With the advent of powerful computer hardware, Computational Fluid Dynamics (CFD) has been vastly used by researches and scientists to investigate flow behavior and its properties. The cost of CFD simulation is very small compared to the experimental arsenal as test facilities and wind-tunnels. In the last years many CFD commercial packages were developed and some of them possess prominence in industry and academia. However, some specific CFD calculations are particular cases and sometimes need special attention due to the complexity of the flow. In these cases, meticulous research becomes necessary. This is the case of turbomachinery flow calculations. The development of CFD codes applied to turbomachinery flow simulations and its implementation issues are not available. A few institutions have this type of knowledge. Each CFD code has its particularities. Developing a CFD code is very interest subject in academia. In this work, a computational code, written in FORTRAN, was developed to calculate internal flows in turbomachines using CFD techniques. The solver is capable of calculating the three-dimensional flows not only for turbomachines. For instance, internal and external flows of nozzles and airfoils can be calculated. The approach used allows the use of unstructured meshes of hexahedral elements. Euler, Navier-Stokes and turbulent equations can be calculated depending on the user settings. Diferent numerical schemes were implemented for time and space integration. Numerical tools to improve the stability and to increase the time-step (local time-step and implicit residual smoothing) were also implemented and all details are described in this work. The origin of this solver is to simulate flows in compressors and turbines. Therefore, both rotating and nonrotating frames of reference are calculated simultaneously. Hence, the verification and validation processes were run for both inertial and non-inertial systems. A step-by-step design procedure is presented in this work. It is very important to mention that to have a complete understanding of the flow physics in compressors and turbines the designer must have a solid knowledge of the operation of gas turbine components.

This work presents the results of shock tunnel tests of a Mach 10 waverider with sharp leading edges. The waverider surface was generated from a conical flowfield with the volume and the viscous lift-to-drag ratio as optimization parameters. A compression and expansion ramps were added to the pure waverider surface in order to simulate the flow over a scramjet engine. The compression ramp was designed so as to provide the ideal conditions for the supersonic combustion of the Hydrogen while the expansion section was derived from an ideal minimum length supersonic nozzle. The experimental data included Schlieren photographs of the flow and the pressure distribution over the compression surface. These data were compared with the inviscid theory. During these investigations, the IEAv's T3 shock tunnel was used to simulate the hypersonic flow. The stagnation conditions as well as the free stream properties were estimated using numerical codes. The tunnel operated at Mach number ranges of 8.9 to 10, Reynolds number from 2.25 x 106 to 8.76 x 106 (m-1) and Knudsen number from 0.06 to 0.19. From the Schlieren photographs it was noted that the inlet flowfield behaves according to the predictions of the hypersonic viscous interaction models. Also, the pressure variation along the compression surface centerline was obtained using piezoelectric pressure sensors. The resulted profile presented the general trend of the flow described by these models.

A Computational Aeroacoustics (CAA) and a novel semi-empirical model is developed for predicting the noise generated by the jet flow through dual stream (coaxial) nozzles, as found in modern turbofan engines. The acoustic source model was developed in a 2D and 3D framework, based on the Lilley's Equations, following the traditional MGBK method from NASA Langley Research Center. The semi-empirical model was based on the Four-Source model from the Institute of Sound and Vibration (ISVR). This suite of methodologies provided a mean of investigating the mechanisms of noise generation and propagation of subsonic coaxial jet flows, as well as the noise prediction at different operating conditions. The work done contributed to the development and improvement of a numerical tool for jet noise prediction of dual-stream exhaust systems, commonly employed in turbofan engines. Such research also subsidies the improvement of semi-empirical methods used in the Center of Reference in Gas Turbine (ITA) for the noise prediction of turbofans in all operating conditions.

