The present work focuses on the system identification method of aerodynamic parameter estimation which is used to calculate the stability and control derivatives required for aircraft flight mechanics. A new rotorcraft parameter estimation technique is proposed which uses a type of artificial neural network (ANN) called radial basis function network (RBFN). Rotorcraft parameter estimation using ANN is an unexplored research topic and the earlier works in this area have used the output error, equation error and filter error methods which are conventional parameter estimation methods. However, the conventional methods require an accurate non-linear rotorcraft simulation model which is not required by the ANN based method. The application of RBFN overcomes the drawbacks of multilayer perceptron (MLP) based delta method of parameter estimation and gives satisfactory results at either end of the ordered set of estimates. This makes the RBFN based delta method for parameter estimation suitable for rotorcraft studies, as both transition and high speed flight regime characteristics can be studied. The RBFN based delta method for parameter estimation is used for computation of aerodynamic parameters from both simulated and real time flight data. The simulated data is generated from an 8-DoF non-linear simulation model based on the Level-1 criteria of rotorcraft simulation modeling. The generated simulated data is used for computation of the quasi-steady and the time-variant stability and control parameters for different flight conditions using the RBFN based delta method. The performance of RBFN based delta method is also analyzed in the presence of state and measurement noise as well as outliers. The established methodology is then applied to compute parameters directly from real time flight test data for a BO 105 S123 helicopter obtained from DLR (German Aerospace Center). The parameters identified using the RBFN based delta method are compared with the identified values for the BO 105 helicopter from published literature which have used conventional parameter estimation techniques for parameter estimation using a 6-DoF and a 9-DoF rotorcraft simulation model. Finally, the estimated parameters are verified from the flight data generated by a frequency sweep pilot control input for assessing the predictive capability of the RBFN based delta method. Since the approach directly computes the parameters from flight data, it can be used for a reliable description of the higher frequency range, which is needed for high bandwidth flight control and in-flight simulation.

Hypersonic vehicles are exposed to severe heating loads during their flight in the atmosphere. In order to minimize the heating problem, a variety of cooling techniques are presently available for hypersonic blunt bodies. Introduction of a forward-facing cavity in the nose tip of a blunt body configuration of hypersonic vehicle is one of the most simple and attractive methods of reducing the convective heating rates on such a vehicle. In addition to aerodynamic heating, the overall drag force experienced by vehicles flying at hypersonic speeds is predominate due to formation of strong shock waves in the flow. Hence, the effective management of heat transfer rate and aerodynamic drag is a primary element to the success of any hypersonic vehicle design. So, precise information on both aerodynamic forces and heat transfer rates are essential in deciding the performance of the vehicle. In order to address the issue of both forces and heat transfer rates, right kind of measurement techniques must be incorporated in the ground-based testing facilities for such type of body configurations. Impulse facilities are the only devices that can simulate high altitude flight conditions. Uncertainties in test flow conditions of impulse facilities are some of the critical issues that essentially affect the final experimental results. Hence, more reliable and carefully designed experimental techniques/methodologies are needed in impulse facilities for generating experimental data, especially at hypersonic Mach numbers. In view of the above, an experimental program has been initiated to develop novel techniques of measuring both the aerodynamic forces and surface heat transfer rates. In the present investigation, both aerodynamic forces and surface heat transfer rates are measured over the test models at hypersonic Mach numbers in IISc hypersonic shock tunnel HST-2, having an effective test time of 800 s. The aerodynamic coefficients are measured with a miniature type accelerometer based balance system where as platinum thin film sensors are used to measure the convective heat transfer rates over the surface of the test model. An internally mountable accelerometer based balance system (three and six-component) is used for the measurement of aerodynamic forces and moment coefficients acting on the different test models (i.e., blunt cone with after body, blunt cone with after body and frustum, blunt cone with after body-frustum-triangular fins and sharp cone with after body-frustum-triangular fins), flying at free stream Mach numbers of 5.75 and 8 in hypersonic shock tunnel. The main principle of this design is that the model along with the internally mounted accelerometer balance system are supported by rubber bushes and there-by ensuring unrestrained free floating conditions of the model in the test section during the flow duration. In order to get a better performance from the accelerometer balance system, the location of accelerometers plays a vital role during the initial