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One idea to stop the increasing threat of global warming is to build a space sunshade, made up of a constellation of satellites to reflect a portion of the sunlight. These satellites need to reach an equilibrium point where they can stay in orbit with minimal adjustments. A few feasibility studies have been made in this area and the location of the equilibrium point, L1' has been found. An area that is still unclear is the best way of getting the satellites to L1'. One idea is to use electric propulsion engines , and utilizing a gravity assist around the Moon on the way to L1'. Gravity assists have mostly been performed with chemical rocket engines, so doing it with electric propulsion causes some unique issues. The goal was to find a trajectory including the gravity assist showing that this could be an effective way of transporting the satellites. To measure how effective this would be, a trajectory without a gravity assist was also created as a means of comparison. The trajectory and orbit simulations were done in the program General Missions Analysis Tool (GMAT), and the resulting trajectory took 51 days and used 83.4kg of fuel for a fuel-to-mass-ratio of 21.5%. This was a worse result compared to the trajectory without a gravity assist, which only used 80.4kg of fuel. Finally a discussion around a potential trajectory which was shown to have a much greater velocity increase was had, which would indicate that a gravity assist maneuver could provide a trajectory that does save on fuel compared to using no gravity assist.

Over the last two decades, Very Low Earth Orbit (VLEO) has gained researchers attention as it provides a significant amount of benefits in the field of earth observation and telecommunications. VLEO provides increased payload performance, improved geospatial accuracy, lower launch mass, simplified end of life disposal, and they reduce space-debris collision risk. However, the utilization of orbits with such low altitudes presents its own set of challenges, denser atmosphere will significantly increase aerodynamic drag, decaying the orbit in a short period of time. Besides increased drag VLEO environment will produce high levels of spacecraft (SC) charging and the presence of atomic oxygen will generate a constant erosion on the surfaces of the SC. An Atmosphere-Breathing Electric Propulsion (ABEP) ingests the residual atmosphere through an intake and uses it as propellant for an electric thruster. Theoretically applicable to any planet with an atmosphere, the system might allow drag compensation for an unlimited time without carrying propellant. In this thesis, different approaches for an intake are introduced, while the modeling, and numerical testing by Direct Simulation Monte Carlo (DSMC) is also presented. The intake is optimized for the RF Helicon-based Plasma Thruster (IPT) developed at IRS and a new concept design takes advantage of new materials properties, for specular surface interactions. Simulation results over different altitudes and conditions used for the verification of the design have been performed achieving a maximum collection efficiency of 94%. / Under de senaste två decennierna har “Very Low Earth Orbit“ VLEO fått stor uppmärksamhet inom forskningsvärlden då det leder till en mängd fördelar inom jordobservation och telekommunikation. VLEO ger ökad nyttolastprestanda, förbättrad geospatial noggrannhet, lägre startmassa, förenklat bortförskaffande och minskning kollisionsrisken för rymdskräp. Användningen av omloppsbanor på så låg höjd medför dessvärre också utmaningar. Den lägre höjden innebär tätare atmosfär och ökar därav det aerodynamiska luftmotståndet avsevärt. Förutom ökad luftmotstånd kommer miljön i VLEO att producera höga nivåer av rymdfarkostladdning och närvaron av atomärt syre leder till en konstant erosion av farkosters ytor. En ABEP leder in den återstående atmosfären genom ett intag och använder den som drivmedel för en elektrisk drivraket. Teoretiskt tillämpbart på alla planeter med en atmosfär, kan systemet tillåta luftmotståndskompensation under obegränsad tid utan något annat drivmedel än den redan befintliga atmosfären. I denna avhandling presenteras olika modeller för ett intag samtidigt som modellering och numerisk testning av DSMC också presenteras. Intaget optimeras för den IPT som utvecklats vid IRS och en ny konceptdesign utnyttjar nya materialegenskaper för speciella ytinteraktioner. Simuleringsresultat över olika höjder och förhållanden som används för verifiering av designen har utförts och resulterat i en maximal insamlingseffektivitet på 94%. / In den letzten zwei Jahrzenten gewann die tiefe Erdumlaufbahn “Very Low Earth Orbit“ (VLEO) durch die erheblichen Vorteile für Erdbeobachtung und Telekommunikation an Aufmerksamkeit in der Wissenschaft. VLEO ermöglicht eine höhere Nutzlastleistung, verbesserte räumliche Genauigkeit, eine geringere Startmasse, vereinfachte “End-of-Life“ - Entsorgung und verringert das Kollisionsrisiko von Weltraumschrott. Die Nutzung von Umlaufbahnen in diesen geringen Höhen stellt jedoch auch eine Reihe von Herausforderungen dar. Die dichtere Atmosphäre im VLEO erhöht den Luftwiderstand erheblich und verringert die Umlaufbahn in kurzer Zeit. Neben dem erhöhten Luftwiderstand tritt auch hohe Raumschiff- oder Satellitenladung auf und durch atomaren Sauerstoff entsteht konstante Erosion an den Oberflächen. Ein atmosphärenatmender elektrischer Antrieb (ABEP) nimmt die Restatmosphäre über einen Einlass auf und verwendet sie als Treibstoff für ein elektrisches Triebwerk. Theoretisch auf jeden Planeten mit Atmosphäre anwendbar, könnte das System so den Widerstand zeitlich unbefristet ohne Treibstoffverwendung kompensieren. In dieser Arbeit werden verschiedene Ansätze für einen Einlass vorgestellt, und die Modellierung und numerischen Tests durch die “Direct Simulation Monte Carlo“ (DSMC) werden präsentiert. Der Einlass ist für den am IRS entwickelten RF Helicon-basierten Plasma Thruster (IPT) optimiert. Ein neues Konzeptdesign nutzt neue Materialeigenschaften für spiegelartige Oberflächen-Reflektionseigenschaften. Simulationsergebnisse verschiedener Höhen und Konditionen wurden zu der Überprüfung des Entwurfs verwendet, wobei eine maximale Einlassammlungswirkungsgrad von 94% erreicht wurde.