O comportamento transiente do escoamento de combustível e do ar no interior de tanques de aeronaves é uma informação que é obtida tradicionalmente somente através de ensaios. Esta dissertação propõe uma metodologia para simular o comportamento do combustível no interior de tanques através de métodos computacionais. Com o estudo do comportamento do combustível para condições de reabastecimento e durante manobras é possível avaliar aspectos importantes para otimizar o projeto dos tanques em função do tipo de operação da aeronave. As ferramentas utilizadas para efetuar as simulações têm como base casos matemáticos e simulações em CFD. Esta dissertação propõe a avaliação do comportamento do combustível em um tanque para duas situações: a primeira considerando somente a ação da gravidade com o intuito de simular o movimento do combustível após uma manobra e a segunda situação considerando o processo de reabastecimento. A avaliação do comportamento do combustível tem como finalidade prover informações sobre o tempo de reabastecimento, tempo de estabilização do movimento do combustível, posicionamento da fronteira combustível-ar, dados de velocidade e pressão. Através das informações referentes ao comportamento do combustível é possível a identificação de problemas em fase prematura de projeto, redução de custos de projeto, e redução da necessidade de ensaios em bancada e em vôo.

The present work is aimed at studying the influence of viscous effects in transonic aeroelastic analyses. To achieve this goal, a two-dimensional and viscous aeroelastic computational solver, for CAE analysis, is developed, which uses unstructured computational meshes and which is able to capture the main aeroelastic phenomena relevant in the transonic regime of flight. The aeroelastic system considered to test the present methodology is the classical typical section model. The system has two structural degrees of freedom. These are pitching and plunging, or heaving. The structural degrees of freedom can be treated within solver in a coupled manner or separately, in a loosely coupled fashion. The typical section model is an approximation to the treatment of a full wing, in which the airfoil at 75% of the semi-span is analyzed. The structural response is obtained by solving a set of a second order ordinary differential equations in time, with aerodynamic forcing. The coupling of the structural degrees of freedom occurs primarily through the aerodynamic forcing terms. The unsteady aerodynamic problem is treated through the numerical solution of the Reynolds-averaged Navier-Stokes equations. These equations are solved using a finite volume method for unstructured computational grids, which uses a second-order centered spatial discretization and a second order time marching scheme. Turbulence closure is achieved through the Spalart-Allmaras one-equation eddy viscosity turbulence model. A reduction of the computational time for the unsteady aerodynamic simulations is obtained through the implmentation of a few convergence acceleration methods, which include the use of a constant CFL number, implicit residual smoothing and unsteady multigrid methods. The aeroelastic problem is solved through the coupling of the aerodynamic and structural formulations. In the present case, the structural equations are cast in a modal formulation and the unsteady aerodynamic responses are represented by aerodynamic states obtained by rational interpolating polynomials. The complete system of equations is written in state space format in the Laplace domain. The aeroelastic stability condition can, then, be determined by standard eigenvalue analyses of the system dynamic matrix.

The primary means of preventing ice formation on wings and engine inlets for modern commercial transport aircraft is by extracting hot air from the compressor and blowing it on the inside surface of the leading edge through small holes drilled in the so-called piccolo tube system. A critical aspect in the design of such system is the prediction of heat transfer of the impinging jets from the piccolo tube. The correct evaluation of the heat transfer rate in such devices is of great interest to optimize both the anti-icing performance and the hot air bleeding from the high-pressure compressor. The history of research in the anti-icing area is rather narrow. A review of the literature reveals that only few experimental and theoretical/numerical studies have been carried out to study the heat transfer and flow in the internal hot-air region. There are some experimental and numerical studies that developed correlations for the average Nusselt number. However, most of the research was performed using a single jet or a group of jets impinging on a flat slat, which is different from the jet impingement on concave surfaces, as the inside surface of a wing. Therefore, the objective of the present work is use the commercial CFD software FLUENT to perform a parametric study of the jet impingement on concave surfaces. The main goal is determine the effect of the Reynolds number on the heat transfer process. At the end of the work, a correlation for the average Nusselt number which account for the Reynolds number is presented.