design of the balance. Hence, axi-symmetric finite element modeling (FEM) of the integrated model-balance system for the missile shaped model has been carried out at 0° angle of attack in a flow Mach number of 8. The drag force of a model was determined using commercial package of MSC/NASTRAN and MSC/PATRAN. For test flow duration of 800 s, the neoprene type rubber with Young’s modulus of 3 MPa and material combinations (aluminum and stainless steel material used as the model and balance) were chosen. The simulated drag acceleration (finite element) from the drag accelerometer is compared with recorded acceleration-time history from the accelerometer during the shock tunnel testing. The agreement between the acceleration-time history from finite-element simulation and measured response from the accelerometer is very good within the test flow domain. In order to verify the performance of the balance, tests were carried out on similar standard AGARD model configurations (blunt cone with cylinder and blunt cone with cylinder-frustum) and the results indicated that the measured values match very well with the AGARD model data and theoretically estimated values. Modified Newtonian theory is used to calculate the aerodynamic force coefficient analytically for various angles of attack. Convective surface heat transfer rate measurements are carried out by using vacuum sputtered platinum thin film sensors deposited on ceramic substrate (Macor) inserts which in turn are embedded on the metallic missile shaped body. Investigations are carried out on a model with and without fin configurations in HST-2 at flow Mach number of 5.75 and 8 with a stagnation enthalpy of 2 MJ/kg for zero degree angle of attack. The measured heating rates for the missile shaped body (i.e., with fin configuration) are lower than the predicted stagnation heating rates (Fay-Riddell expression) and the maximum difference is about 8%. These differences may be due to the theoretical values of velocity gradient used in the empirical relation. The experimentally measured values are expressed in terms of normalized heat transfer rates, Stanton numbers and correlated Stanton numbers, compared with the numerically estimated results. From the results, it is inferred that the location of maximum heating occurs at stagnation point which corresponds to zero velocity gradient. The heat-transfer ratio (q1/Qo)remains same in the stagnation zone of the model when the Mach number is increased from 5.75 to 8. At the corners of the blunt cone, the heat transfer rate doesn’t increase (or) fluctuate and the effects are negligible at two different Mach numbers (5.75 and 8). On the basis of equivalent total enthalpy, the heat-transfer rate with fin configuration (i.e., at junction of cylinder and fins) is slightly higher than that of the missile model without fin. Attempts have also been made to evaluate the feasibility of using forward facing cavity as probable technique to reduce the heat transfer rate and to study its effect on aerodynamic coefficients on a 41° apex angle missile shaped body, in hypersonic shock tunnel at a free stream Mach number of 8. The forward-facing circular cavities with two different diameters of 6 and 12 mm are chosen for the present investigations. Experiments are carried out at zero degree angle of attack for heat transfer measurements. About 10-25 % reduction in heat transfer rates is observed with cavity at gauge locations close to stagnation region, whereas the reduction in surface heat transfer rate is between 10-15 % for all other gauge locations (which is slightly downstream of the cavity) compared with the model without cavity. In order to understand the influence of forward facing cavities on force coefficients, measurement of aerodynamic forces and moment coefficients are also carried out on a missile shaped body at angles of attack. The same six component balance is also being used for subsequent investigation of force measurement on a missile shaped body with forward facing cavity. Overall drag reductions of up to 5 % is obtained for a cavity of 6 mm diameter, where as, for the 12 mm cavity an increase in aerodynamic drag is observed (up to about 10%). The addition of cavity resulted in a slight increase in the missile L/D ratio and did not significantly affect the missile lateral components. In summary, the designed balances are found to be suitable for force measurements on different test models in flows of duration less than a millisecond. In order to compliment the experimental results, axi-symmetric, Navier-Stokes CFD computations for the above-defined models are carried out for various angles of attack using a commercial package CFX-Ansys 5.7. The experimental free stream conditions obtained from the shock tunnel are used for the boundary conditions in the CFD simulation. The fundamental aerodynamic coefficients and heat transfer rates of experimental results are shown to be in good agreement with the predicted CFD. In order to have a feeling of the shock structure over test models, flow visualization experiments have been carried out by using the Schlieren technique at flow Mach numbers of 5.75 and 8. The visualized shock wave pattern around the test model consists of a strong bow shock which is spherical in shape and symmetrical over the forebody of the cone. Experimentally measured shock stand-off distance compare well with the computed value as well as the theoretically estimated value using Van Dyke’s theory. These flow visualization experiments have given a factual proof to the quality of flow in the tunnel test section.

