Hypoplastic left heat syndrome (HLHS) is a complex cardiac malformation in neonates suffering from congenital heart disease and occurs in nearly 1 per 5000 births. HLHS is uniformly fatal within the first hours or days after birth as the severly malformed anatomies of the left ventricle, mitral, and aortic valves, and ascending aorta are not compatable with life. The regularly implemented treatment, the Norwood operation, is a complex open heart procedure that attempts to establish univenticular circulation by removing the atrial septum ( communicating the right and left ventricle), reconstructing the malformed aortic arch, and connecting the main pulmonary artery into the reconstructed arch to allow direct perfusion from the right ventricle into the systemic circulation. A relatively new treatment being utilized,the Hybrid Norwood procedure, involves a less invasive strategy to establish univentricular circulation that avoids a cardiopulmonary bypass (heart-lung machine), deliberate cardiac arrest, and circulatroy arrest of the patient during the procedure. The resulting systemic-pulmonary circulation is unconventional; blood is pumped simotaneously and in parallel to the systemic and pulmonary arteries after the procedure. Cardiac surgeons are deeply interested in understanding the global and local hemodynamics of this anotomical configuration. To this end, a multiscale model of the entire circulatory system was developed utilizing an electrical lumped parameter model for the peripheralor distal circulation coupled with a #D Computational Fluid Dynamics (CFD) model to understand the local hemodynamics. The lumped parameter (LP) model is mainly a closed loop circut comprised of RLC comartments that model cardiac function as well as the viscous drag, flow intertia, and compliance of the different atrial and venous beds in the body. A system of 32 first-order differential equations is formulated and solved for the LP model using a fourth-order adaptive Runge-Kutta solver. The output pressure and flow waveforms obtained from the LP model are imposed as boundary conditions on the CFD model. Coupling of the two models is done through an iterative process where the parameters in the LP model are adjusted to match the CFD solution. The CFD model domain is a representative HLHS anatomy of an infant after undergoing the Hybrid Norwood procedure and is comprised of the neo-aorta, pulmonary roots, aortic arch with branching arteries, and pulmonary arteries. The flow field is solved over several cardiac cycles using an implicit-unsteady RANS equation solver with the k-epsilon turbulence model.; Hypoplastic left heart syndrome (HLHS) is a complex cardiac malformation in neonates suffering from congenital heart disease and occurs in nearly 1 per 5000 births. HLHS is uniformly fatal within the first hours or days after birth as the severely malformed anatomies of the left ventricle, mitral and aortic valves, and ascending aorta are not compatible with life. The regularly implemented treatment, the Norwood operation, is a complex open heart procedure that attempts to establish univentricular circulation by removing the atrial septum (communicating the right and left ventricle), reconstructing the malformed aortic arch, and connecting the main pulmonary artery into the reconstructed arch to allow direct perfusion from the right ventricle into the systemic circulation. A relatively new treatment being utilized, the Hybrid Norwood procedure, involves a less invasive strategy to establish univentricular circulation that avoids a cardiopulmonary bypass (heart-lung machine), deliberate cardiac arrest, and circulatory arrest of the patient during the procedure. The resulting systemic-pulmonary circulation is unconventional; blood is pumped simultaneously and in parallel to the systemic and pulmonary arteries after the procedure. Cardiac surgeons are deeply interested in understanding the global and local hemodynamics of this anatomical configuration. To this end, a multiscale model of the entire circulatory system was developed utilizing an electrical lumped parameter model for the peripheral or distal circulation coupled with a 3D Computational Fluid Dynamics (CFD) model to understand the local hemodynamics. The lumped parameter (LP) model is mainly a closed loop circuit comprised of RLC compartments that model cardiac function as well as the viscous drag, flow inertia, and compliance of the different arterial and venous beds in the body.

