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Administration of nasal cannula therapy tasks providers with periodically monitoring their patients and adjusting settings according to patient needs. Conventionally, providers monitor a patient's oxygen demand using pulse oximetry and a qualitative assessment of the patient's work of breathing. The motivation for this research is to augment the traditional qualitative assessment of work of breathing with a quantitative measurement of a patient's tidal volume, the volume of air inhaled with each breath. This thesis presents a novel approach to measure tidal volume using a nasal cannula with built-in pressure sensors. Pressure waveforms obtained from continuous measurement of the pressure at the tip of the cannula are used to estimate nasal flowrates, and these nasal flowrates are time-integrated to estimate tidal volumes. Computational fluid dynamics (CFD) models were used to simulate fluid flow in a simplified nasal passage undergoing nasal cannula therapy. These simulations used a range of flow conditions characteristic of both low-flow and high-flow nasal cannula treatments. The simulations produced a transformation from cannula tip pressure to instantaneous nasal flowrate, and this transformation was evaluated using a matching empirical experiment. This empirical experiment used a matching physical geometry with a similar range of flow conditions, and the transformation obtained from CFD was able to estimate the actual tidal volumes with 85% accuracy. This study showed that continuous pressure measurement at the tip of a nasal cannula produces enough information to estimate nasal flowrates and tidal volumes. No similar studies were found during the literature review, so an accuracy of 85% is promising for this stage. If this technique could be made more accurate and deployed in an unobtrusive way, the resulting nasal cannula device could be used to continuously, comfortably monitor patients' tidal volumes. / Master of Science / Oxygen therapy is the most common prescription in hospitals across the United States, and the most common form of oxygen therapy is nasal cannula therapy. Administration of nasal cannula therapy requires providers to periodically assess their patients' oxygen saturations and work of breathing. Oxygen saturation can be quantitatively monitored using pulse oximetry but work of breathing must be qualitatively monitored using visual exams or walking tests. The motivation of this research is to augment this qualitative assessment with a quantitative metric. In our research, we chose the volume of inhaled air (the "tidal volume") as a proxy metric for a patient's work of breathing. This thesis presents our attempt to use a nasal cannula augmented with pressure sensors to estimate the tidal volume of a mannequin undergoing nasal cannula therapy. Our concept is that more intense inhalations/exhalations produce larger pressure swings at the tip of the nasal cannula. For this proof-of-concept study, a simplified nasal passage geometry was used. Pressure waveforms obtained from continuous measurement of the pressure at the tip of the cannula are used to estimate nasal flowrates, and these nasal flowrates are time-integrated to estimate tidal volumes. Computational fluid dynamics (CFD) simulations were used to predict how the cannula tip pressure changes as a function of nasal flowrates and cannula flowrates, then this relationship was tested using a matching empirical experiment. This matching empirical experiment showed that our technique of estimating tidal volumes was 85% accurate. This study showed that continuous pressure measurement at the tip of a nasal cannula produces enough information to estimate nasal flowrates and tidal volumes. No similar studies were found during the literature review, so an accuracy of 85% is promising for this stage. If this technique could be made more accurate and deployed in an unobtrusive way, the resulting nasal cannula device could be used to continuously, comfortably monitor patients' tidal volumes.

