Atomization of fuel is essential in controlling combustion inside a direct injection engine. Controlling combustion helps in reducing emissions and boosting efficiency. Cavitation is one of the factors that significantly affect the nature of spray in a combustion chamber. Typical fuel injector nozzles are small and operate at a very high pressure, which limit the study of internal nozzle behavior. The time and length scales further limit the experimental study of a fuel injector nozzle. Simulating cavitation in a fuel injector will help in understanding the phenomenon and will assist in further development. The construction of any simulation of cavitating injector nozzles begins with the fundamental assumptions of which phenomena will be included and which will be neglected. To date, there has been no consensus about whether it is acceptable to assume that small, high-speed cavitating nozzles are in thermal or inertial equilibrium. This diversity of opinions leads to a variety of modeling approaches. If one assumes that the nozzle is in thermal equilibrium, then there is presumably no significant delay in bubble growth or collapse due to heat transfer. Heat transfer is infinitely fast and inertial effects limit phase change. The assumption of inertial equilibrium means that the two phases have negligible slip velocity. Alternatively, on the sub-grid scale level, one may also consider the possibility of small bubbles whose size responds to changes in pressure. Schmidt et al. developed a two dimensional transient homogeneous equilibrium model which was intended for simulating a small, high speed nozzle flows. The HEM uses the assumption of thermal equilibrium to simulate cavitation. It assumes the two-phase flow inside a nozzle in homogeneous mixture of vapor and liquid. This work presents the simulation of high-speed nozzle, using the HEM for cavitation, in a multidimensional and parallel framework. The model is extended to simulate the non-linear effects of the pure phase in the flow and the numerical approach is modified to achieve stable result in multidimensional framework. Two-dimensional validations have been presented with simulation of a venturi nozzle, a sharp nozzle and a throttle from Winklhofer et al. Three-dimensional validations have been presented with simulation of ‘spray A’ and ‘spray H’ injectors from the Engine Combustion Network. The simulated results show that equilibrium assumptions are sufficient to predict the mass flow rate and cavitation incidence in small, high-speed nozzle flows.

Scramjet engines are key for sustained hypersonic flights. Analytic models play a critical role in the preliminary design of a scramjet engine configuration. The objective of this research is to develop and validate a quasi-1D model for the scramjet engine encompassing inlet, isolator and combustor, to evaluate the impact of flight conditions and design parameters on the engine functionality. The model is developed assuming isentropic flow in the inlet with a single turn; modified Fanno-flow equations in the isolator that account for the area change of the core flow; and the combustor is modeled using Heiser-Pratt equations accounting for the fuel mixing efficiency. The isolator and combustor models are validated against experimental results. The model accounts for twelve parameters allowing for a decent range of possible configurations. Finally, the model was applied to five sets of parametric studies to evaluate the effect of multiple parameters on the engine functionality.

<p>  </p> <p>Monolithic catalyst beds have been used extensively in other industries and are gaining interest for space propulsion applications. Additive manufacturing of monolithic supports allows for catalyst beds with a wider range of geometries than could be produced using conventional methods, potentially allowing for higher performance monoliths that can compete with conventional packed beds in performance. Achieving these gains requires a consistent, even, and well-adhering washcoating procedure for the additively manufactured supports, one which works well on varied geometry and on support materials that can be readily printed. I conducted an extensive development process on improving methods of surface preparation and coating for high temperature ceramic monoliths that resulted in improvements in the state of the art. The materials and methods used are appropriate for rocket grade hydrogen peroxide, hydrazine, or other monopropellants with similar operating temperatures. Using existing published coating methods resulted in uneven coating distribution and poor adhesion. I demonstrate that this was due to the substrate surface morphology producing a hydrophobic effect. Surface morphology plays a significant role in coating coverage and adhesion and differences in initial support surfaces likely account for much of the variation in results seen across the literature. I present a method of controlled thermochemical surface etching using pure sodium hydroxide at 420°C that can reliably produce a roughened hydrophilic surface from a variety of starting morphologies. I also present several modifications to the primer formulation that improve evenness of coverage, the most significant of which is the inclusion of a surfactant at a concentration of 1 g per 36 g water. The surface treatment and coating formulation improvements combine well and produce an even coating with strong adhesion to the substrate. I also conducted preliminary work on the investigation of novel geometric designs for monolithic catalyst beds, and on the reactivity of different transition metal oxide catalysts for rocket grade hydrogen peroxide decomposition. </p>

