A novel parallel block-based anisotropic adaptive mesh refinement (AMR) technique for multi-block body-fitted grids is proposed and described. Rather than adopting the more usual isotropic approach to mesh refinement, an anisotropic refinement procedure is proposed which allows refinement of grid blocks in each coordinate direction in an independent fashion. This allows for more efficient and accurate treatment of narrow layers and/or discontinuities which occur, for example, in the boundary and mixing layers of viscous flows, and in regions of strong non-linear wave interactions with shocks. The anisotropic AMR technique is implemented within an existing finite-volume framework, which encompasses both explicit and implicit solution methods, and is capable of performing calculations with second- and higher-order spatial accuracy. To clearly demonstrate the feasibility of the proposed technique, it is applied to the unsteady and steady-state solutions of both the advection diffusion equation, as well as the Euler equations, in two space dimensions.

A novel parallel block-based anisotropic adaptive mesh refinement (AMR) technique for multi-block body-fitted grids is proposed and described. Rather than adopting the more usual isotropic approach to mesh refinement, an anisotropic refinement procedure is proposed which allows refinement of grid blocks in each coordinate direction in an independent fashion. This allows for more efficient and accurate treatment of narrow layers and/or discontinuities which occur, for example, in the boundary and mixing layers of viscous flows, and in regions of strong non-linear wave interactions with shocks. The anisotropic AMR technique is implemented within an existing finite-volume framework, which encompasses both explicit and implicit solution methods, and is capable of performing calculations with second- and higher-order spatial accuracy. To clearly demonstrate the feasibility of the proposed technique, it is applied to the unsteady and steady-state solutions of both the advection diffusion equation, as well as the Euler equations, in two space dimensions.

O presente trabalho tem por objetivo o estudo da otimização multiobjetivo aplicada ao projeto de perfis aerodinâmicos em regime transônico, analisando comparativamente diferentes formas de definir as funções objetivo. A otimização é efetuada pelo algoritmo genético NSGA-II. Os resultados são avaliados utilizando métricas de diversidade da população e otimalidade das soluções, das quais duas são propostas. As funções objetivo são constituidas de diferentes parametrizações da geometria e diferentes técnicas de simulação numérica. A parametrização da geometria é feita utilizando a paramentrização Parsec ou a parametrização baseada em pontos de controle. A discretização do domínio espacial é feita utilizando malha estruturada conformada ao perfil e suavização por EDP elíptica. As duas técnicas de volumes finitos com diferentes modelos para o cálculo do fluxo na face do volume implementadas foram o método de Jameson (esquema centrado) e o método de Roe (esquema upwind). As comparações feitas são as seguintes: utilização de modelo viscoso e invíscido, com o uso do código Mses com a parametrização por ponto de controle; a utilização da parametrização por ponto de controle e parametrização Parsec usando o método de Jameson; e a comparação entre o método centrado e o upwind, utilizando a parametrização Parsec. Conclui-se dos resultados obtidos que a utilização da parametrização por pontos de controle é melhor. Entretanto, ainda é necessária a utilização de uma parametrização que garanta maior suavidade ou a imposição de restrições sobre a suavidade da solução. A utilização do modelo viscoso torna os resultados da otimização melhores do ponto de vista da otimalidade. Na utilização de modelos de correção viscosa, como no caso do Mses, é necessária a utilização de métodos invíscidos que forneçam resultados com maior representatividade física / The objective of present study is analyze the multi-objective optimization applied to transonic airfoils project comparing different ways to define the objective functions. The optimization is evaluated by the genetic algorithm NSGA-II. The results is analyzed using metrics of diversity and optimality for multi-objective problems, which two are proposed. The objective functions are defined by different parametrizations of geometry and different techniques of numerical simulation. The geometry parametrization was made by two distinct forms: using Parsec parametrization; and the control points based parametrization. The space domain discretization was made using structured body-fitted mesh with elliptical PDE smooth. A finite volume code with two different techniques for calculations of flux interface had been implemented: the Jamesons method (centered); and the Roes method ( it upwind). For viscous model usage analysis was used the Mses code that has implemented a finite volumes technique with viscous model correction. The following comparisons has been made: viscous and inviscid model using the Mses code with the control points parametrization; the control points and Parsec parametrizations using the Jamesons method; and the comparison among the centered method and upwind using the parametrization Parsec. From the results, it is concluded that the used of control points parametrization is interesting. Although, is still needed the used of a parametrization that guarantees a better smoothness or the imposes of a geometrical or property distribution restriction. The uses of viscous model gives better optimizations results in optimality requirement. It is needed the uses of inviscid method that forces better physical representation when using viscous correction model