Spacecraft thermal protection systems are at risk of being damaged due to airflow produced from Environmental Control Systems. There are inherent uncertainties and errors associated with using Computational Fluid Dynamics to predict the airflow field around a spacecraft from the Environmental Control System. This paper describes an approach to quantify the uncertainty in using Computational Fluid Dynamics to predict airflow speeds around an encapsulated spacecraft without the use of test data. Quantifying the uncertainty in analytical predictions is imperative to the success of any simulation-based product. The method could provide an alternative to traditional “validation by test only” mentality. This method could be extended to other disciplines and has potential to provide uncertainty for any numerical simulation, thus lowering the cost of performing these verifications while increasing the confidence in those predictions. Spacecraft requirements can include a maximum airflow speed to protect delicate instruments during ground processing. Computational Fluid Dynamics can be used to verify these requirements; however, the model must be validated by test data. This research includes the following three objectives and methods. Objective one is develop, model, and perform a Computational Fluid Dynamics analysis of three (3) generic, non-proprietary, environmental control systems and spacecraft configurations. Several commercially available and open source solvers have the capability to model the turbulent, highly three-dimensional, incompressible flow regime. The proposed method uses FLUENT, STARCCM+, and OPENFOAM. Objective two is to perform an uncertainty analysis of the Computational Fluid Dynamics model using the iv methodology found in “Comprehensive Approach to Verification and Validation of Computational Fluid Dynamics Simulations”. This method requires three separate grids and solutions, which quantify the error bars around Computational Fluid Dynamics predictions. The method accounts for all uncertainty terms from both numerical and input variables. Objective three is to compile a table of uncertainty parameters that could be used to estimate the error in a Computational Fluid Dynamics model of the Environmental Control System /spacecraft system. Previous studies have looked at the uncertainty in a Computational Fluid Dynamics model for a single output variable at a single point, for example the re-attachment length of a backward facing step. For the flow regime being analyzed (turbulent, three-dimensional, incompressible), the error at a single point can propagate into the solution both via flow physics and numerical methods. Calculating the uncertainty in using Computational Fluid Dynamics to accurately predict airflow speeds around encapsulated spacecraft in is imperative to the success of future missions

A comprehensive computational and experimental analysis has been conducted to characterize the flow dynamics and periodic structures formed in the wake of complex turning vanes. The vane packs were designed by the StreamVane swirl distortion generator technology, a design system that can efficiently reproduce swirl distortion for compressor rig and full turbofan engine testing. StreamVanes consist of an array of turning vanes that commonly contain variations in turning angle along their span, a nonaxisymmetric profile about the centerline, and vane-to-vane intersections or junctions to accurately generate the desired distortion. In this study, vane packs are considered complex if they contain two out of three of these features, a combination seen in other turbomachinery components outside of StreamVane design. Similar to all stator vanes or rotor blades, StreamVane vane packs are constructed using a series of cross-sectional airfoil profiles with blunt trailing edges and finite thicknesses. This, in turn, introduces periodic vortex structures in the wake, commonly known as trailing edge vortex shedding. To fully understand how the dynamics and coherent wake formations within vortex shedding impact both the flow distortion and structural durability of StreamVanes, it is first necessary to characterize the corresponding wakes in three dimensions. The current study provides an in-depth analysis to predict and measure the trailing edge vortex development using high-fidelity computational fluid dynamics and stereoscopic time-resolved particle image velocimetry experiments. Two testcase StreamVane geometries were specifically designed with complex features to evaluate their influence on the dynamics and coherence of the respective vane wakes. Fully three-dimensional, unsteady computational fluid dynamics simulations were performed using a Reynolds-Averaged Navier-Stokes solver coupled with a standard two-equation turbulence model and a hybrid, scale-resolving turbulence model. Both models predicted large-scale wake frequencies within 1—14% of experiment, with a mean difference of less than 3.2%. These comparisons indicated that