Climate change and rising temperatures is a well-known problem. To tackle global warming a transition from fossil fuels to renewable and reliable energy sources is necessary. Hydrogen, in fuel cells, is proposed to replace diesel and gasoline in the transport sector. Hydrogen is a pure fuel and the fuel cells only emit water and heat as a byproductbyproduct. Combined with electric motors, the hydrogen fuel cell can be 2-3 times more efficient compared to combustion engines fueled by gasoline. The performance of the fuel cell is affected by how the individual parts of the cell are designed. There are some difficulties in manufacturing complex geometries which requires require a forming in more than one step.  The goal isis to investigate, with the help of COMSOL Multiphysics software, how the performance of the fuel cell is affected by the shaped geometry at the cathode side of the flow plate. A numerical model is developeding will be made with varying parameters on the measurement of the cross-sections of the channels where pressure drop and flow distribution for ten different geometries isarewill be investigated. The model iswas built in the COMSOL Multiphysics 6.1 Software and includes a three-dimensional geometry consisting of a gas channel and a gas diffusion layer. The flow is laminar and the gas diffusion layer is set as a porous medium.  The results show that geometries with less sharp edges have lower pressure drops and more uniform flow distribution compared to geometries with sharper edges. The geometry with the sharpest edges has the highest pressure drop of 4.8 Pa/mm and the geometry with rounder edges has the lowest of 3.8 Pa/mm. A relationship between pressure drop and cross-sectional area can be found. With increasing radius and increasing cross-sectional area will the pressure drop decrease. The Reynolds number is higher for sharper geometries since the average velocity in the channels is higher, which also gives a lower friction factor. The length of the top flat becomes less for rounder geometries, which positively affects uniform flow distribution. The geometries with rounder edges have the most uniform distribution at the top of the gas diffusion layer and the sharpest geometry has the least uniform distribution. The deviation from the mean velocity is lower for sharper geometries, mainly because the velocities in the gas diffusion layer are lower. Sensitivity analysis was made over the mass flow rate and mesh, showing that the pressure drop is proportional to the mass flow rate and it becomes higher with less fine mesh.  Less fine mesh also gives lower velocities in the gas diffusion layer. Further studies can be made on how the gas diffusion layer behaves in the fuel cell when adding clamping force to the stack when putting it together and investigate if and how it affects pressure drop and flow distribution. The environmental benefit can be crucial if the performance of the fuel cells improves and motivates the investments which is are needed for, among other things, the infrastructure. / Klimatförändringar och stigande temperaturer är ett välkänt problem. För att ta itu med den globala uppvärmningen är en övergång från fossila bränslen till förnybara och pålitliga energikällor nödvändigt. Vätgas,  går bland annat att användas ii bland annat bränsleceller och , skulle kunna ersätta diesel och bensin inom transportsektorn. Vätgas är ett rent bränsle och bränslecellerna släpper bara ut vatten och värme som biprodukter. I kombination med elmotorer kan vätgasbränslecellen vara 2–3 gånger mer effektiva jämfört med förbränningsmotorer som drivs av bensin. Bränslecellens prestanda påverkas av hur de enskilda delarna av cellen är utformade. Det finns vissa svårigheter att tillverka komponenter med komplexa geometrier som kräver formning i fler än ett steg. Målet är att med hjälp av programvaran COMSOL Multiphysics undersöka hur bränslecellens prestanda påverkas av den formade geometrin på katodsidan av flödesplattan. En numerisk modellering kommer att utförasgöras utifrån med varierande parametrar därpå måtten hos kanalernas tvärsnitt varieras. Tdär tryckfall och flödesfördelning hos tio olika geometrier kommer att undersökas. Modellen byggdes i COMSOL Multiphysics 6.1 Software och inkluderar en tredimensionell geometri bestående av en gaskanal och ett gasdiffusionsskikt. Flödet är laminärt och gasdiffusionsskiktet antas vara ett poröst medium. Resultaten visar att geometrier med mindre skarpa kanter ger lägre tryckfall och jämnare flödesfördelning jämfört med geometrier med skarpare kanter. Geometrin med skarpast kanter har det högsta tryckfallet på 4.8 Pa/mm och geometrin med rundare kanter har ett tryckfall på 3.8 Pa/mm. Ett samband mellan tryckfall och tvärsnittsarea kan hittas då ökad radie och ökad tvärsnittsarea ger en minskning i tryckfall. Reynoldstalet är högre för skarpare geometrier eftersom medelhastigheten i kanalerna är högre, vilket också ger en lägre friktionsfaktor. Längden på toppen av kanalerna blir mindre för rundare geometrier, vilket påverkar flödesfördelningen positivt. Geometrierna med rundare kanter har den mest jämna fördelningen i toppen av gasdiffusionsskiktet och den skarpaste geometrin har den minst jämna fördelningen. Avvikelsen från medelhastigheten är lägre för skarpare geometrier, främst på grund av att hastigheterna i gasdiffusionslagret är lägre. Känslighetsanalys gjordes över storleken på massflödet och noggrannheten i meshen, vilket visar att tryckfallet är proportionellt mot massflödet och att det blir högre med mindre noggrann mesh. Mindre noggrann mesh ger också lägre hastigheter i gasdiffusionsskiktet. Ytterligare studier kan göras om hur gasdiffusionslagret beter sig i bränslecellen vid sammanpressning av alla delar i cellen och undersöka om och hur det påverkar tryckfall och flödesfördelning. Fördelen för miljönMiljönyttan kan vara stor om bränslecellernas prestanda förbättras och på så vis kan motivera de investeringar som behövs för utbyggnaden av bland annat infrastrukturen.