Dans cette thèse, nous nous intéressons aux équations des gaz sans pression avec contrainte de congestion qui soulèvent encore de nombreuses questions. La stratégie que nous proposons repose sur des précédents travaux sur le mouvement de foule dans le cadre de l'espace de Wasserstein, et sur un modèle granulaire avec des collisions inélastiques.Elle consiste en l'étude d'un schéma discrétisé en temps dont les suites doivent approcher les solutions de ces équations.Le schéma se présente de la manière suivante : à chaque pas de temps, le champ des vitesses est projeté sur un ensemble lui permettant d'éviter les croisements entre particules, la densité est ensuite déplacée selon le nouveau champ des vitesses, puis est projetée sur l'ensemble des densités admissibles (inférieures à une valeur seuil donnée).Enfin, le champ des vitesses est mis à jour en tenant compte du parcours effectué par les particules. En dimension 1, les solutions calculées par le schéma coïncident avec les solutions connues pour ce système. En dimension 2, les solutions calculées respectent les propriétés connues des solutions des équations de gaz sans pression avec contrainte de congestion. De plus, on retrouve des similarités entres ces solutions et celles du modèle granulaire microscopique dans des cas où elles sont comparables. Par la suite, la discrétisation en espace pose des problèmes et a nécessité l'élaboration d'un nouveau schéma de discrétisation du coût Wasserstein quadratique. Cette méthode que nous avons baptisée méthode du balayage transverse consiste à calculer le coût en utilisant les flux de masses provenant d'une certaine cellule et traversant les hyperplans définis par les interfaces entre les cellules. / In this thesis, we consider the pressureless Euler equations with a congestion constraint.This system still raises many open questions and aside from its one-dimensional version,very little is known. The strategy that we propose relies on previous works of crowd motion models withcongestion in the framework of the Wasserstein space, and on a microscopic granularmodel with inelastic collisions. It consists of the study of a time-splitting scheme. The first step is about the projection of the current velocity field on a set, avoiding the factthat trajectories do not cross during the time step. Then the scheme moves the density with the new velocity field. This intermediate density may violate the congestion constraint. The third step projects it on the set of admissible densities. Finally, the velocity field is updated taking into account the positions of physical particles during the scheme. In the one-dimensional case, solutions computed by the algorithm matchwith the ones that we know for these equations. In the two-dimensional case, computed solutions respect some properties that can be expected to be verified by the solutions to these equations. In addition, we notice some similarities between solutions computed by the scheme and the ones of the granular model with inelastic collisions. Later, this scheme is discretized with respect to the space variable in the purpose of numerical computations of solutions. The resulting algorithm uses a new method to discretize the Wasserstein cost. This method, called Transverse Sweeping Method consists in expressing the cost using the mass flow from any cell and crossing hyperplanes defined by interfaces between cells.

Parciální diferenciální rovnice hrají důležitou v inženýrských aplikacích. Často je možné tyto rovnice řešit pouze přibližně, tj. numericky. Z toho důvodu vzniklo množství diskretizačních metod pro řešení těchto rovnic. Uvedená nespojitá Galerkinova metoda se zdá jako velmi obecná metoda pro řešení těchto rovnic, především pak pro hyperbolické systémy. Naším cílem je řešit úlohy aeroakustiky, přičemž šíření akustických vln je popsáno pomocí linearizovaných Eulerových rovnic. A jelikož se jedná o hyperbolický systém, byla vybrána právě nespojitá Galerkinova metoda. Mezi nejdůležitější aspekty této metody patří schopnost pracovat s geometricky složitými oblastmi, možnost dosáhnout metody vysokého řádu a dále lokální charakter toho schématu umožnuje efektivní paralelizaci výpočtu. Nejprve uvedeme nespojitou Galerkinovu metodu v obecném pojetí pro jedno- a dvoudimenzionalní úlohy. Algoritmus následně otestujeme pro řešení rovnice advekce, která byla zvolena jako modelový případ hyperbolické rovnice. Metoda nakonec bude testována na řadě verifikačních úloh, které byly formulovány pro testování metod pro výpočetní aeroakustiku, včetně oveření okrajových podmínek, které, stejně jako v případě teorie proudění tekutin, jsou nedílnou součástí výpočetní aeroakustiky.

