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[pt] Entradas de ar submersas são utilizadas em diversos sistemas de uma aeronave, tais como motor, ar-condicionado, ventilação e turbinas auxiliares. Neste trabalho visa-se estudar, através de simulações numéricas, a influência do uso de um gerador de vórtices sobre a espessura da camada limite a montante de uma entrada de ar submersa, com o intuito de reduzi-la e, assim, aumentar o desempenho deste tipo de entrada. O escoamento em uma entrada NACA convencional é analisado numericamente e seus resultados são tomados como referência para comparações subseqüentes. Em seguida, o gerador de vórtices é projetado e acoplado à entrada NACA convencional. Uma análise paramétrica numérica da influência da posição horizontal, do ângulo de ataque e da área do gerador de vórtices sobre a estrutura do escoamento e sobre os parâmetros de desempenho da entrada de ar é apresentada. Finalmente, um mastro de suporte do gerador de vórtices é projetado e são realizadas simulações do conjunto entrada NACA com gerador de vórtices e mastro para três ângulos de derrapagem do mastro. Os resultados mostram que a presença do gerador de vórtices livre leva a reduções consideráveis da espessura da camada limite e, consequentemente, a ganhos significativos nos parâmetros de desempenho da entrada de ar. Para o caso da entrada NACA com gerador de vórtices livre, os ganhos obtidos em relação à entrada NACA convencional, em termos de eficiência e de vazão mássica, são de até 58% e 21%, respectivamente. No caso da entrada NACA com gerador de vórtices e mastro, o melhor resultado exibe ganhos da ordem de 53%, em termos de eficiência, e de 19%, em termos da vazão mássica que ingressa na entrada de ar. A contribuição do arrasto provocado pela presença do gerador de vórtices com mastro no arrasto total do conjunto entrada NACA com gerador de vórtices e mastro é pequena, menor que 10%. / [en] Submerged air inlets are used for several systems of an aircraft, such as engine, air conditioning, ventilation, and auxiliary turbines. This work intends, through numerical simulations, to study the influence of the use of a vortex generator upon the boundary layer that develops upstream of a submerged air intake, with the aim of decreasing its thickness and, thus, to increase the inlet performance. The flow in a conventional NACA inlet is analyzed numerically and its results are considered as a reference for subsequent comparisons. Then, the vortex generator is designed and assembled to the conventional NACA inlet. A parametric analysis of the influence of the horizontal position, the angle of attack, and the area of the vortex generator on the flow field structure and on the performance parameters of the air inlet is presented. Finally, a support mast of the vortex generator is designed, and simulations are performed for the ensemble NACA inlet with vortex generator and mast for three sideslip angles of the support. The results show that the presence of the vortex generator is responsible for considerable reductions of the boundary layer thickness and, consequently, significant improvements of the performance parameters of the NACA inlet. The improvements, relative to the conventional NACA intake, in terms of ram recovery ratio and mass flow rate, may reach of 58% and 21%, respectively, for the case of the NACA inlet with the freely standing vortex generator. For the case of the NACA inlet with the vortex generator and support, improvements of up to 53%, in terms of ram recovery ratio, and 19%, in terms of mass flow rate ingested by the intake, were obtained. The contribution of the drag induced by the presence of the vortex generator with support on the total drag of the ensemble is smaller than 10%.

In this thesis, we investigate flow past an array of micro-electro-mechanical-type (MEMS-type) heating elements placed on a flat surface, where MEMS devices have hump-shaped surfaces, using the triple deck theory. In this work we start by investigating the problem with a single heating element. MEMS devices can be used to control the fluid dynamics over the surface. Hence, we present a review of the boundary layer and the triple deck theories, followed by a literature review of the problem of flow past an array of MEMS devices. Next, we formulate our problem with the aid of the method of matched expansions for supersonic and subsonic flows. Thirdly, we solve analytically the linear version of the problem for supersonic flows. Thereafter, the non-linear problem is solved numerically where a detailed description of a hybrid method to solve the formulated non-linear problem for supersonic flow is exhibited. Fourthly, for subsonic flows we continue investigating flow past a heating element placed on a flat surface. Linear analysis of this problem is conducted. A novel numerical method to solve the non-linear problem for subsonic flows is described. The results are then discussed. In a similar context, we formulate a problem which can be considered as an the extension of previous subsonic flow problem to the three dimensional case. Analytical results are obtained using the Fourier transform where the linear approximation of the problem is considered and numerical results are then obtained using the Fast Fourier Transform. Finally, we consider a case of transonic flow past a heating element placed on a flat surface, where MEMS device has a hump-shaped surface. This transonic flow problem is non-linear in the upper deck and the lower deck equations where they should be solved simultaneously. Hence, a numerical method is required where we will use a finite difference method in stream-wise direction and Chebyshev collocation method in the wall normal direction. The results are then analysed. In conclusion, the use of localised heating elements in boundary layers for flow types considered in the thesis can contribute to the possibility of favourably controlling the fluid flow perturbations.

