on the ionization channel and exit plane. A shield design which should significantly reduce the radial field strength at the exit plane without affecting the ionization channel field was modelled and implemented. The HEMPT was tested in a vacuum chamber with and without the shield to characterize any change to performance characteristics. Data were collected using a nude Faraday probe and retarding potential analyzer. The data show a significant increase in centerline current with the application of the shield, but due to RPA malfunction and thruster failure the actual change in performance could not be concluded. The unshielded HEMPT was characterized, however, and was found to produce 12.1 +/- 1.3 mN of thrust with a specific impulse of 1361 +/- 147s. The thruster operated with a total efficiency of 10.63 +/- 3.66%, an efficiency much lower than expected. A large contributor to this low efficiency is likely the use of argon in place of xenon. Its lower mass and higher ionization energy make it a less efficient propellant choice. Further, the thruster is prone to overheating, indicating that significant thermal losses are present in this design.

This master thesis deals with the design of a handbike with electric propulsion. The final design is created in regard to knowledge from design and technical analysis and also to desficiencies of contemporary products. The purpose of this work is to create design of electrohandbike which will respect user and his needs from ergonomical and esthetic point of view.

Since 2000, a relatively new electric thruster concept has been in research, development, and production at Thales Electron Devices in Germany. This High Efficiency Multistage Plasma Thruster, or HEMPT, has promising lifetime capabilities due to its plasma confinement system. However, the permanent magnet system that offers this and other benefits also creates a hollow plume, where ions are accelerated at angles rather than up the thruster centerline, causing a dip in ion current along the centerline. A laboratory model, built at JPL, was run at Cal Poly to characterize this plume shape and implement a shield to restore a conical shape to the plume. A similar solution was used on a different type of thruster, a cylindrical hall thruster, at Princeton with excellent results. A shield was designed to shunt the magnetic field outside the thruster, where the Princeton experiments have identified a radial magnetic field as the cause for this hollow plume. The thruster was run with and without the shield, taking measurements of the ion current in the plume using a linear probe drive. The shield fixed the plume shape, increasing centerline current by 48%, however it also had detrimental effects on thruster performance, causing a decrease in thrust, specific impulse, and cut the total efficiency in half. The shield design was reexamined and a new design has been suggested for future testing of the HEMPT to restore performance while still fixing the plume shape.

The future of aviation technology is transitioning to cleaner, more efficient and higher endurance aircraft solutions. As fully electric propulsion systems still fall short of the operational requirements of modern day aircraft, there is increasing pressure and demand for the aviation industry to explore alternatives to fossil fuel driven propulsion systems. The primary focus of this research is to experimentally evaluate hybrid electric propulsion systems (HEPS) for Unmanned Aerial Vehicles (UAV) which combine multiple power sources to improve performance. HEPS offer several potential benefits over more conventional propulsion systems such as a smaller environmental impact, lower fuel consumption, higher endurance and novel configurations through distributed propulsion. Advanced operating modes are also possible with HEPS, increasing the vehicle’s versatility and redundancy in case of power source failure. The primary objective of the research is to combine all of the components of a small-scale HEPS together in a modular test bench for evaluation. The test bench uses components sized for a small-scale UAV including a 2.34kW two-stroke 35cc engine and a 1.65kW brushless DC motor together with an ESC capable of regenerative braking. Individual components were first tested to characterize performance, and then all components were assembled together in a parallel configuration to observe system-level performance. The parallel HEPS is capable of functioning in the four required operating modes: EM Only, ICE Only, Dash Mode (combined EM and ICE power) as well as Regenerative Mode where the onboard batteries get recharged. Further, the test bench was implemented with a supervisory controller to optimize system performance and run each component in the most efficient region to achieve torque requirements programmed into mission profiles. The logic based controller operates with the ideal operating line (IOL) concept and is implemented with a custom LabView GUI. The system is able to run on electric power or ICE power interchangeably without making any modifications to the transmission as the one-way bearing assembly engages for whichever power source is rotating at the highest speed. The most impressive of these sets of tests is the Dash mode testing where the output torque of the propeller is supplied from both the EM and ICE. Working in tandem, it was proved that the EM was drawing 19.9A of current which corresponds to an estimated 0.57Nm additional torque to the propeller for a degree of hybridization of 49.91%. Finally, the regenerative braking mode was proven to be operational, capable of recharging the battery systems at 13A. All of these operating modes attest to the flexibility and convenience of having a hybrid-electric propulsion system. The results collected from the test bench were validated against the models created in the aircraft simulation framework. This framework was created in MATLAB to simulate the performance of a small UAV and compare the performance by swapping in various propulsion systems. The purpose of the framework is to make direct comparisons of HEPS performance for parallel and series architectures against conventional electric and gasoline configuration UAVs, and explore the trade-offs. Each aircraft variable in the framework was modelled parametrically so that parameter sweeps could be run to observe the impact on the aircraft’s performance. Finally, rather than comparing propulsion systems in steady-state, complex mission profiles were created that simulate real life applications for UAVs. With these experiments, it was possible to observe which propulsion configurations were best suited for each mission type, and provide engineers with information about the trade-offs or advantages of integrating hybrid-electric propulsion into UAV design. In the Pipeline Inspection mission, the exact payload capacities of each aircraft configuration could be observed in the fuel burn versus CL,cruise parameter sweep exercise. It was observed that the parallel HEPS configuration has an average of 3.52kg lower payload capacity for the 35kg aircraft (17.6%), but has a fuel consumption reduction of up to 26.1% compared to the gasoline aircraft configuration. In the LIDAR Data collection mission, the electric configuration could be suitable for collection ranges below 100km but suffers low LIDAR collection times. However, at 100km LIDAR collection range, the series HEPS has an endurance of 16hr and the parallel configuration has an endurance of 19hr. In the Interceptor mission, at 32kg TOW, the parallel HEPS configuration has an endurance/TOW of 1.3[hr/kg] compared to 1.15[hr/kg] for the gasoline aircraft. This result yields a 13% increase in endurance from 36.8hr for gasoline to 41.6hr for the parallel HEPS. Finally, in the Communications Relay mission, the gasoline configuration is recommended for all TOW above 28kg as it has the highest loiter endurance. / Graduate