Since the 50's, with the appearance of the turbojet engines, the jet noise is being exhaustively studied, because it is one of the most important source of aircraft noise. Many attempts have been made to reduce the jet noise, including higher by-pass turbofan engines. Chevron nozzles also have been used by the industry to try to reduce the jet noise with a low performance and weight penalty. This work shows a computation procedure to assess how this noise suppression devices impact on both fluid dynamics and acoustics of single isothermal jets. Towards this goal, different chevron nozzles, with 6, 8 and 12 lobes have been analyzed. The calculation procedure is based on a Reynolds Average Navier-Stokes calculation, followed by a stochastic noise generation and radiation method, resulting in a relatively fast noise calculation procedure. The simulations have been carried out using the commercial software CFD++. The calculation procedure has predicted the expected fluid dynamic and acoustic behavior for chevron nozzles, e.g., shortening the potential core length, high frequency noise increase and low frequency noise attenuation. The parametric study of the number of lobes has shown that this parameter impacts the mixing region. Moreover, varying this parameter is a way to attain different low frequency reductions, without great impacts on the highest frequencies. Although the procedure did not capture correctly the absolute values of the acoustic response, the results show that this relatively simple and quick analysis reproduced important parameters in designing new nozzles and can be used as a way to better understand the influence of chevrons.

Este trabalho tem por finalidade analisar um método de fluxo potencial para cálculo de coeficientes aerodinâmicos através de uma comparação com resultados experimentais. O método utilizado é o método malha de dipolos, cujas vantagens figuram sobretudo em sua alta velocidade de cálculo e de análise, e razoável precisão nos resultados, razões pelas quais a metodologia desenvolvida neste trabalho é direcionada principalmente para as fases iniciais de projeto aeronáutico.

Neste trabalho apresenta-se uma metodologia para a estimativa dos efeitos de interferência aerodinâmica sobre empenagens devido a jatos turbulentos, subsônicos e aquecidos. Primeiramente emprega-se um método de diferenças finitas para resolver numericamente as equações de Reynolds com as aproximações de camada limite. As equações de reynolds são resolvidas simultaneamente com a equação diferencial parcial oriundado modelo (K) da energia cinética turbulenta desenvolvido por Prandtl. Desta forma são calculados o campo de velocidade e de temperatura resultante. Na segunda etapa modela-se o campo de velocidade radial induzido pelo jato através de uma distribuição linear de sumidouros. Esta distribuição é utilizada com o "Vortex Lattice Method" no cálculo da interferência aerodinâmica sobre uma empenagem horizontal. Finalmente, são obtidas as curvas de sustentação de uma empenagem horizontal em função das principais variáveis do escoamento do jato.

The present work is primarily concerned with studying the influence of an upwind spatial discretization on the capability of representing turbulent flows on aerospace applications, in the context of a flow simulation code that is fairly close to a production code. Therefore, the work addresses the issues of implementing and validating an advanced turbulence model for high Reynolds number aerospace applications in the context of an existing flux-vector splitting simulation tool, which incorporates several advances in current CFD practice, including parallel processing. The flow simulation tool used in the present work was originally developed for high speed, high altitude, hypersonic applications. Hence, the code did not include any provisions for turbulence modeling, since most flows at these conditions can be adequately treated as laminar flows. Moreover, due to the presence of strong shock waves, which are typical of hypersonic applications, a very dissipative spatial discretization scheme, based on the upwind flux vector splitting concept, was employed in the construction of the inviscid numerical fluxes. Therefore, the use of such a tool for the simulation of turbulent aerospace flows requires the implementation of a turbulence closure, as well as an adequate treatment of the excessive artificial dissipation automatically generated by the original spatial discretization scheme. In the present case, the flows of interest are simulated using the three-dimensional Reynolds-averaged Navier-Stokes equations. The turbulence closure considered is the one-equation, eddy viscosity, Spalart-Allmaras model. The work discusses in detail the theoretical and numerical formulation of the selected model, as well as the validation studies. The work also demonstrates how the spatial discretization scheme is selectively modified such that the flow simulation tool remains robust for high speed applications at the same time that it can accurately compute turbulent boundary layers. Furthermore, the work also addresses the parallelization and other high performance computational issues, demonstrating that the resultant flow simulation code can achieve adequate performance on current multi-CPU, multi-core computational clusters. Finally, the work discusses issues that could be considered for the continuation of the research effort here undertaken.