lower fidelity simulations were capable of accurately capturing such flows for complex vane packs. Additionally, structural and modal analyses were conducted using finite element models to determine the correlations between dominant structural modes and dominant wake (flow) modes. The simulations predicted that vortex shedding modes generally contained frequencies 300% larger than dominant structural modes, and therefore, vortex induced vibrations were unlikely to occur. Lastly, mode decomposition methods were applied to the experimental results to extract energy ratios and reveal dynamic content across high-order wake modes. The vortex shedding modes generated more than 80% of the total wake energy for both complex vane packs, and dynamic decomposition methods revealed unique structures within the vane junction wake. In all analyses, comparisons were made between different vane parameters, such as trailing edge thickness and turning angle, where it was found that trailing edge thickness was the dominant vortex shedding parameter. The motivation, methodology, and results of the following research is presented to better understand the wake interactions, computational predictive capabilities, and structural dynamics associated with vortex shedding from complex vane packs. Although the results directly relate to StreamVane distortion generator technology, the qualitative and quantitative comparisons between the selected methods, geometry parameters, and flow conditions can be extrapolated to modern turbomachinery components in general. Therefore, this dissertation aims to benefit distortion generator and turbomachinery designers by providing insight into the underlying physics and overall modeling techniques of the wake dynamics in highly three-dimensional, complex components. / Doctor of Philosophy / A comprehensive analysis has been completed to characterize the unsteady wake flow produced by complex turning vane systems in three dimensions. Turning vanes are a common component utilized in the field of fluid dynamics and aerospace propulsion to effectively turn and manipulate the working fluid to the desired condition. For propulsion applications, similar vanes can alleviate performance losses by improving the overall aerodynamics and mitigating flow distortions entering the compressor of a jet engine. Conversely, complex turning vanes can also be used to reproduce the distortion for engineers to evaluate jet engine components when subjected to nonuniform flow ingestion. The distinct geometry features that make these vanes complex are also present in other turbomachinery systems outside of distortion generation. In any case, the cross-sectional profiles of the turning vanes commonly contain blunt ends or trailing edges due to engineering limitations and/or restrictions. This geometric feature introduces periodic wake structures, known as vortex shedding, that can negatively effect the performance of the overall system. It is therefore a necessity to characterize both the dynamics and coherence of vortex shedding to fully understand the flow features in highly three-dimensional flows. In the presented research, this is achieved by applying computational simulations and experimental measurements to extract the corresponding wake dynamics of complex vane packs. The selected testcases where designed using the StreamVane technology, a mature system that generates tailored turning vanes to reproduce flow distortion in jet engine or fan rig ground-testing facilities. The fluid simulations captured the expected wake flow and largescale structures convecting downstream of the vane packs. A comparison between two different flow models and the experimental results revealed minimal quantitative differences in the large-scale dynamics, which gave insight into the model selection to predict such flows. Additional structural simulations were performed to estimate the forcing and response of the vane packs when subjected to the aerodynamic loading. The results showed vortex shedding was highly unlikely to cause large amplitude vibrations and structural failures. In all analyses, the primary results were correlated with common vane parameters and operating conditions to evaluate their impact on the wake dynamics. The motivation, methodology, and results of the following research is presented to better understand the wake interactions, computational predictive capabilities, and structural dynamics associated with vortex shedding from complex vane packs. Although the results directly relate to StreamVane distortion generator technology, the qualitative and quantitative comparisons between the selected methods, geometry parameters, and flow conditions can be extrapolated to modern turbomachinery components in general. Therefore, this dissertation aims to benefit distortion generator and turbomachinery designers by providing insight into the underlying physics and overall modeling techniques of the wake dynamics in highly three-dimensional, complex components.