Engineering applications of unconventional fuels like HFOs require a detailed understanding of the physics associated with their evaporation. The processing of HFOs involves forming a spray; therefore, studying droplets is of particular interest. The work described in this dissertation tackles two of the most obscure aspects associated with HFOs modelling. The first aspect is the identification of a valid chemical description of the structure of the fuel. In particular, the author focused on finding a methodology that allows identifying a discrete surrogate to describe the complex pool of molecules of which the fuels are made. The second part of the work was devoted to understand and model thermally-induced secondary breakup, which is the primary cause of deviation from the "d2" that multi-component droplet experience. The formulation of a surrogate was successfully achieved by developing and implementing a new algorithm that allows building a surrogate from a set of easily accessible physical properties. A new methodology for the post-processing of experimental data was formulated. The methodology consists of studying the evolution of the normalized distance of the interface from the droplet’s centroid instead of its diameter. The new approach allowed the separation between interface deformation and expansion/shrinking. The information was then processed using the dynamic mode decomposition to separate the stochastic contribution associated with secondary atomization and the deterministic contribution of vaporization. Finally, thermally induced secondary atomization was studied using a CFD code appositely developed. The code is based on the geometric Volume of Fluid (VoF) method and consists of a compressible, multi-phase, multi-component solver in which phase change is considered. The novelty in the proposed approach is that the evaporation source term and the surface tension force are evaluated directly from the geometrically reconstructed interface. The code was validated against the exact solution of analytically solvable problems and experimental data. The solver was then used to study HFO secondary breakup and perform a parametric analysis that helped to understand the problem’s physics. A possible application of this framework is the formulation of sub-models to be applied in spray calculations.

This thesis reports on a numerical investigation into the three-dimensional flow divergence around a high solidity vertical axis wind turbine. Three-dimensional unsteady Reynolds averaged Navier-Stokes simulations of an H-type vertical axis wind turbine were used to examine the impact of turbine aspect ratio and tip speed ratio on the flow divergence. The turbine height was changed to alter the turbine aspect ratio, while keeping the diameter constant, to ensure that the solidity and tip speed ratio values were comparable between the different aspect ratios tested. The power output of the turbine consistently increased with aspect ratio and the optimal tip speed ratio for peak performance was negligibly affected. The flow divergence results showed that larger aspect ratio turbines had significantly more flow divergence with a 1 m/s entrance velocity difference between the smallest and largest cases. These two results where contradictory as a larger aspect ratio turbine was more efficient even though it had a smaller fraction of the upstream flow entering the upwind pass. The reason for this result was that impact of the tip effects, which caused a power reduction near the end of the blades. The distance from the blade tips that experienced a power reduction was constant for turbines of aspect ratio one and greater, resulting in a smaller turbine having a greater fraction of its height effected by the tips. This caused the overall power output for a smaller aspect ratio turbine to be lower even though its centre performance was higher, due to an increased entrance velocity. The change in flow divergence with tip speed ratio was also examined to better understand the driving force behind the divergence. It was found that the turbine power output was not the direct cause of flow divergence. The blade forces, specifically the force generated in the upstream direction had a strong linear correlation with the upstream flow divergence. / Thesis / Master of Applied Science (MASc)

Yes / In a novel development on previous computational fluid dynamics studies, the work reported here used an Eulerian two-fluid model with the shear stress transport k–ω turbulence closure model and bubble interaction models to simulate aeration tank performance at full scale and to identify process performance issues resulting from design parameters and operating conditions. The current operating scenario was found to produce a fully developed spiral flow. Reduction of the airflow rates to the average and minimum design values led to a deterioration of the mixing conditions and formation of extended unaerated fluid regions. The influence of bubble-induced mixing on the reactor performance was further assessed via simulations of the residence time distribution of the fluid. Internal flow recirculation ensured long contact times between the phases; however, hindered axial mixing and the presence of dead zones were also identified. Finally, two optimization schemes based on modified design and operating scenarios were evaluated. The adjustment of the airflow distribution between the control zones led to improved mixing and a 20% improvement to the mass transfer coefficient. Upgrading the diffuser grid was found to be an expensive and ineffective solution, leading to worsening of the mixing conditions and yielding the lowest mass transfer coefficient compared with the other optimization schemes studied. / College of Engineering and Physical Sciences, University of Birmingham, UK