Indiana University-Purdue University Indianapolis (IUPUI) / Testing of a wave-rotor constant-volume combustor (WRCVC) showed the viability of the application of wave rotors as a pressure gain combustor. The aero-thermal design of the WRCVC rig had originally been performed with a time-dependent, one-dimensional model which applies a single-step reaction model for the combustion process of the air-fuel mixture. That numerical model was validated with experimental data with respect of matching the flame propagation speed and the pressure traces inside the passages of the WRCVC. However, the numerical model utilized a single progress variable representing the air-fuel mixture, which assumes that fuel and air are perfectly mixed with a uniform concentration; thus, limiting the validity of the model. In the present work, a two-step reaction model is implemented in the combustion model with four species variables: fuel, oxidant, intermediate and product. This combustion model is developed for a more detailed representation for the combustion process inside the wave rotor. A two-step reaction model presented a more realistic representation for the stratified air-fuel mixture charges in the WRCVC; additionally it shows more realistic modeling for the partial combustion process for rich fuel-air mixtures. The combustion model also accounts for flammability limits to exert flame extinction for non-flammable mixtures. The combustion model applies the eddy-breakup model where the reaction rate is influenced by the turbulence time scale. The experimental data currently available from the initial testing of the WRCVC rig is utilized to calibrate the model to determine the parameters, which are not directly measured and no directly related practice available in the literature. A prediction of the apparent ignition the location inside the passage is estimated by examination of measurements from the on-rotor instrumentations. The incorporation of circumferential leakage (passage-to-passage), and stand-off ignition models in the numerical model, contributed towards a better match between predictions and experimental data. The thesis also includes a comprehensive discussion of the governing equations used in the numerical model. The predictions from the two-step reaction model are validated using experimental data from the WRCVC for deflagrative combustion tests. The predictions matched the experimental data well. The predicted pressure traces are compared with the experimentally measured pressures in the passages. The flame propagation along the passage is also evaluated with ion probes data and the predicted reaction zone.

The main purpose of this work is to prove short-time structural stability of compressible vortex sheets with surface tension. The main result can be summarised as follows. Assume we start with an initial vortex-sheet configuration which consists of two inviscid fluids with density bounded below flowing smoothly past each other, where a strictly positive fixed coefficient of surface tension produces a surface tension force across the common interface, balanced by the pressure jump. We assume the fluids are modelled by the compressible Euler equations in three space dimensions with a very general equation of state relating the pressure, entropy and density in each fluid such that the sound speed is positive. Then, for a short time, which may depend on the initial configuration, there exists a unique solution of the equations with the same structure, that is, two fluids with density bounded below flowing smoothly past each other, where the surface tension force across the common interface balances the pressure jump. The mathematical approach consists of introducing a carefully chosen artificial viscosity-type regularisation which allows one to linearise the system so as to obtain a collection of transport equations for the entropy, pressure and curl together with a parabolic-type equation for the velocity. We prove a high order energy estimate for the non-linear equations that is independent of the artificial viscosity parameter which allows us to send it to zero. This approach loosely follows that introduced by Shkoller et al in the setting of a compressible liquid-vacuum interface. Although already considered by Shkoller et al, we also make some brief comments on the case of a compressible liquid-vacuum interface, which is obtained from the vortex sheets problem by replacing one of the fluids by vacuum, where it is possible to obtain a structural stability result even without surface tension.

This work is devoted to the numerical solution of equations describing the dynamics of flocks of birds. Specifically, we pay attention to the Euler equations for compressible flow with a right-hand side correction. This model is based on the work Fornasier et al. (2010). Due to the complexity of the model, we focus only on the one-dimensional case. For the numerical solution we use a semi-implicit discontinuous Galerkin method. Discretization of the right-hand side is chosen so that we preserve the structure of the semi-implicit scheme for the Euler equations presented in the work Feistauer, Kučera (2007). The proposed numerical scheme was implemented and numerical experiments showing the robustness of the scheme were carried out. Powered by TCPDF (www.tcpdf.org)