As a contribution to advancing turbofan engine ground test technology in support of propulsion system integration in modern conceptual aircraft, a novel inlet distortion generator (ScreenVaneTM) was invented. The device simultaneously reproduces combined inlet total pressure and swirl distortion elements in a tailored profile intended to match a defined turbofan engine inlet distortion profile. The device design methodology was intended to be sufficiently generic to be utilized in support of any arbitrary inlet distortion profile yet adequately specific to generate high-fidelity inlet distortion profile simulation. For the current investigation, a specific inlet distortion profile was defined using computational analysis of a conceptual boundary layer ingesting S-duct turbofan engine inlet. The resulting inlet distortion profile, consisting of both total pressure and swirl distortion elements, was used as the objective profile to be matched by the ScreenVane in a turbofan engine ground test facility. A ScreenVane combined inlet total pressure and swirl distortion generator was designed, computationally analyzed, and experimentally validated. The design process involved specifying a total pressure loss screen pattern and organizing a unique arrangement of swirl inducing turning vanes. Computational results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Computational full-field total pressure recovery and swirl angle profiles matched within approximately 1% and 2.5° (RMSD), respectively. Experimental turbofan engine ground test results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Experimental full-field total pressure recovery and swirl angle profiles matched within approximately 1.25% and 3.0° (RMSD), respectively. Following the successful reproduction of the S-duct turbofan engine inlet manufactured distortion profile, a turbofan engine response evaluation was conducted using the validated ScreenVane inlet distortion generator. Flow measurements collected at discrete planes immediately upstream and downstream of the fan rotor isolated the component for performance analysis. Based on the results of this particular engine and distortion investigation, the adiabatic fan efficiency was negligibly altered while operating with distorted inflow conditions when compared to nominal inflow conditions. Fuel flow measurements indicated that turbofan engine inlet air mass flow specific fuel consumption increased by approximately 5% in the presence of distortion. While a single, specific turbofan engine inlet distortion profile was studied in this investigation, the ScreenVane methodology, design practices, analysis approaches, manufacturing techniques, and experimental procedures are applicable to any arbitrary, realistic combined inlet total pressure and swirl distortion. / Doctor of Philosophy / As a contribution to advancing turbofan engine ground test technology in support of propulsion system integration in modern conceptual aircraft, a novel inlet distortion generator (ScreenVaneTM) was invented. The device simultaneously reproduces combined inlet total pressure and swirl distortion elements in a tailored profile intended to match a defined turbofan engine inlet distortion profile. The device design methodology was intended to be sufficiently generic to be utilized in support of any arbitrary inlet distortion profile yet adequately specific to generate high-fidelity inlet distortion profile simulation. For the current investigation, a specific inlet distortion profile was defined using computational analysis of a conceptual boundary layer ingesting S-duct turbofan engine inlet. The resulting inlet distortion profile, consisting of both total pressure and swirl distortion elements, was used as the objective profile to be matched by the ScreenVane in a turbofan engine ground test facility. A ScreenVane combined inlet total pressure and swirl distortion generator was designed, computationally analyzed, and experimentally validated. The design process involved specifying a total pressure loss screen pattern and organizing a unique arrangement of swirl inducing turning vanes. Computational and experimental results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Following the successful reproduction of the S-duct turbofan engine inlet manufactured distortion profile, a turbofan engine response evaluation was conducted using the validated ScreenVane inlet distortion generator. Flow measurements collected at discrete planes immediately upstream and downstream of the fan rotor isolated the component for performance analysis. Based on the results of this particular engine and distortion investigation, the adiabatic fan efficiency was negligibly altered while operating with distorted inflow conditions when compared to nominal inflow conditions. Fuel flow measurements indicated that turbofan engine inlet air mass flow specific fuel consumption increased in the presence of distortion. While a single, specific turbofan engine inlet distortion profile was studied in this investigation, the ScreenVane methodology, design practices, analysis approaches, manufacturing techniques, and experimental procedures are applicable to any arbitrary, realistic combined inlet total pressure and swirl distortion.