Over the past few years, electric propulsion systems have been widely used in automotive applications. The next decade is likely to see the electrification of aerial vehicles. In the past 20 years, the passengers demand in the aviation industry has increased by roughly 5% annually. Drastic increment in the passengers demand leads to many problems such as emission, noise pollution, airports capacity shortage, and high fuel consumption. An electric airplane that can take off and land in an extremely short runway can solve all the mentioned problems. Also, an airplane that is smaller and lighter with the ability to take off and land from an extremely short runway can be used as a new transportation system in congested cities and solve the urban road traffic and compensate for people’s time wasted in traffic. With this in mind, in this thesis, the feasibility of converting a conventional fixed-wing direct-drive propeller airplane to an electric extremely short takeoff and landing airplane has been examined. An overview of the history of electric aerial vehicles and flying cars is conducted where some of these vehicles are still under development phase. The main aim of this thesis is to address the effect of takeoff and landing runway length on the electric motor main specifications, including power, torque, and speed. Also, the effect of cruising speed on the motor specifications are investigated, and it is observed that there is a considerable difference between the amount of required power for the cruising mode and takeoff mode. In the end, the impact of the braking system and airplane weight on the landing distance are examined, and It is found that for an airplane with a cruise-efficient propeller, usage of thrust reverser is not practical and hence it is not recommended. Although if the propeller is designed to have high efficiency at takeoff and landing, the thrust reverser can be a good solution to make the landing runway shorter. / Thesis / Master of Applied Science (MASc)

As an alternative solution to global warming, this thesis explores the possibility of aspace-based geoengineering scheme that may prove worthwhile to implement in parallel toother environmental efforts that help mitigate impact of climate change. One suggestionof a geoengineering solution is deploying a large number of sunshades in the vicinity ofthe first Lagrange point of the Sun-Earth system, and this prospective sunshade projectwould serve to shield Earth from incident solar radiation. This thesis is an extension ofa feasibility study for the implementation of this large-scale mission, and has a focus oncomparing electric thrusters to solar sailing as a means of propulsion. Background onelectric propulsion systems and spaceflight mechanics is provided. The investigation wasperformed by defining the spacecraft configurations, and then computing trajectories toa point of escape from Earth and from there to the final equilibrium point.Our results show that in order to meet the propellant demands of the electric thrusters,the launch mass would need to increase by around 15-25 % compared to the solar sailingimplementation, equating to around 1010 kg. Nevertheless, electric propulsion could stillbe a beneficial choice since it would allow shorter transfer times for each shade whichreduces the radiation exposure and subsequent degradation of the spacecraft’s systems.It was found that the transfer time with electric propulsion would be about one-half orone-fifth that of solar sailing, depending on spacecraft parameters. Additionally, electricpropulsion allows a much lower initial parking orbit, and while this would increase the ra-diation exposure it would also reduce the launch costs due to the higher payload capacityto lower altitudes. However, electric propulsion of this scale require prior advancementsin xenon or other inert propellant extraction methods and possibly a wide-scale construc-tion of air separation plants.