Cancer is the second leading cause of death in the United States and its onset is highly incident in the lungs, with very low long-term survival rates. Chemotherapy plays a significant role for lung cancer treatment, and pulmonary delivery may be a potential route for anticancer drug delivery to treat lung tumors. Coenzyme Q₁₀ (CoQ₁₀) is a poorly-water soluble compound that is being investigated for the treatment of carcinomas. In this work, we hypothesize that formulations of CoQ10 may be developed for pulmonary delivery with a satisfactory pharmacokinetic profile that will have the potential to improve a pharmacodynamic response when treating lung malignancies. The formulation design was to use a vibrating-mesh nebulizer to aerosolize aqueous dispersions of CoQ₁₀ stabilized by phospholipids physiologically found in the lungs. In the first study, a method was developed to measure the surface tension of liquids, a physicochemical property that has been shown to influence the aerosol output characteristics from vibrating-mesh nebulizers. Subsequently, this method was used, together with analysis of particle size distribution, zeta potential, and rheology, to further evaluate the factors influencing the capability of this nebulizer system to continuously and steadily aerosolize formulations of CoQ₁₀ prepared with high pressure homogenization. The aerosolization profile (nebulization performance and in vitro drug deposition of nebulized droplets) of formulations prepared with soybean lecithin, dimyristoylphosphatidylcholine (DMPC), dipalmitoylphosphatidylcholine (DPPC) and distearoylphosphatidylcholine (DSPC) were evaluated. The rheological behavior of these dispersions was found to be the factor that may be indicative of the aerosolization output profile. Finally, the pulmonary deposition and systemic distribution of CoQ₁₀ prepared as DMPC, DPPC, and DSPC dispersions were investigated in vivo in mice. It was found that high drug amounts were deposited and retained in the mouse lungs for at least 48 hours post nebulization. Systemic distribution was not observed and deposition in the nasal cavity occurred at a lower scale than in the lungs. This body of work provides evidence that CoQ₁₀ may be successfully formulated as dispersions to be aerosolized using vibrating-mesh nebulizers and achieve high drug deposition in the lungs during inhalation. / text

Αντικείμενο της εργασίας είναι η αεροδυναμική και αεροακουστική ανάλυση στροφείων ανεμοκινητήρων οριζοντίου άξονα (α-ο-α). Ο υπολογισμός του πεδίου ροής και των αεροδυναμικών συντελεστών του στροφείου ενός ανεμοκινητήρα επιτυγχάνεται κατά δύο τρόπους, με σκοπό την άμεση σύγκριση των αποτελεσμάτων με κριτήρια αφενός την ακρίβεια και αφετέρου την ευκολία ή πρακτικότητα που προσδιορίζεται κύρια σε όρους χρόνου υπολογισμού και διαθεσιμότητας υπολογιστικών πόρων. Οι δύο επιλεγμένοι τρόποι που διαφοροποιούνται στην φυσικο-μαθηματική μοντελοποίηση του προβλήματος ροής γύρω από το στροφείο του ανεμοκινητήρα, αποτελούν δύο δοκιμασμένες μεθοδολογίες ή τεχνικές ανάλυσης και σχεδιασμού περιστρεφόμενων στροφείων, τα οποία μπορούν να λειτουργούν ως κινητήριες μηχανές ή ως εργομηχανές, είναι η μέθοδος των επιφανειακών στοιχείων και η αριθμητική επίλυση των εξισώσεων Navier-Stokes. Για την αξιολόγηση των υπολογιστικών αποτελεσμάτων επιλέχθηκε ως στροφείο αναφοράς, ο πειραματικός ανεμοκινητήρας NREL phase II. Ο αλγόριθμος των επιφανειακών στοιχείων συμπλέχτηκε με ολοκληρωτικά σχήματα πρόλεξης και υπολογισμού του οριακού στρώματος με σκοπό να συμπεριληφθούν τα φαινόμενα συνεκτικότητας της ροής. Πραγματοποιήθηκε παραμετρική ανάλυση του δρομέα του ανεμοκινητήρα για διαφορετικές συνθήκες λειτουργίας του. Η σύγκριση των αποτελεσμάτων των συντελεστών πίεσης των περιστρεφόμενων πτερυγίων για τέσσερις θέσεις κατά το εκπέτασμα του πτερυγίου με τα πειραματικά δεδομένα δείχνει ικανοποιητική συμφωνία. Για την ανάλυση του πεδίου ροής που παράγεται γύρω από περιστρεφόμενους δρομείς α-ο-α χρησιμοποιήθηκε η μέθοδος της υπολογιστικής ρευστοδυναμικής (CFD). Πραγματοποιήθηκαν RANS προσομοιώσεις για διαφορετικές συνθήκες λειτουργίας του ανεμοκινητήρα και για τέσσερα διαφορετικά μοντέλα τύρβης. Το k-ω SST μοντέλο τύρβης έχει τις μικρότερες αποκλίσεις με τα πειραματικά αποτελέσματα. Η αεροακουστική ανάλυση του στροφείου ενός ανεμοκινητήρα επιτυγχάνεται με την