There is increasing interest in design methods and performance prediction for turbine engines operating at low Reynolds numbers. In this regime, boundary layer separation may be more likely to occur in the turbine flow passages. For accurate CFD predictions of the flow, correct modeling of laminar-turbulent boundary layer transition is essential to capture the details of the flow. To investigate possible improvements in model fidelity, both two-dimensional and three-dimensional CFD models were created for the flow over several low pressure turbine blade designs. A new three-equation eddy-viscosity type turbulent transitional flow model originally developed by Walters and Leylek was employed for the current RANS CFD calculations. Flows over three low pressure turbine blade airfoils with different aerodynamic characteristics were simulated over a Reynolds number range of 15,000-100,000, and predictions were compared to experiments. The turbulent transitional flow model sensitivity to inlet turbulent flow parameters showed a dependence on free-stream turbulence intensity and turbulent length scale. Using the total pressure loss coefficient as a measurement of aerodynamic performance, the Walters and Leylek transitional flow model produced adequate prediction of the Reynolds number performance in the Lightly Loaded blade. Furthermore, the correct qualitative flow response to separated shear layers was observed for the Highly Loaded blade. The vortex shedding produced by the separated flow was largely two-dimensional with small spanwise variations in the separation region. The blade loading and separation location was sufficiently predicted for the Aft-Loaded L1A blade flowfield. Investigations of the unsteady flowfield of the Aft-Loaded L1A blade showed the shear layer produced a large separation region on the suction surface. This separation region was located more downstream and significantly reduced in size when impinged upon by the upstream wakes, thus improving the aerodynamic performance consistent with experiments. For all cases investigated, the Walters and Leylek transitional flow model was judged to be sufficient for understanding the separation and transition characteristics, and superior to other widely-used turbulence models in accuracy of describing the details of the transitional and separated flow. This research characterized and assessed a new model for low Reynolds number turbine aerodynamic flow prediction and design improvement. / Ph. D.

The first goal of this work was to develop models based on nonlinear ordinary-differential equations or nonlinear algebraic equations, which produce the lift and drag coefficients on a cylinder or a cylinder-like structure. We introduced an improved wake oscillator for the lift, which combines the van der Pol and Duffing equations. We proposed a two-term quadratic model that relates the drag and lift coefficients, which reproduces the phase relationship between the drag and lift and its variation with the Reynolds number. We found that a mixed-type (external and parametric) forcing is needed to represent the effects of the cylinder motion. The second goal of this work was to develop a deeper understanding of the shedding and fluid forces on a cylinder and how they depend on its oscillatory motion within and outside the synchronization (or lock-in) band of frequencies. We performed extensive CFD (computational fluid dynamics) simulations and solved the unsteady Reynolds-averaged Navier-Stokes equations that govern the flow fields around fixed and moving (in either the cross-flow or in-line direction) cylinders. We identified various wake modes that can exist, depending on the cylinder motion (direction, amplitude, and frequency) by using modern methods of nonlinear dynamics. The possible responses can be period-one, periodic with large period, quasiperiodic, or chaotic. Moreover, we found that the route to chaos is torus breakdown. We investigated how four frequency sweeps of the cross-flow motion affect the response curves and the hysteresis phenomenon. We studied in detail the effect of the in-line motion on the wake and related this effect to the reduction in the lift and mean drag due to a synchronization type that is very different from the one due to cross-flow motion. / Ph. D.

A new noise metric has been developed that may be used for optimization problems involving aerodynamic noise from a clean wing. The modeling approach uses a classical trailing edge noise theory as the starting point. The final form of the noise metric includes characteristic velocity and length scales that are obtained from three-dimensional, steady, RANS simulations with a two- equation k-omega turbulence model. The noise metric is not the absolute value of the noise intensity, but an accurate relative noise measure as shown in the validation studies. One of the unique features of the new noise metric is the modeling of the length scale, which is directly related to the turbulent structure of the flow at the trailing edge. The proposed noise metric model has been formulated so that it can capture the effect of different design variables on the clean wing airframe noise such as the aircraft speed, lift coefficient, and wing geometry. It can also capture three-dimensional effects which become important at high lift coefficients, since the characteristic velocity and the length scales are allowed to vary along the span of the wing. Noise metric validation was performed with seven test cases that were selected from a two-dimensional NACA 0012 experimental database. The agreement between the experiment and the predictions obtained with the new noise metric was very good at various speeds, angles of attack, and Reynolds Number, which showed that the noise metric is capable of capturing the variations in the trailing edge noise as a relative noise measure when different flow conditions and parameters are changed. Parametric studies were performed to investigate the effect of different design variables on the noise metric. Two-dimensional parametric studies were done using two symmetric NACA four-digit airfoils (NACA 0012 and NACA 0009) and two supercritical (SC(2)-0710 and SC(2)-0714) airfoils. The three-dimensional studies were performed with two versions of a conventional transport wing at realistic approach conditions. The twist distribution of the baseline wing was changed to obtain a modified wing which was used to investigate the effect of the twist on the trailing edge noise. An example study with NACA 0012 and NACA 0009 airfoils demonstrated a reduction in the trailing edge noise by decreasing the thickness ratio and the lift coefficient, while increasing the chord length to keep the same lift at a constant speed. Both two- and three-dimensional studies demonstrated that the trailing edge noise remains almost constant at low lift coefficients and gets larger at higher lift values. The increase in the noise metric can be dramatic when there is separation on the wing. Three-dimensional effects observed in the wing cases indicate the importance of calculating the noise metric with a characteristic velocity and length scale that vary along the span. The twist change does not have a significant effect on the noise at low lift coefficients, however it may give significant noise reduction at higher lift values. The results obtained in this study show the importance of the lift coefficient on the airframe noise of a clean wing and favors having a larger wing area to reduce the lift coefficient for minimizing the noise. The results also point to the fact that the noise reduction studies should be performed in a multidisciplinary design and optimization framework, since many of the parameters that change the trailing edge noise also affect the other aircraft design requirements. It's hoped that the noise metric developed here can aid in such multidisciplinary design and optimization studies. / Ph. D.