Multidisciplinary interactions are expected to play a significant role in the design of future high-performance aircraft (Blended-Wing Body, Truss-Braced wing, High Speed Civil transport, High-Altitude Long Endurance aircraft and future military aircraft). Also, the availability of supercomputers has made it now possible to employ high-fidelity models (Computational Fluid Dynamics for fluids and detailed finite element models for structures) at the preliminary design stage. A necessary step at that stage is to calculate the wing angle-of-attack at which the wing will generate the desired lift for the specific flight maneuver. Determination of this angle, a simple affair when the wing is rigid and the flow regime linear, becomes difficult when the wing is flexible and the flow regime non-linear. To solve this inherently nonlinear problem, a Newton's method type algorithm is developed to simultaneously calculate the deflection and the angle of attack. The present algorithm requires the sensitivity of the aerodynamic pressure with respect to each of the generalized displacement coordinates needed to represent the structural displacement. This sensitivity data is easy to determine analytically when the flow regime is linear. The present algorithm uses a finite difference method to obtain these sensitivities and thus requires only the pressure data and the surface geometry from the aerodynamic model. This makes it ideally suited for nonlinear aerodynamics for which it is difficult to obtain the sensitivity analytically. The present algorithm requires the CFD code to be run for each of the generalized coordinates. Therefore, to reduce the number of generalized coordinates considerably, we employ the modal superposition approach to represent the structural displacements. Results available for the Aeroelastic Research Wing (ARW) are used to evaluate the performance of the modal superposition approach. Calculations are made at a fixed angle of attack and the results are compared to both the experimental results obtained at NASA Langley Research Center, and computational results obtained by the researchers at NASA Ames Research Center. Two CFD codes are used to demonstrate the modular nature of this research. Similarly, two separate Finite Element codes are used to generate the structural data, demonstrating that the algorithm is not dependent on using specific codes. The developed algorithm is tested for a wing, used for in-house aeroelasticity research at Boeing (previously McDonnell Douglas) Long Beach. The trim angle of attack is calculated for a range of desired lift values. In addition to the Newton's method algorithm, a non derivative method (NDM) based on fixed point iteration, typical of fixed angle of attack calculations in aeroelasticity, is employed. The NDM, which has been extended to be able to calculate trim angle of attack, is used for one of the cases. The Newton's method calculation converges in fewer iterations, but requires more CPU time than the NDM method. The NDM, however, results in a slightly different value of the trim angle of attack. It should be noted that NDM will converge in a larger number of iterations as the dynamic pressure increases. For one value of the desired lift, both viscous and inviscid results were generated. The use of the inviscid flow model while not resulting in a markedly different value for the trim angle of attack, does result in a noticeable difference both in the wing deflection and the span loading when compared to the viscous results. A crude (coarse-grain) parallel methodology was used in some of the calculations in this research. Although the codes were not parallelized, the use of modal superposition made it possible to compute the sensitivity terms on different processors of an IBM SP/2. This resulted in a decrease in wall clock time for these calculations. However, even with the parallel methodology, the CPU times involved may be prohibitive (approximately 5 days per Newton iteration) to any practical application of this method for wing analysis and design. Future work must concentrate on reducing these CPU times. Two possibilities: (i) The use of alternative basis vectors to further reduce the number of basis vectors used to represent the structural displacement, and (ii) The use of more efficient methods for obtaining the flow field sensitivities. The former will reduce the number of CFD analyses required the latter the CPU time per CFD analysis. NOTE: (03/2007) An updated copy of this ETD was added after there were patron reports of problems with the file. / Ph. D.