επίλυση της ακουστικής εξίσωσης Ffowcs-Williams Hawkings, μέσω ενός υπολογιστικού κώδικα που αναπτύχθηκε γι’ αυτό το σκοπό. Από τα αποτελέσματα των προσομοιώσεων, φάνηκε στα ροδογράμματα κατευθυντικότητας του ήχου, τα επίπεδα της ακουστικής πίεσης να είναι υψηλότερα για θέσεις παρατηρητή ανάντη και κατάντη του ανεμοκινητήρα. / The aim of this study is to represent the aerodynamic and aeroacoustic analysis of horizontal axis wind turbine (ΗAWT) rotors. The calculation of the flow field and the aerodynamic coefficients over the wind turbine rotor are performed using two methodologies, the panel method and the numerical solution of Navier-Stokes equations. These two methodologies are differentiated in the mathematical modeling approach of the flow around the rotor and are utilized in the design and manufacturing phases of horizontal axis wind turbine rotors. Moreover, the results of these two methodologies are compared in terms of the accuracy and the computational time required. For the evaluation of the computational results the experimental wind turbine NREL phase II is chosen as the reference rotor. An invicid/viscous interaction algorithm is developed using integral boundary layer equations coupled with the low order panel method solution in order to account the viscous effects. A parametric analysis of the wind turbine rotor is conducted for different operating conditions. The comparison of the results of the pressure coefficients of the rotating blades for four spanwise positions along the blade with the experimental data shows satisfactory agreement. The analysis of the near and far flow field of HAWT is obtained via CFD by RANS simulations of four different turbulence models (Spalart-Allmaras, k-ε, k-ε RNG and k-ω SST). From the conducted study, it is confirmed the ability of analysis of a HAWT rotor flow field with the RANS equations and the good agreement of the computations with experimental data, when the k-ω SST turbulence model is used. The aeroacoustic analysis of the HAWT is based on the solution of the Ffowcs Williams-Hawkings (FW-H) equation via a computer code developed for this purpose. The radiation patterns of the calculated aeroacoustic noise show that high level amplitudes are calculated for upwind and downwind positions.

Current lean-premixed fuel injector designs have shown great potential in terms of reducing emissions of pollutants, but such designs are susceptible to combustion instabilities in which aerodynamic instability plays a major role and also has an effect on mixing of air and fuel. In comparison to prototype testing with combustors running in operating conditions, computational approaches such as Large Eddy Simulations (LES) offer a much more cost-effective alternative in the design stage. However, computational models employed by LES require validation by experimental data. This is one of the main motivations behind the present experimental study. Combined particle image velocimetry (PIV) and planar laser induced fluorescence (PLIF) instrumentation allowed simultaneous measurements of velocity vector and a conserved scalar introduced into the fuel stream. The results show that the inner swirl shear layer features two pairs of vortices, which draw high concentration fuel mixture from the central jet into the swirl stream and causes it to rotate in their wakes. Such periodic entrainment also occurs with the characteristic frequencies of the vortices. This has clear implications for temporal variations in fuel/air ratio in a combusting flow; these bursts of mixing, and hence heat release, could be a possible cause of mixing-induced pressure oscillation in combusting tests. For the first time in such a flow, all 3 components of the turbulent scalar flux were available for validation of LES-based predictions. A careful assessment of experimental errors, particularly the error associated with spatial filtering, was carried out. Comparison of LES predictions with experimental data showed very good agreement for both 1st and 2nd moment statistics, as well as spectra and scalar pdfs. It is particularly noteworthy that comparison between LES computed and measured scalar fluxes was very good; this represents successful validation of the simple (constant Schmidt number) SGS model used for this complex and practically important fuel injector