Uncertainty is prevalent throughout the design, analysis, and optimization of aerospace products. When scientific computing is used to support these tasks, sources of uncertainty may include the freestream flight conditions of a vehicle, physical modeling parameters, geometric fidelity, numerical error, and model-form uncertainty, among others. Moreover, while some uncertainties may be treated as probabilistic, aleatory sources, other uncertainties are non-probabilistic and epistemic due to a lack of knowledge, and cannot be rigorously treated using classical statistics or Bayesian approaches. An additional complication for propagating uncertainty is that many aerospace scientific computing tools may be computationally expensive; for example, a single high-fidelity computational fluid dynamics solution may require several days or even weeks to complete. It is therefore necessary to employ uncertainty propagation strategies that require as few solutions as possible. The Non-Intrusive Polynomial Chaos (NIPC) method has grown in popularity in recent decades due to its ability to propagate both aleatory and epistemic parametric sources of uncertainty in a computationally efficient manner. While traditional Monte Carlo methods might require thousands to millions of function evaluations to achieve statistical convergence, NIPC typically requires tens to hundreds for problems with similar numbers of uncertain dimensions. Despite this efficiency, NIPC is limited in one important aspect: it can only propagate uncertainty at a particular point in a design space or flight envelope. For optimization or aerodynamic database problems that require uncertainty estimates at many more than one point, the use of NIPC quickly becomes computationally intractable. This dissertation introduces a new method entitled Spatially Accurate Polynomial Chaos (SAPC) that extends the original NIPC approach for the spatial regression of aleatory and epistemic parametric sources of uncertainty. Throughout the dissertation, the SAPC method is applied to various aerospace problems of interest. These include the regression of aerodynamic force and moment uncertainties throughout the flight envelope of a commercial aircraft, the design under uncertainty of a two-stream propulsive mixer device, and the robust design of a low-boom supersonic demonstrator aircraft. Collectively the results suggest that SAPC may be useful for a large variety of engineering applications. / Doctor of Philosophy / Uncertainty is prevalent throughout the design, analysis, and optimization of aerospace products. When scientific computer simulations are used to support these tasks, sources of uncertainty may include the speed of an aerospace vehicle, the direction of the wind, physical modeling constants or assumptions, and the vehicle shape, among others. As a result of these sources uncertainty, assessments of vehicle performance are also uncertain. For example, if the speed of a vehicle is not known precisely, then computer simulations will predict a lift force which is also imprecisely known. A challenge when assessing the uncertainty in aerospace vehicle performance is that the computer simulations which predict performance may take a long time to run, even on state-of-the-art super computers. Traditional statistical methods may require thousands or millions of simulations for the prediction of uncertainty, which does not fit within the computational budget of most aerospace analyses. A newer method called Non-Intrusive Polynomial Chaos (NIPC) is more efficient, typically requiring only tens to hundreds of simulations; however, NIPC only provides uncertainty estimates at a single point in an aircraft flight envelope or design condition. In this dissertation, a new method called Spatially Accurate Polynomial Chaos (SAPC) is developed. The SAPC method combines desirable features of NIPC with regression methods for an efficient estimation of uncertainty throughout a vehicle flight envelope or design space. Throughout the dissertation, the SAPC method is applied to various aerospace problems of interest. These include the regression of aerodynamic force and moment uncertainties throughout the flight envelope of a commercial aircraft, the design under uncertainty of a two-stream propulsive mixer device, and the robust design of a low-boom supersonic demonstrator aircraft. Collectively the results suggest that SAPC may be useful for a large variety of engineering applications.