A mathematical model for the determination of the aerodynamic forces acting on an aircraft is presented. The mathematical model is based on the generalization of the idea of aerodynamically steady motions. One important use of these results is the determination of steady (time-invariant) aerodynamic forces and moments. Such aerodynamic forces can be determined using computer simulation by determining numerically the associated steady flows around the aircraft when it is moving along such generalized steady trajectories. The method required the extension of standard (inertial) CFD formulations to general non-inertial reference frames. Generalized Navier-Stokes and Euler equations have been derived. The formulation is valid for all ranges of Mach numbers including transonic flow. The method was implemented numerically for the planar case using the generalized Euler equations. The developed computer codes can be used to obtain numerical flow solutions for airfoils moving in general steady motions (i.e. circular motions). From these numerical solutions it is possible to determine the variation of the lift, drag and pitching moment with respect to the pitch rate at different Mach numbers and angles of attack. One of the advantages of the mathematical model developed here is that the aerodynamic forces become well-defined functions of the motion variables (including angular rates). In particular, the stability derivatives are associated with partial derivatives of these functions. These stability derivatives can be computed using finite differences or the sensitivity equation method. / Ph. D.

Computational solutions are presented for flows ranging from incompressible viscous flows to inviscid transonic flows. The viscous flow problems are solved using the incompressible Navier-Stokes equations while the inviscid solutions are attained using the full potential equation. Results for the viscous flow problems focus on turbulence modeling when separation is present. The main focus for the inviscid results is the development of an unstructured solution algorithm. The subject dealing with turbulence modeling for separated flows is discussed first. Two different test cases are presented. The first flow is a low-speed converging-diverging duct with a rapid expansion, creating a large separated flow region. The second case is the flow around a stationary hydrofoil subject to small, oscillating hydrofoils. Both cases are computed first in a steady state environment, and then with unsteady flow conditions imposed. A special characteristic of the two problems being studied is the presence of strong adverse pressure gradients leading to flow detachment and separation. For the flows with separation, numerical solutions are obtained by solving the incompressible Navier-Stokes equations. These equations are solved in a time accurate manner using the method of artificial compressibility. The algorithm used is a finite volume, upwind differencing scheme based on flux-difference splitting of the convective terms. The Johnson and King turbulence model is employed for modeling the turbulent flow. Modifications to the Johnson and King turbulence model are also suggested. These changes to the model focus mainly on the normal stress production of energy and the strong adverse pressure gradient associated with separating flows. The performance of the Johnson and King model and its modifications, along with the Baldwin-Lomax model, are presented in the results. The modifications had an impact on moving the flow detachment location further downstream, and increased the sensitivity of the boundary layer profile to unsteady flow conditions. Following this discussion is the numerical solution of the full potential equation. The full potential equation assumes inviscid, irrotational flow and can be applied to problems where viscous effects are small compared to the inviscid flow field and weak normal shocks. The development of a code is presented which solves the full potential equation in a finite volume, cell centered formulation. The unique feature about this code is that solutions are attained on unstructured grids. Solutions are computed in either two or three dimensions. The grid has the flexibility of being made up of tetrahedra, hexahedra, or prisms. The flow regime spans from low subsonic speeds up to transonic flows. For transonic problems, the density is upwinded using a density biasing technique. If lift is being produced, the Kutta-Joukowski condition is enforced for circulation. An implicit algorithm is employed based upon the Generalized Minimum Residual method. To accelerate convergence, the Generalized Minimum Residual method is preconditioned. These and other problems associated with solving the full potential equation on an unstructured mesh are discussed. Results are presented for subsonic and transonic flows over bumps, airfoils, and wings to demonstrate the unstructured algorithm presented here. / Ph. D.