flow. In addition to providing benchmark data for the validation of LES predictions, a new experimental technique has been developed that is capable of providing spatially resolved residence time data. Residence times of combustors have commonly been used to help understand NOx emissions and can also contribute to combustion instabilities. Both the time mean velocity and turbulence fields are important to the residence time, but determining the residence time via analysis of a measured velocity field is difficult due to the inherent unsteadiness and the three dimensional nature of a high-Re swirling flow. A more direct approach to measure residence time is reported here that examines the dynamic response of fuel concentration to a sudden cutoff in the fuel injection. Residence time measurement was mainly taken using a time-resolved PLIF technique, but a second camera for PIV was added to check that the step change does not alter the velocity field and the spectral content of the coherent structures. Characteristic timescales evaluated from the measurements are referred to as convection and half-life times: The former describes the time delay from a fuel injector exit reference point to a downstream point of interest, and the latter describes the rate of decay once the effect of the reduced scalar concentration at the injection source has been transported to the point of interest. Residence time is often defined as the time taken for a conserved scalar to reduce to half its initial value after injection is stopped: this is equivalent to the sum of the convection time and the half-life values. The technique was applied to a high-swirl fuel injector typical of that found in combustor applications. Two test cases have been studied: with central jet (with-jet) and without central jet (no-jet). It was found that the relatively unstable central recirculation zone of the no-jet case resulted in increased transport of fuel into the central region that is dominated by a precessing vortex core, where long half-life times are also found. Based on this, it was inferred that the no-jet case may be more prone to NOx production. The technique is described here for a single-phase isothermal flow field, but with consideration, it could be extended to studying reacting flows to provide more insight into important mixing phenomena and relevant timescales.

Today's globally competitive market and its associated environmental and social issues of sustainable development are major challenges for the automobile industry. To meet them, the industry needs to invest in high performance development tools. For improving engine performance in terms of consumption and emission, the interactions between the subsystems of the engine air-path need to be understood. This thesis followed two major axes of research in this context. First, the problems related to the modeling of the global air-path system were studied, which include the airflow characteristics between the different subsystems of the air-path, high frequency combustion modeling and pulsating airflow, and estimation of the exhaust aerodynamic force on the vanes of variable geometry turbochargers (VGT). The detailed modeling study was used for developing an engine air-path simulator, which takes into account these interactions and predicts the influence of subsystems on the global air-path. The second axis of research was focused on modeling of mechatronic actuators of the air-path, taking into account their nonlinear behavior due to friction and changes in operating conditions. A generic nonlinear dynamic model was developed and included in the simulator. This model can be adapted to most commercial actuators. The complete simulator has been implemented using AMESim for engine and air-path modeling, and Simulink for control. It has been parameterized according to the specifications of a commercial diesel engine and validated against experimental data. Finally, robust local controllers were studied for actuator position control, aimed at guaranteeing the performance of the actuators under parametric uncertainty and external disturbances. An advanced controller was developed, which adapts to changes in friction characteristics of the actuator and external load changes. The performance of all controllers has been demonstrated experimentally.