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Energy efficiency and propulsive characteristics of a 10 cm undulatory rainbow trout (oncorhynchus mykiss) swimming in a stationary position are considered. Two CFD simulations are performed utilizing dynamic grid meshing (FLUENT 6.3). The first simulation uses a laminar flow model with an added hydrofoil shape in order to test if thrust and drag can be brought to unity. The second simulation uses a Large-Eddy Simulation (LES) turbulence model to determine if transition to turbulence along the fish's surface leads to boundary layer separation. The expected results caused by adding these two features to earlier simulations do not occur. Thrust and drag are not found to be equal with usage of the thicker fish shape; instead both thrust and drag increase by 40-80% while diverging in value. Evidence of boundary layer separation is not present with usage of the LES turbulence model. Swimming energy efficiency is calculated to be 70% in both simulations. A brief analyses of boundary layer and downstream wake are included, showing general agreement with earlier studies. Limitations of the simulation are discussed. Future work regarding the author's preparation for an additional simulation of a rainbow trout utilizing a swimming method known as the Karman Gait is also considered. This preparation includes the creation of a 2-D grid domain and programs to define the kinematics of the fish and produce a specified vortex inlet condition.

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.

This dissertation investigates the possibility of using compact and heaterless calcium aluminate electride hollow cathodes in different electric propulsion systems for space applications. As conventional hollow cathodes generally require a heater to reach the high operating temperatures necessary to thermally emit electrons, research on low temperature heaterless hollow cathodes as electron sources has been increasing. Efforts at Technische Universität Dresden have resulted in an operational hollow cathode design that can be reliably used for low current plasma discharges. Hollow cathodes are crucial components in electric propulsion systems to ionize the propellant and neutralize the extracted ion beam. The successful development of an operational hollow cathode opens the possibility of using the design in different low-power electric propulsion systems. As the electron emission properties of C12A7:e- are still not well understood, a volume-averaged hollow cathode model has been developed as part of this thesis to obtain an improved insight into the plasma processes governing the cathode discharge. The model consists of two computational domains in which the plasma properties are volume-averaged. A lumped-node thermal model coupled with the plasma model provides the cathode temperature distribution for different operating points. The model moreover provides the discharge voltage which can be directly compared to experimental data. The thermal model was compared to thermal measurements to derive adequate values for free model parameters. The discharge voltage fits well for a 1 A discharge but diverges from measurement data at higher currents. The model is a starting point for further modeling efforts and needs to be verified using extensive plasma diagnostics. The first electric propulsion system developed as part of this thesis is an electrothermal device that takes advantage of high particle temperatures in a hollow cathode discharge. A performance model and preliminary test series were used to derive design parameters for a prototype that was used for an extensive parameter study. The thruster reliably generates thrust over a current range between 1 A – 3 A. The thrust achieved with this device is in the high micronewton to low millinewton range. The specific impulse is on the order of 100 s, which is low for electric propulsion systems, and the high discharge voltages of approximately 50 V limit the achievable efficiency to <1%. The second thruster concept is a DC discharge gridded ion thruster using a C12A7:e- hollow cathode as the discharge cathode and the neutral gas inlet. An analytical discharge model combined with a particle-in-cell simulation for ion extraction by electrostatically biased grids was used to design a modular testing prototype. The concept requires a low discharge current on the order of 200 mA. Operating the cathodes in a milliamp discharge current range proved to be difficult and was accompanied by high discharge voltages. Extracting an ion beam from the testing prototype was not successful. The third propulsion system is a magnetoplasmadynamic thruster (MPDT) that takes advantage of a strong magnetic field generated by permanent magnets and an orthogonal current in a plasma discharge using a C12A7:e- hollow cathode. Conventional MPDTs require high current discharges to generate a sufficiently strong self-induced magnetic field. The developed concept is a design alternative to expand the operational envelope to lower powers. A major advantage is the comparatively easy scalability of the device. One prototype for the low amp current range was developed and successfully operated. The generated thrust is in the low millinewton range with a specific impulse up to 1,200 s. The test series highlighted thermal problems with the design. Consequently, a sub-amp version of the concept was developed. The thruster was successfully operated but required high mass flow rates, lowering the specific impulse and efficiency.