The ability of Reynolds Averaged Navier Stokes (RANS) models to predict the characteristics of a non-Body of Revolution (non-BOR) Ellipsoidal model is studied to establish the feasibility of utilizing RANS as a non-BOR concept design tool. Data unable to be obtained experimentally, such as streamwise and spanwise pressure gradients and yaw turn boundary layer characteristics, are also established. A range of conditions are studied including ahead, pitched up, steady 10 and 15 degree yaw turns, and unsteady 10 and 15 degree yaw turns. Simulation results show good agreement for ahead and pitched forces and moments. Straight ahead skin friction values also showed good agreement, providing even improved agreement over an LES model which utilized wall functions. Yaw turn conditions also showed good agreement for roll angles up to 10 degrees. Steady maneuvering forces and moments showed good agreement up to 10 degrees roll and separation calculations also showed good agreement up to 10 degrees roll. Unsteady maneuvering characteristics showed mixed results, with the normal force and pitching moment trends generally agreeing with experimental data, whereas the unsteady rolling moment did not tend to follow experimental trends. Two primary conditions, the change in curvature between the mid-body and elliptical ends and the accuracy of modeling of 3D flows with RANS, are discussed as sources of discrepancies between the experimental data and steady simulations greater than 10 degrees roll and unsteady rolling simulations. / Master of Science

Consider a bottom-hinged Oscillating Wave Surge Converter (OWSC): This device oscillates due to the hydrodynamic forces applied on it by the action of ocean waves. The focus of this thesis is to build upon the in-house multi-block generalized coordinate finite volume solver GenIDLEST using a collocated grid arrangement within the framework of the fractional-step method to make it compatible to simulate such systems. The first step in this process is to deploy a convection scheme which differentiates between air and water. This process is further complicated by the 1:1000 density and 1:100 viscosity ratio between the two fluids. For this purpose, a phase field method is chosen for its ease of implementation and proven boundedness and conservativeness properties. Extensive validation and verification using standard test cases, such as droplet in shear flow, Rayleigh Taylor instability, and the Dam Break Problem is carried out. This development is then coupled with the present Immersed Boundary Module which is used to simulate the presence of moving bodies and again verified against test cases, such as the Dam Break problem with a vertical obstacle and heave decay of a partially submerged buoyant cylinder. Finally, a relaxation zone technique is used to generate waves and a numerical beach technique is used to absorb them. These are then used to simulate the Oscillating Surge Wave Converter. / Master of Science / An Oscillating Wave Surge Converter can be best described as a rectangular flap, hinged at the bottom, rotating under the influence of ocean waves from which energy is harvested. The singular aim of this thesis is to model this device using Computational Fluid Dynamics (CFD). More specifically, the aim is to model this dynamic device with the full Navier Stokes Equations, which include inertial forces, arising due to the motion of the fluid, viscous forces which dissipate energy, and body forces such as gravity. This involves three key steps: 1. Modelling the air-water interface using a convection scheme. A phase field method is used to differentiate between the two fluids. This task is made more challenging because of the very large density and viscosity differences between air and water. 2. Model dynamic moving geometries in a time-dependent framework. For this, we rely on the Immersed Boundary Method. 3. Develop a numerical apparatus to generate and absorb ocean waves. For this, we rely on the Relaxation Zone and Numerical Beach Method. These developments are validated in different canonical problems and finally applied to a two-dimensional oscillating surge wave energy converter.