A numerical study of a proposed innovative jet interaction configuration is presented. This work aimed at improving present-day jet interaction configurations in their applications as control thrusters on hypersonic vehicles. Jet thrusters are a useful control system for fast-moving vehicles flying in the upper layers of the atmosphere because of their effectiveness and responsiveness. They produce a strong and responsive lateral force on the vehicle through the interaction of two main mechanisms. The first mechanism comes from the momentum of the injectant itself, basically the thrust of the jet. The second and subtler contribution comes from the jet interaction flowfield, the interaction of the expanding injectant with the crossflow. This interaction produces areas of high pressure ahead of the injector and areas of low pressure in the region aft of the jet. The combination of the high-pressure regions in front of and low-pressure regions aft of the injector produces an undesirable nose-down pitching moment on the vehicle. In order to counterbalance the nose-down attitude, modern-day thruster designs include a large secondary injector far aft of the center of gravity of the vehicle. The thrust of this second injector acting far aft of the primary injector neutralizes the nose-down pitching moment. This is not an efficient method to obviate the problem since it requires the vehicle to be designed to carry two large thrusters and double the quantity of fuel necessary for one thruster. In light of these considerations, this study aimed at developing a jet interaction configuration that can dispense from the need of a large secondary injector to compensate for the nose-down pitching moment. The cases studied here were first a primary jet alone and then a primary jet with pairs of smaller jets. This configuration was based on the notion that the interaction of the secondary jets, conveniently located immediately aft of the thruster, with the barrel shock and the wake of the primary jet can drastically reduce the nose-down pitching moment. Because of the complexity of the jet interaction flowfield the investigation of the feasibility and the assessment of the efficiency of the new jet interaction configurations combined the present numerical effort with experimental studies of jet interaction flowfields performed in the supersonic wind tunnel at Virginia Tech. During the present numerical study the jet interaction flowfield associated with the sonic injection of a gas into a high-speed crossflow was simulated by numerically solving the Reynolds Averaged Navier Stokes (RANS) equations. Turbulence was modeled through a first-order model, the Wilcox's 1988 k-w turbulence model. The computations made use of the finite volume code General Aerodynamic Simulation Program (GASP) Version 4. For simplicity and to keep the study general, the jet interaction flowfield was studied on a flat plate instead of a body of revolution as on a vehicle. Calculations were run for a number of jet interaction configurations consisting of a primary jet alone, a primary jet and one pair of secondary jets, and a primary jet and two pairs of secondary jets. The flow conditions of the simulations ranged from a Mach number of 2.1 up to a Mach number of 4.5 and jet total pressure to freestream static pressure ratios of 14 to 680. A large effort was dedicated to the development of an efficient computational grid that could capture most of the flow-physics with a minimum number of cells. To this end , Chimera or overset grids were employed in the simulation of the secondary injectors. Grid convergence was shown to be achieved for the case of single injection by conducting a thorough convergence study. The discretization error was calculated through a modified Richardson extrapolation to be low. The numerical solutions were compared to the experimental results in order to assess the capability of RANS equations and of first-order turbulence models to properly simulate the complex flowfield. The k-w turbulence model proved to be reliable and robust and the results it provided for this type of flowfield were accurate enough from an engineering standpoint to make informed decisions about the configuration layout. In spite of the overall good performance, the k-w turbulence model failed to correctly predict the flow in the regions of strong adverse pressure gradients. Comparisons with experimental results showed that the separation region was often under-predicted thus highlighting the need to employ better turbulence models for more accurate results. The RANS equations were found accurate enough to provide physical mean-flow solutions. Further, the numerical simulations provided information about the detailed physics of the flowfield that is impossible to obtain through experimental work. The analysis of the numerical solutions highlighted the existence of a complex system of counter-rotating trailing vortices that are responsible for the mixing of the injectant with the freestream. The typical features of the flowfield created by an under-expanded jet exhausting in a quiescent medium were visible in the jet interaction flowfield with the difference of the existence of a crossflow and a non-uniform back-pressure. The region of low pressure aft of the injector was shown to be generated by the detachment of the barrel shock from the surface of the flat plate that leaves a large volume to be filled by the surrounding fluid. The simulations showed that the innovative configuration with one primary jet and an array of smaller secondary jets can effectively decrease the nose-down pitching moment by as much as 160%. In some cases, it also increased the total normal force acting on the flat plate (namely the thrust) by as much as 3%. This effect was found to be caused by the reduction in size and intensity of the low-pressure region aft of the primary injector. / Ph. D.