Combustion emissions from aviation operations contribute significantly to climate change and air pollution. Accordingly, there is increasing interest in advancing battery-powered propulsion for aviation applications to reduce emissions. As batteries continue to improve, it is essential to recognize breakthroughs in battery specific energy in the context of air transport vehicles. Most electric aircraft designs and programs have focused on small aircraft because of restrictive battery performance. This work presents a feasibility assessment for an all-electric airliner based on an Airbus A220-100 with turbofan engines replaced by electric motors and propellers. The analysis compares the performance characteristics of the electric airliner to the A220-100 and establishes several configurations with varying battery pack-specific energy. The short-term electric airliner could replace conventional aircraft on very short, high-density missions. In contrast, the long-term electric airliner requires significant battery technology improvements that are not currently foreseen. The alternative long-term electric airliner could complete half of the A220-100’s missions, but the necessary specific energy value is also not anticipated shortly. All-electric airliners would significantly impact manufacturing, operations, costs, and emissions but are commercially infeasible with current battery technology. Additional development of more advanced battery technology is required to increase the specific energy of battery packs, enhance battery safety and reliability, and develop lighter high-power electric motors.

Investigation of complex multirotor aerodynamic phenomena via wind tunnel experimentation is becoming extremely important with the rapid progress in advanced distributed propulsion VTOL concepts. Much of this experimentation is being performed in large, highly advanced tunnels. However, the proliferation of this class of vehicles extends to small aircraft used by small businesses, universities, and hobbyists without ready access to this level of test facility. Therefore, there is a need to investigate whether multirotor vehicles can be adequately tested in smaller wind tunnel facilities. A test rig for a 2.82-pound quadcopter was developed to perform powered testing in the Cal Poly Aerospace Department’s Low Speed Wind Tunnel, equipped with a 3-foot tall by 4-foot wide test section. The results were compared to data from similar tests performed in the U.S. Army 7-by 10-ft Wind Tunnel at NASA Ames. The two data sets did not show close agreement in absolute terms but demonstrated similar trends. Due to measurement uncertainties, the contribution of wind tunnel interference effects to this discrepancy in measurements was not able to be properly quantified, but is likely a major contributor. Flow visualization results demonstrated that tunnel interference effects can likely be minimized by testing at high tunnel speeds with the vehicle pitched 10-degrees or more downward. Suggestions towards avoiding the pitfalls inherent to multirotor wind tunnel testing are provided. Additionally, a modified form of the conventional lift-to-drag ratio is presented as a metric of electric multirotor aerodynamic efficiency.

In order to support UCLA’s development of the 3 cm Miniature Xenon Ion (MiXI) thruster, Cal Poly has a 3 cm thruster under development. This version, called MiXI Cal Poly Version 1 (MiXI-CPv1), is complete and has been utilized in vacuum chamber thermal validation testing. Testing on this version was used to check the validity of heat transfer simulations modeled in SolidWorks. Investigations of the 3 cm ion thruster configuration were intended to discover the driving factors affecting the thermal behavior of the discharge chamber and surrounding design space. Numerical simulations indicate that the heating of the samarium cobalt permanent magnets can be mitigated through the implementation of two proposed modifications. The first modification is to implement a 2% thoriated tungsten filament cathode. This design exhibited maximum permanent magnet temperatures of 325°C, twenty-five degrees below the maximum upper temperature of 350°C. Since some magnetic degaussing effects have been observed at temperatures above 300°C, the aforementioned solution can be combined with a thruster design modification to achieve a reduced permanent magnet temperature of 298°C. This modification would involve increase the anode wall thickness from approximately 0.7 mm to 2 mm below the permanent magnet ring, creating a stepped anode design. Additionally, less effective solutions were proposed and modeled and are presented for completeness.

