Cavitation is detrimental to the performance of ships and submarines, causing noise, erosion, and vibration. This study seeks to understand cavitation inception and delay on a typical ducted propulsor by utilizing the SimCenter's unstructured simulation and design system: U2NCLE. Specifically, three fundamental questions are addressed: 1. What are the macroscale flow physics causing cavitation inception? 2. How does cavitation inception scale with Reynolds number? 3. How can tip-leakage vortex cavitation inception be suppressed? To study the physics of cavitation inception, a ducted propulso simulation is developed and extensively validated with experimental results. The numerical method is shown to agree very well with experimental measurements made in the vortex core. It was discovered that the interaction of the leakage and trailing edge vortices cause the pressure to drop to a local minimum, providing ideal conditions for inception to occur. However, experimental observation shows that inception does not occur at the minimum pressure location, but rather at the point where the two vortices completely coalesce. At the point of coalescence, the simulation reveals that the streamwise core velocity decelerates, causing the air nuclei to stretch and burst. A Reynolds number scaling analysis is performed for the minimum pressure and maximum velocity in the vortex core. First, the numerical method is validated on a flate plate at various Reynolds numbers to assess the ability of typical turbulence models to predict Reynolds numbers ranging from one million to one billion. This scaling analysis methodology is then applied to the propulsor simulation, revealing that the minimum pressure in the vortex core is much less dependent on Reynolds number than was previously hypothesized. Lastly, to investigate means of delaying cavitation inception, the propulsor is parameterized and studied using design optimization theory. Concepts of vortex alleviation evident in nature are used to suggest suitable parameterizations. Also, dimension reduction is used to reduced the number of design variables. Finally, the concepts are implemented, evaluated, and shown to completely decouple the two vortices causing cavitation inception. Moreover, the minimum pressure in the vortex core is significantly increased.

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This work presents a preliminary design of a liquid rocket engine test platform to support research into liquid propulsion systems and rocket engine components, including injectors, ignition systems, combustion chambers and engine cooling systems. The liquid propellants, specifically liquid oxygen and ethanol, are pressure-fed using gaseous nitrogen. The test platform supports engine thrust values up to 1 kN, as well as varying oxidizer/fuel ratios up to 4.0 and varying ethanol concentrations between 70 and 100%. The test platform will integrate with a mobile control centre, which was designed concurrently, and provides remote control of the test procedures and data acquisition of all relevant pressure, temperature, mass flow and thrust data. The propellant feed assembly can support both cold and hot fire testing campaigns and is equipped with numerous safety features including inert gas purge lines, emergency drain lines and emergency shut-down and de-pressurization procedures.

Vacuum arc thrusters offer a relatively simple and cheap form of satellite propulsion, especially suitable for nanosatellites such as CubeSats or even smaller. This thesis focuses on vacuum arc thruster design considering the thruster’s manufacturing, assembly and integration into the spacecraft, and proposes a new anode geometry easing thruster production. Vacuum arc thruster research is traditionally experimental in nature due to a lack of accurate models. This work follows this approach, and studies experimentally the effect of several geometric design parameters on thruster performance. The outcome confrms findings from several papers, and suggests specifc improvements towards existing models for predicting the effect of the thruster’s geometry on its thrust. The chosen experimental approach raised the need for a micro-thrust measurement stand. Two distinct measurement stands have been designed, realized and used to test various thruster prototypes. One test stand is more accurate. However, the other setup allows for considerably faster testing.