Cancer is the second leading cause of death in the United States and its onset is highly incident in the lungs, with very low long-term survival rates. Chemotherapy plays a significant role for lung cancer treatment, and pulmonary delivery may be a potential route for anticancer drug delivery to treat lung tumors. Coenzyme Q₁₀ (CoQ₁₀) is a poorly-water soluble compound that is being investigated for the treatment of carcinomas. In this work, we hypothesize that formulations of CoQ10 may be developed for pulmonary delivery with a satisfactory pharmacokinetic profile that will have the potential to improve a pharmacodynamic response when treating lung malignancies. The formulation design was to use a vibrating-mesh nebulizer to aerosolize aqueous dispersions of CoQ₁₀ stabilized by phospholipids physiologically found in the lungs. In the first study, a method was developed to measure the surface tension of liquids, a physicochemical property that has been shown to influence the aerosol output characteristics from vibrating-mesh nebulizers. Subsequently, this method was used, together with analysis of particle size distribution, zeta potential, and rheology, to further evaluate the factors influencing the capability of this nebulizer system to continuously and steadily aerosolize formulations of CoQ₁₀ prepared with high pressure homogenization. The aerosolization profile (nebulization performance and in vitro drug deposition of nebulized droplets) of formulations prepared with soybean lecithin, dimyristoylphosphatidylcholine (DMPC), dipalmitoylphosphatidylcholine (DPPC) and distearoylphosphatidylcholine (DSPC) were evaluated. The rheological behavior of these dispersions was found to be the factor that may be indicative of the aerosolization output profile. Finally, the pulmonary deposition and systemic distribution of CoQ₁₀ prepared as DMPC, DPPC, and DSPC dispersions were investigated in vivo in mice. It was found that high drug amounts were deposited and retained in the mouse lungs for at least 48 hours post nebulization. Systemic distribution was not observed and deposition in the nasal cavity occurred at a lower scale than in the lungs. This body of work provides evidence that CoQ₁₀ may be successfully formulated as dispersions to be aerosolized using vibrating-mesh nebulizers and achieve high drug deposition in the lungs during inhalation.

There is currently a great deal of interest within the aviation community in the design of small, slow-flying but manoeuvrable uninhabited vehicles for reconnaissance, surveillance, and search and rescue operations in urban environments. Inspired by observation of birds, insects, fish and cetaceans, flapping wings are being actively studied in the hope that they may provide greater propulsive efficiencies than propellers and rotors at low Reynolds numbers for such Micro-Air Vehicles (MAVs). Researchers have posited the Strouhal number (combining flapping frequency, amplitude and forward speed) as the parameter controlling flapping wing aerodynamics in cruising flight, although there is conflicting evidence. This thesis explores the effect of flapping frequency and amplitude on forces and wake structures, as well as physical mechanisms leading to optimum propulsive efficiency. Two-dimensional rigid airfoils are considered at Reynolds number 2,000 ??? 40,000. A compressible Navier-Stokes simulation is combined with numerical and analytical potential flow techniques to isolate and evaluate the effect of viscosity, leading and trailing edge vortex separation, and wake vortex dynamics. The wake structures of a plunging airfoil are shown to be sensitive to the flapping frequency independent of the Strouhal number. For a given frequency, the wake of the airfoil exhibits ???vortex lock-in??? as the amplitude of motion is increased, in a manner analogous to an oscillating circular cylinder. This is caused by interaction between the flapping frequency and the ???bluff-body??? vortex shedding frequency apparent even for streamlined airfoils at low Reynolds number. The thrust and propulsive efficiency of a plunging airfoil are also shown to be sensitive to the flapping frequency independent of Strouhal number. This dependence is the result of vortex shedding from the leading edge, and an interaction between the flapping frequency and the time for vortex formation, separation and convection over the airfoil surface. The observed propulsive efficiency peak for a pitching and plunging airfoil is shown to be the result of leading edge vortex shedding at low flapping frequencies (low Strouhal numbers), and high power requirements at large flapping amplitudes (high Strouhal numbers). The efficiency peak is governed by flapping frequency and amplitude separately, rather than the Strouhal number directly.