To access and operate in space, a wide range of propulsion systems has been developed, from high-thrust chemical propulsion to low-thrust electrical propulsion, and new kind of systems are considered, such as solar sails and nuclear propulsion. Recently, interest in hybrid rocket engines has been renewed due to their attractive features (safe, cheap, flexible) and they are now investigated and developed by research laboratories such as ONERA.This master’s thesis work is in line with their development at ONERA and aims at finding a methodology to study numerically the liquid oxidizer injection using a Lagrangian solver for the liquid phase. For this reason, it first introduces a model for liquid atomiser developed for aeronautical applications, the FIMUR model, and then focuses on its application to a hybrid rocket engine configuration.The FIMUR model and the Sparte solver have proven to work fine with high mass flow rates on coarse grids. The rocket engine simulations have pointed out the need of an initialisation of the flow field. The methodology study has proven that starting with a reduced liquid mass flow rate is preferable to a simulation with a reduced relaxation between the coupled solvers. The former could not be brought to conclusion due to lack of time but gives an encouraging path to further investigate.

I samband med den globala uppvärmningen ökar efterfrågan för rena och förnybara bränslen alltmer i dagens samhälle. Eftersom flygindustrin idag är ansvarig för samma mängd växthusgaser som all motortrafik i Sverige, skulle ett byte till en avgasfri energikälla för flygfarkoster vara ett stort framsteg. Därför har projektet genom modellering framtagit ett hybridsystem av ett batteri och en bränslecell och undersökt hur kombinationen av olika storlekar på dem presterar i en driftcykel. Då batterier har hög specifik effekt men är tunga, kompletteras de med fördel av bränsleceller, som är lättviktiga och bidrar med uthållig strömförsörjning. På så sätt blir hybriden optimal för flygfarkoster. Kandidatarbetet är en del av projektet Green Raven, ett tvärvetenskapligt samarbete mellan instutitionerna Tillämpad Elektrokemi, Mekatronik och Teknisk Mekanik på Kungliga Tekniska Högskolan. Driftcykelmodelleringen gjordes i Simulink, och flera antaganden gjordes beträffande effektprofilen, samt bränslecellens mätvärden och effekt. Tre olika energihushållningsscheman skapades, vilka bestämde bränslecellseffekten beroende på vätgasnivån och batteriets laddningstillstånd. Skillnaden på systemen var vilka intervall av laddningstillstånd hos batteriet som genererade olika effekt hos bränslecellen.  Det bästa alternativet visade sig vara 0/100-systemet, eftersom det var det enda som inte orsakede någon degradering av bränslecellens kapacitet. / In today’s society, with several environmental challenges such as global warming, the demand for cleanand renewable fuels is ever increasing. Since the aviation industry in Sweden is responsible for the sameamount of greenhouse gas emissions as the motor traffic, a change to a non-polluting energy source forflying vehicles would be considerable progress. Therefore, this project has designed a hybrid system of abattery and a fuel cell and investigated how different combinations of battery and fuel cell sizes perform ina drive cycle, through computer modelling. As batteries possess a high specific power but are heavy, thefuel cells with high specific energy complement them with a sustained and lightweight power supply,which makes the hybrid perfect for aviation. The bachelor thesis is a part of Project Green Raven, aninterdisciplinary collaboration with the institutions of Applied Electrochemistry, Mechatronics andEngineering Mechanics at KTH Royal Institute of Techology. The drive cycle simulations were done inSimulink, and several assumptions regarding the power profile, fuel cell measurements and power weremade. Three different energy management strategies were set up, determining the fuel cell powerdepending on hydrogen availability and state of charge of the battery. The strategies were called 35/65,20/80 and 0/100, and the difference between them was at which state of charge intervals the fuel cellchanged its power output. The best strategy proved to be 0/100, since it was the only option which causedno degradation of the fuel cell whatsoever.