An integrated aerodynamic-ramp-injector/plasma-torch-igniter of original design was tested in a Mâ = 2, unvitiated, heated flow facility arranged as a diverging duct scramjet combustor. The facility operated at a total temperature of 1000 K and total pressure of 330 kPa. Hydrogen (H2), ethylene (C2H4) and methane (CH4) were used as fuels, and a wide range of global equivalence ratios were tested. The main data obtained were wall static pressure measurements, and the presence of combustion was determined based on the pressure rises obtained. Supersonic and dual-mode combustion were achieved with hydrogen and ethylene fuel, whereas very limited heat release was obtained with the methane. Global operability limits were determined to be 0.07 < Ï < 0.31 for hydrogen, and 0.14 < Ï < 0.48 for ethylene. The hydrogen fuel data for the aeroramp/torch system was compared to data from a physical 10 unswept compression ramp injector and similar performance was found with the two arrangements. With hydrogen and ethylene as fuels and the aeroramp/plasma-torch system, the effect of varying the air total temperature was investigated. Supersonic combustion was achieved with temperatures as low as 530K and 680K for the two fuels, respectively. These temperatures are facility/operational limits, not combustion limits. The pressure profiles were analyzed using the Ramjet Propulsion Analysis (RJPA) code. Results indicate that both supersonic and dual-mode ramjet combustion were achieved. Combustion efficiencies varied with Ï from a high of about 75% to a low of about 45% at the highest Ï . With a theoretical diffuser and nozzle assumed for the configuration and engine, thrust was computed for each fuel. Fuel specific impulse was on average 3000 and 1000 seconds for hydrogen and ethylene respectively, and air specific impulse varied from a low of about 9 sec to a high of about 24 sec (for both fuels) for the To = 1000K test condition. The GASP RANS code was used to numerically simulate the injection and mixing process of the fuels. The results of this study were very useful in determining the suitability of the selected plasma torch locations. Further, this tool can be used to determine whether combustion is theoretically possible or not. / Ph. D.

A method in which the transport and evolution of the bubble population in a surface ship wake is numerically simulated is presented. The simulation is accomplished by constructing an advective-diffusive transport model for the scalar bubble field and solving this model for late times after ship passage. The bubble population model requires convection velocities and turbulent diffusion information that is supplied by solving the Reynolds-averaged parabolized Navier-Stokes equations with a <i>k</i> - ∊ turbulence model. The mean flow equations are solved by approximating the differential equations with a second order accurate finite difference scheme. The resulting large, sparse, banded matrix is solved by applying a version of the conjugate gradient method. The method has proven to be efficient and robust for the free shear flow problems of interest here. The simulation is initiated with given information in a plane at some point downstream of the ship from which the solution is propagated. The model is executed for a single and a twin propeller ship at 15 knots. The simulation shows that the development of the hydrodynamic and bubble near wake is dominated by ship geometry via strong advective transport. The far wake is dominated by diffusion and bubble rise and dissolution. Thus relatively large changes in geometry have a limited influence on the far wake. / Ph. D.

A theoretical model and associated computer program are developed to compute hydrodynamic interaction forces and moments on passing ships in a shallow asymmetric canal, by applying generalized Lagally's theorem. Steady lift force on the ships are estimated following standard slenderbody approach. Passing ships are assumed to travel at constant speeds, parallel to the canal walls. Ships are assumed to be slender, the fluid is ideal and wave making effects are ignored. Numerical calculations for several passing cases show good agreement with existing model test results. / M.S.

This thesis presents an identification scheme for the dynamic model of a ship at sea. We determine the form of the governing differential equations for a ship which is free to pitch and roll, but constrained in all other degrees of freedom, using a perturbation-energy technique. This technique approximates energy expressions and applies Lagrange's equations for quazi-coordinates to develop the equations of motion. When formulating the energies, we take advantage of the ship's symmetry to reduce the number of terms. The equations of motions are approximated such that they contain quadratic and cubic nonlinear terms. Having the form of the governing equations, we set up the parametric identification procedure. Using the method of multiple scales, we exploit resonances and obtain expressions containing subsets of the parameters to be identified. Then we outline a scheme which uses these expressions in conjunction with experimental data to identify the ship parameters. / M.S.

A new particle-in-cell algorithm was developed for plasma simulations involving complex boundary conditions. The new algorithm is based on the three-dimensional immersed finite element method which is developed in this thesis, and a modified legacy particle-in-cell code. The model also applies a new meshing technique that separates the field solution mesh from the particle pushing mesh in order to increase the computational eciency of the model. The new simulation model is used in two applications of great importance to the development of ion propulsion technology: the ion optics performance and the interaction between spacecraft and the ion thruster. The first application is ion optics simulations. Simulations are performed to investigate ion optics plasma flow for a whole subscale NEXT ion optics. The operating conditions modeled cover the entire cross-over to perveance limit range. The results of the ion optics simulations demonstrated good agreement with the available experimental data. The second application is ion thruster plume simulations. Simulations are performed to investigate ion thruster plume - spacecraft interactions for the Dawn spacecraft. Plume induced contaminations on the solar array are studied for a variety of ion thruster configurations including multiple thruster firings. / Ph. D.