This work presents the coupling between two different formulations applied to flow simulation and analysis of wind rotors, integral and differential formulations. First, for the integral formulation is defined a control volume where the variables problem are defined, as well as the necessaries working hypothesis, then a proposed mathematical modeling is defined. Simulations through NACA airfoils, using Computational Dynamic Fluids, are performed in order to evaluate drag and lift coefficients, to be used in the integral formulation. The Navier-Stokes equations are solved in house and the Smagorinsky turbulence model with Van Driest damping function is retained. The computational code is implemented with structured cartesian mesh, where the airfoil is modeled using the Immersed Boundary Methodology. The results of simulation through a NACA0012 airfoil are shown for several attack angles and Re = 10000. Results of energetic efficiency are presented for a horizontal axis wind turbine using the integral formulation, where the coefficients are given by differential formulations. / Este trabalho apresenta o acoplamento entre as duas formulações diferentes aplicadas à simulação do escoamento e análise de rotores eólicos, formulações integral e diferencial. Primeiramente para a formulação integral é definido um volume de controle onde as variáveis do problema são definidas, bem como as hipóteses simplificadoras necessárias, para então ser proposta uma modelagem matemática. Simulações do escoamento em aerofólios NACA, utilizando Dinâmica dos Fluidos Computacional, são realizadas a fim de determinar os coeficientes de arrasto e sustentação, os quais foram utilizados na formulação integral. As equações de Navier-Stokes são resolvidas em um código computacional e o modelo de turbulência de Smagorinsky com função de amortecimento de Van Driest é utilizado. O código computacional é implementado com uma malha cartesiana estruturada, e o aerofólio é modelado utilizando a Metodologia da Fronteira Imersa. Os resultados da simulação através de um aerofólio NACA0012 são mostrados para vários ângulos de ataque e Re = 10000. Os resultados de eficiência energética são apresentados para uma turbina eólica de eixo horizontal utilizando a formulação integral, onde os coeficientes são dados pelas formulações diferenciais. / Mestre em Engenharia Mecânica

Les chantiers de maintenance et d’assainissement dans les centrales nucléaires nécessitent la mise en place d’enceintes ventilées autour des zones contaminées afin de limiter la propagation de la contamination à l’environnement extérieur. L’air rentre dans l’enceinte aux travers d’ouvertures sous la forme d’un écoulement directionnel, orienté de l’extérieur vers l’intérieur, assurant ainsi le confinement dynamique. En raison des opérations qui se déroulent à l’intérieur de l’enceinte et des perturbations externes, il est possible que l’écoulement de confinement dynamique aux ouvertures soit perturbé et subisse des inversions locales et instationnaires, conduisant ainsi à transporter la contamination à l’extérieur de l’enceinte. La présente étude s’intéresse aux petites ouvertures de type fentes minces rectangulaires où l’écoulement au droit de celles-ci est généralement turbulent. Les principaux objectifs de la thèse sont d’une part d’identifier les conditions aérodynamiques susceptibles de produire le phénomène de rétrodiffusion aux ouvertures, d’autre part d’évaluer la capacité des approches de modélisation de la turbulence URANS et LES à reproduire les instabilités liées à ce type d’écoulement. Il a été montré que l’apparition du phénomène de rétrodiffusion est principalement liée à la présence d’une perturbation aéraulique additionnelle, de type jet turbulent ou sillage, en compétition avec l’écoulement initial de confinement dynamique. Des expériences de traçage gazeux ont été mises en place sur une maquette expérimentale dans le but de quantifier la rétrodiffusion en fonction des différentes conditions aérauliques à l’ouverture et des caractéristiques de celle-ci. Des visualisations des écoulements à l’ouverture ont également été réalisées à l’aide d’un dispositif de tomographie laser. Enfin, l’analyse des résultats des simulations CFD a démontré que les approches de type RANS ou URANS ne permettaient pas de reproduire les instabilités de l’écoulement conduisant au phénomène de rétrodiffusion, contrairement aux simulations des grandes échelles de la turbulence (LES) qui reproduisent fidèlement les structures locales et instantanées à l’origine du phénomène. / Operations of decommissioning and decontamination in nuclear facilities require setting up depressurized enclosures around contaminated areas in order to prevent leakage of radioactive materials, to the surrounding environment. Air passes through openings which generates a directional airflow ensuring the aerodynamic containment of hazardous material inside the enclosure. Due to operating activities inside or outside the enclosure, the directional flow might be disturbed. Consequently, local and unsteady backflows may occur at the opening leading to the outward transport of contamination. The current study is focused on airflow dynamics through small openings, such as rectangular slits where the initial inflow stream is turbulent. The main purposes of this work are to identify the required aerodynamic conditions likely to generate unsteady flow inversions at the studied opening and also to verify the ability of CFD simulations to predict this type of flow by using URANS and LES approaches. Results have shown that an additional flow, such as a turbulent jet or a wake in competition with inward flow, is the main cause leading to the leakage at the opening. Experiments, using gas tracer detection techniques, are conducted in order to quantify outflow leakage in the near field of the opening under different aerodynamic configurations and openings characteristics. A laser tomography technique is also implemented to visualize the external leakage airflow in the middle plane of the opening. CFD simulations have shown that a qualitative description of instantaneous leakage flow patterns at the opening can be achieved. This is characterized by the occurrence of local coherent structures transporting passive tracer outwards. Moreover, velocities obtained from CFD results (Large Eddy Simulations) are compared to those obtained from experimental measurements.