While vertical takeoff and landing aircraft show promise for urban air transport, distributed electric propulsion on existing aircraft may offer an immediately implementable alternative. Dis- tributed electric propulsion has the potential of increasing the aircraft thrust-to-weight ratio and lift coefficient high enough to enable takeoff distances of less than 100 meters. While fuel based propulsion technologies generally increase in specific power with increasing size, electric propul- sion typically can be decreased in size without a decrease in specific power. The smaller but highly power-dense propulsion units enable alternative designs including many small units, optionally powered units, and vectored thrust from the propulsion units, which can all contribute to better runway performance, decreased noise, adequate cruise speed, and adequate range. This concep- tual study explores a retrofit of continuously powered, invariant along the wingspan, open bladed electric propulsion units. To model and explore the design space we used a set of validated models including a blade element momentum method, a vortex lattice method, linear beam finite element analysis, classical laminate theory, composite failure, empirically-based blade noise modeling, mo- tor mass and motor controller empirical mass models, and nonlinear gradient-based optimization. We found that while satisfying aerodynamic, aerostructural, noise, and system constraints, a fully blown wing with 16 propellers could reduce the takeoff distance by over 50% when compared to the optimal 2 propeller case. This resulted in a conceptual minimum takeoff distance of 20.5 meters to clear a 50 ft (15.24 m) obstacle. We also found that when decreasing the allowable noise to 60 dBa, the fully blown 8 propeller case performed the best with a 43% reduction in takeoff distance compared to the optimal 2 propeller case. This resulted in a noise-restricted conceptual minimum takeoff distance of 95 meters.Takeoff distances of this length could open up thousands of potential urban runway locations to make a retrofit distributed electric aircraft an immediately implementable solution to the urban air transport challenge.

The feasibility of relocating a small (~500,000 kg) Near-Earth Asteroid (NEA) to High Earth Orbit via Solar Electric Propulsion (SEP) is evaluated with the orbital simulation software General Mission Analysis Tool (GMAT). Using prior research as a basis for the mission parameters, a retrieval mission to NEA 2008 HU4 is simulated in two parts: approach from Earth and return of the Asteroid Redirect Vehicle (ARV) with the asteroid in tow. Success of such a mission would pave the way for future missions to larger NEAs and other deep space endeavors. It is shown that for a hypothetical launch time of 24 May 2016, the ARV could arrive within 25 km of 2008 HU4 on 28 Jun 2017 with a Delta V of 0.406 km/s, begin return maneuver on 08 Dec 2017 and reach Earth altitude of 450,000 km by 23 Apr 2026 with a Delta V of 44.639 m/s.

Electric propulsion systems utilize electrical energy to produce thrust for spacecraft propulsion. These systems have multiple applications ranging from Earth orbit North-South station keeping to solar system exploratory missions such as NASA’s Discovery, New Frontiers, and Flagship class missions that focus on exploring scientifically interesting targets. In an electromagnetic thruster, a magnetic field interacting with current in an ionized gas (plasma) accelerates the propellant to produce thrust. Pulsed inductive thrusters rely on an electrodeless discharge where both the magnetic field in the plasma and the plasma current are induced by a time-varying current in an external circuit. The multi-dimensional acceleration model for a pulsed inductive plasma thruster consists of a set of circuit equations describing the electrical behavior of the thruster coupled to a one-dimensional momentum equation that allow for estimating thruster performance. Current models lack a method to account for the time-varying energy distribution in an inductive plasma accelerator.