In this research, subsonic rarefied flows over a flat-plate at constant pressure are investigated using the direct simulation Monte Carlo (DSMC) technique. An infinitely thin plate (either finite or semi-infinite) with zero angle of attack is considered. Flows with a Mach number of 0.102 and 0.4 and a Reynolds number varying between 0.063 and 246 are considered covering most of the transitional regime between the free-molecule and the continuum limits. A two-dimensional DSMC code of G.A. Bird is used to simulate these flows, and the code is modified to examine the effects of various inflow and outflow boundary conditions. It is observed that simulations of the subsonic rarefied flows are sensitive to the applied boundary conditions. Several extrapolation techniques are considered for the evaluation of the flow properties at the inflow and outflow boundaries. Among various alternatives, four techniques are considered in which the solutions are found to be relatively less sensitive. In addition to the commonly used extrapolation techniques, in which the flow properties are taken from the neighboring boundary cells of the domain, a newly developed extrapolation scheme, based on tracking streamlines, is applied to the outflow boundaries, and the best results are obtained using the new extrapolation technique together with the Neumann boundary conditions. With the new technique, the flow is not distorted even when the computational domain is small. Simulations are performed for various freestream conditions and computational domain configurations, and excellent agreement is obtained with the available experimental data.

This thesis investigates the turbulent flow in the wake of the wing-body junction of the NASA Common Research Model to further reveal its complex vortical structure and to contribute to the reference database used for Computational Fluid Dynamics validation activities. Compressible flows near two wall-boundary layers occurs not only at the wing-body junction but at every control surface of an airplane, therefore increased knowledge about this complex flow structure could potentially improve the estimates of drag performance and control surface efficiency, primarily for minimizing the environmental impact of commercial flight. The airplane model is modified by adding an inboard flap to investigate the influence from the deflection on the vorticity and velocity field. Future flap designs and settings are discussed from a performance improvement point of view, with the investigated flow influence in mind. The experimental measurements for this thesis were collected using a Cobra Probe, a dynamic multi-hole pressure probe, for Reynolds numbers close to one million based on the wing root chord. A pre-programmed three-dimensional grid was used to cover the most interesting parts of the junction flow. The facility used for the tests is a 120 cm by 80 cm indraft, subsonic wind tunnel at NASA Ames Research Center’s Fluid Mechanics Lab, which provides an on-set flow speed of around Mach 0.15, corresponding to approximately 48 m/s. / Den här avhandlingen undersöker det turbulenta flödet runt övergången mellan flygplanskropp och vinge på en NASA Common Research Model för att vidare utforska den komplexa, tredimensionella strukturen av flödet och bidra till NASA’s officiella databas för jämförelser med simulerade flöden. Kompressibla flöden nära tvåväggsgränsskikt uppkommer inte bara vid övergången mellan flygplanskropp och vinge utan även vid varje kontrollyta på ett flygplan. Ökad kunskap om flödets beteende vid sådana områden kan därför bidra till en bättre uppskattning av prestanda och effektivitet av kontrollytorna och flygplanet i sin helhet, vilket kan bidra till minskad miljöpåverkan från kommersiell flygtrafik. Flygplansmodellen är modifierad genom montering av en vingklaff på den inre delen av vingen, detta för att undersöka hur olika vinklar på klaffarnas nedböjning påverkar flödets struktur och hastighetsfält. Framtida klaffdesigner och inställningar för ökad prestanda diskuteras även utifrån denna påverkan. Mätningarna i vindtunneln gjordes med en Cobra Probe, ett dynamisk tryckmätningsinstrument, speciellt designad för turbulenta och instabila flöden. Reynoldsnumren som generades av den subsoniska, indrags-vindtunneln var ungefär en miljon baserad på vingrotens längd, vilket motsvarar knappt en tiondel av normala flygförhållanden för samma flygplansmodell.