O estudo de métodos de modelagem de escoamentos aerodinâmicos em regime transônico é de grande importância para a engenharia aeronáutica. O maior desafio no tratamento desses escoamentos está na sua característica não linear devido aos efeitos de compressibilidade e formação de ondas de choque. Tais efeitos não lineares influenciam no desempenho de superfícies aerodinâmicas em geral, bem como são responsáveis pelo aparecimento de fenômenos danosos para a resposta aeroelástica de aeronaves. O equacionamento para esses tipos de escoamentos pode ser obtido via as equações básicas da mecânica dos fluidos. No entanto, apenas soluções numéricas de tais equações são possíveis de ser obtidas de forma prática no presente momento. Para o caso específico do tratamento de problemas transônicos, as equações de Euler formam um conjunto de equações diferenciais a derivadas parciais capazes de capturar os efeitos não lineares de escoamentos compressíveis, porém os efeitos da viscosidade não são levados em consideração. O objetivo desse trabalho é implementar uma rotina computacional capaz de resolver numericamente escoamentos em regime transônico em torno de aerofólios. Para isso as equações de Euler não lineares são utilizadas e o campo de fluido ao redor de um perfil aerodinâmico é discretizado pelo método das diferenças finitas. Uma malha estruturada do tipo C discretizando o fluido ao redor de um aerofólio NACA0012 é considerada. A metodologia para solução numérica é baseada no método explícito de MacCormack de segunda ordem de precisão no tempo e espaço. Baseados na aproximação upwind, termos de dissipação artificial com coeficientes não lineares também são adicionados ao método. A solução do escoamento transônico estacionário ao redor do aerofólio NACA0012 é obtida e as principais propriedades do escoamento são apresentadas. Observa-se a formação de ondas de choque através de contornos de número de Mach ao redor do aerofólio. Gráficos das distribuições de pressão no intra e extradorso do aerofólio são mostrados, onde se identificam aos efeitos da brusca variação de pressão devido as ondas de choque. Os resultados são validados com valores de distribuição de pressão para o mesmo aerofólio encontradas na literatura técnica. Os resultados obtidos combinam bem com os fornecidos em códigos computacionais para solução do mesmo problema aerodinâmico / The study of aerodynamic modeling methods for the transonic flow regime is of great importance in aeronautical engineering. Major challenge on the treatment of those flows is on their nonlinear features due to compressibility effects and shock waves (appearance). Such nonlinear effects present a strong influence on aerodynamic performance, as well as they are responsible for harmful aeroelastic response phenomena in aircraft. Equations for transonic flows can be obtained from the basic fluid mechanic equations. However, only numerical methods are able to attain practical solutions for those set of differential equations at the present moment. For the specific case of treating transonic flow problems, the nonlinear Euler equations provide a set of partial differential equations with features to capture nonlinear effects of typical compressible flows, despite of not accounting for viscous flows effects. The aim of this work is to implement a computational routine for the numerical solution of transonic flows around airfoils. The Euler equations are used and the flow field around a aerodynamic profile is discretized by finite difference method. A C-type structured mesh is used to discretize the flow around a NACA0012 airfoil. The methodology for numerical solution is based on the explicit MacCormack method which has second order accuracy in time and space. Based on the upwind approximation, artificial dissipation with nonlinear coefficients is incorporated to the method. The steady transonic flow around the NACA0012 airfoil numerical solution is assessed and the main flow properties are presented. Shock wave structure can also be observed by means of the Mach number contours around the airfoil. Pressure distributions on upper and lower surfaces for different flow conditions are also shown, thereby allowing the observation of the effects of the abrupt pressure change due to shock waves. The results are validated using data presented in the technical literature. The present code solutions agree well with the solution obtained in other computational codes used for the same problem

Neste trabalho, desenvolve-se um método de fronteira imersa para o estudo de escoamento compressível modelado pelas equações de Euler bidimensionais. O método de discretização de diferenças finitas é empregado, usando o método de Steger-Warming de ordem dois para discretizar as variáveis espaciais e o esquema de Runge-Kutta de ordem quatro para discretizar as variáveis temporais. O método da fronteira imersa foi empregado para o estudo de aeroelasticidade computacional em uma seção típica de aerofólio bidimensional com dois movimentos prescritos: torsional e vertical, com o objetivo de se verifcar a eficiência do método e sua aplicabilidade para problemas em aeroelasticidade computacional. Neste estudo desenvolveu-se também um programa de computador para simular escoamentos compressíveis de fluido invíscido utilizando a metodologia proposta. A verificação do código gerado foi feita utilizando o método das soluções manufaturadas e o problema de reflexão de choque oblíquo. A validação foi realizada comparando-se os resultados obtidos para o escoamento ao redor de uma seção circular e de uma seção de aerofólio NACA 0012 com os resultados experimentais, para cada caso. / In this work, an immersed boundary method is developed to study compressible flow modeled by the two-dimensional Euler equations. The finite difference method is employed, using the second order Steger-Warming method to discretizate the space variables and the fourth order Runge-Kutta method to discretizate the time variables. The immersed boundary method was employed to study computational aeroelasticity on a typical two-dimensional airfoil section with two prescribed motion: pitching and plunging, in order to verify the efficiency of the numerical method and its applicability in computational aeroelasticity problems. In this work, a computer program was developed to simulate compressible flows for inviscid fluids using the methodology proposed. The verification of the computational code was performed using the method of manufactured solutions and the oblique shock wave reflection problem. The validation was performed comparing the obtained results for flows around a circular section and a NACA 0012 airfoil section with the experimental results, for each case.