The power requirements for future spacecraft power systems will be on the order of a few hundred kilowatts to a few megawatts. Because of these power levels, a high-voltage, high-power distribution subsystem may be utilized to transmit power from the source to the different loads. Using current state-of-the-art power conditioning electronics, complex series and parallel configurations will be required at the interface between the source and the distribution subsystem and between the distribution subsystem and the loads. The dynamic response of such a spacecraft power system may be obtained using a general purpose program such as SPICE2. However, for large and complex spacecraft power systems, the input file will be large and complex with correspondingly large computation times. As an alternative, the spacecraft power system can be considered as an interconnection of modular components. Each component is treated as a two-port network, and a state model is written with the port voltages as the inputs. The state model of each component is solved using the state transition matrix and assuming that the port voltages are . clamped for each time step. This calculation proceeds as if all two-port networks are decoupled. After the state variables have been updated, the inputs to all components are calculated using network analysis principles. The solution procedure alternates between solving the dynamic model of all components and the network equations for the component inputs. The modular state variable approach and SPICE2 are compared using two example systems. This comparison shows the advantages of the modular state variable approach. First, for the modular state variable approach the system is considered as an interconnection of modular components. In SPICE2, the system is treated as an interconnection of circuit elements. As a result, the system description for large and complex spacecraft power systems is much _ larger and more complex than a modular state variable description. Secondly, the modular state variable approach requires less CPU time than SPICE2. For one of the example systems presented here, the modular state variable approach uses one-twentieth of the CPU time used by SPICE2. / Ph. D.

As the world moves towards environmental sustainability, the civil aviation enterprise has responded by setting challenging goals for significantly increased energy efficiency and reduced harmful emissions into the atmosphere as codified by National Aeronautics and Space Administration (NASA) and Advisory Council for Aircraft Innovation and Research in Europe (ACARE). The airline industry supports these goals because of their positive impact on operational cost and the environment. Achieving such goals requires introduction of novel technologies and aircraft concepts. Previous studies have shown that electrified aircraft can be effective in meeting these challenges.While there are several mechanisms to incorporate novel technologies for electrified aircraft, two such technologies: turbo-electric propulsion and distributed propulsion, are used in this research. Integration of these two technologies with the airframe leverages the well-known favorable interference between the wing and the tractor propeller wake to provide increased lift during takeoff.In the present research, the advantages and disadvantages of integrating a turbo-electric distributed propulsion (TEDiP) system are assessed for a regional transport aircraft (RTA). With near term motor technology, an improvement in trip fuel burn was observed on the four and six propeller variants of the TEDiP aircraft. The takeoff field length(TOFL) also improved in all three design variants which is a direct result of the working of distributed propulsion leading to better aerodynamic performance at takeoff conditions.The approach and findings for this research are reported in this thesis. / Master of Science / While air transportation system is considered the fastest means to travel, the avi-ation industry is responsible for 2.1% of all human-induced CO2 emissions, whichputs a renewed emphasis on environmental sustainability. There is heightenedinterest in exploring alternative propulsion technologies for aviation to mitigatethe effects of ever increasing demand for air travel coupled with fossil fuel pricevolatility.Ambitious plans have been outlined by leading aerospace organizations to reduceharmful emissions into the atmosphere. Achieving these ambitious goals requiresdevelopment and introduction of game changing technologies and aircraft con-cepts. Few such concepts include novel propulsion systems like all electric andhybrid-electric propulsion, distributed propulsion, and boundary layer ingestion.The X-57 is a novel all-electric aircraft being developed by NASA as a technologydemonstrator and makes use of multiple electric motors and propellers placedon the wing.Owing to battery technology limitations, all-electric and hybrid-electric propul-sion are not considered as viable options. In the near term, incorporatingdistributed propulsion alongside turbo-electric propulsion, for a Turbo-ElectricDistributed Propulsion(TEDiP) system may be a promising option in the near--to-mid-term. The overall goal of the present study is to investigate potentialbenefits and penalties of TEDiP systems for regional transport aircraft (RTA).To perform this study, the aerodynamics module of Pacelab Aircraft PreliminaryDesign (APD) Multi-Disciplinary Optimization (MDAO) framework is alteredto account for changes in wing-propeller aerodynamics due to the interactionof wing and multiple propellers. This required selection of a cost-effective toolthat captures aerodynamic data for multiple propellers and wing. VSPAEROis the aerodynamic tool of choice for this research. Aerodynamic data fromVSPAERO is coupled to APD and three TEDiP design variants with four, sixand eight propeller are designed with the ATR 72-500 as the baseline. Thebenefits and penalties of integrating the TEDiP system onto these variants isinvestigatedThe results show that a performance comparable to the baseline can be achievedin the near term with the four propeller variant even with current electricalsystems technology trends with a small weight penalty, and in the medium termon a six propeller variant. A decrease in trip fuel burn and improved takeofffield length(TOFL) performance justifies the usage of TEDiP systems.