The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag reduction of the vehicle along with the minimized heat transfer rate. Using forward facing solid aerospike is the simplest way among the existing drag reduction methodologies for hypersonic blunt cone bodies. But the flow oscillations associated with this aerospike makes it difficult to implement. When analyzing this flow, it can be understood that this oscillating flow can be compared to conical cavity flow. Therefore in the spiked flows, it is decided to implement the technique used in reducing the flow oscillation of the cavities. Based on this method the shallow conical cavity flow generated by the aerospike fixed ahead of a 120o blunt cone body is fissured as multiple cavities by so many disks formed from 10o cone. Now the deep conical cavities had the length to mean depth ratio of unity; this suppresses the unnecessary oscillations of the shallow cavity. The total length of the telescopic aerospike is fixed as 100mm. And one another conical tip plain aerospike of same length is designed for comparing the telescopic spike’s performance at hypersonic flow Mach numbers of 5.75 and 7.9. A three component force balance system capable of measuring drag, lift and pitching moment is designed and mounted internally into the skirt of the model. Drag measurement is done for without spike, conical tip plain spiked and telescopic spiked blunt cone body. The three configurations are tested at different angles of attack from 0 to 10 degree with a step of 2. A discrete iterative deconvolution methodology is implemented in this research work for obtaining the clean drag history from the noisy drag accelerometer signal. The drag results showed the drag reduction when compared to the without spike blunt cone body. When comparing to the plain spiked, the telescopic spiked blunt cone body has lesser drag at higher angles of attack. Heat transfer measurements are done over the blunt cone surface using the Platinum thin film gauges formed over the Macor substrate. These results and the flow visualization give better understanding of the flow and the heat flux rate caused by the flow. The enhancement in the heat flux rate over the blunt cone surface is due to the shock interaction. And in recirculation region the heat flux rate is very much lesser when compared to without spike blunt cone body. It is observed that the shock interaction in the windward side is coming closer towards the nose of the blunt cone as the angle of attack increases and the oscillation of the oblique shock also decreases. Schlieren visualization showed that there is dispersion in the oblique shock, particularly in the leeward side. In the telescopic spike there are multiple shocks generated from each and every disk which coalesces together to form a single oblique shock. And the effect of the shock generated by the telescopic spike is stronger than the effect of the shock generated by the conical tip plain spike.

Helicopter performance is greatly influenced by its drag. Pylons, fuselage, landing gear, and especially the rotor hub of a helicopter experience large separated flow regions, even under steady level flight conditions the vehicle has been designed for, contributing to the helicopter drag. Several passive and active flow control concepts have been studied for reducing helicopter drag. While passive flow control methods reduce drag, they do so at one optimized design condition. Therefore, passive drag reduction methods may not work for helicopters that operate under widely varying flight conditions. Active flow control (AFC) methods overcome this disadvantage and consequently are widely being pursued. The present investigator has studied some of these AFC methods using computational fluid dynamics (CFD) techniques and has found synthetic (or pulsed) jets as one of the more effective drag reduction devices. Two bluff bodies, representative of helicopter components, have been studied and the mechanism behind drag reduction has been analyzed. It was found that the increase in momentum due to the jet, and a resultant reduction in the separated flow region, is the main reason for drag reduction in these configurations. In comparison with steady jets, synthetic jets were found to use less power for a greater drag reduction. The flow inside these synthetic jet devices is incompressible. It is computationally inefficient to use compressible flow solvers in incompressible regions. In such regions, using Lattice Boltzmann equations (LBE) is more suitable compared to solving the incompressible Navier-Stokes equations. The length scales close to the synthetic jet devices are very small. LBE may be used to better resolve these small length scale regions. However, using LBE throughout the whole domain would be computationally expensive since the grid spacing in the LBE solver has to be of the order of the mean free path. To address this need, a coupled Lattice Boltzmann-Navier-Stokes (LB-NS) methodology has been developed. The LBE solver has been successfully validated in a standalone manner for several benchmark cases. The solver has also been shown to be of second order accuracy. This LBE solver has been subsequently coupled with an existing Navier-Stokes (NS) solver. Validation of the coupled methodology has been done for analytical problems with known closed form solution. This LB-NS methodology is further used to simulate the flow past a cylinder where synthetic jet devices have been used to reduce drag. The LBE solver is used in the cavity of the synthetic jet nozzle while the NS solver is employed in the rest of the domain. The cylinder configuration was chosen to demonstrate drag reduction on helicopter hub shape geometries. Significant drag reduction is observed when synthetic jets are used, compared to the baseline no flow control case.

A body in a hypersonic flow field will experience very high heating especially during re-entry. Conventionally this problem is tackled to some extent by the use of large angle blunt cones. At the cost of increased drag, the heat transfer rate is lower over most parts of the blunt body, except in a region around the stagnation point. Thus even with blunt cones, management of heat transfer rates and drag on bodies at hypersonic speeds continues to be an interesting research area. Various thermal protection systems have been proposed in the past, like heat sink cooling, ablation cooling and aerospikes. The ablative cooling system becomes extremely costly when reusability is the major concern. Also the shape change due to ablation can lead to issues with the vehicle control. The aerospikes themselves may become hot and ablate at hypersonic speeds. Hence an alternate form of cooling system is necessary for hypersonic flows, which is more feasible, cost effective and efficient than the conventional cooling systems. Injection of a mass of cold fluid into the boundary layer through the surface is one of the potential cooling techniques in the hypersonic flight corridors. These kinds of thermal protection systems are called mass transfer cooling systems. The injection of the mass may be through discrete slots or through a porous media. When the coolant is injected through a porous media over the entire surface, the coolant comes out as a continuous mass. Such a cooling system is also referred as “transpiration cooling system”. When the fluid is injected through discrete slots, the system is called as “film cooling system”. In either case, the coolant absorbs the incoming heat through its rise in enthalpy and thus modifies the boundary layer characteristics in such a way that the heat flow rate to the surface is less. Injection of a forward facing jet (opposite to the freestream direction) from the stagnation point of a blunt body can be used for mitigating both the aerodynamic drag and heat transfer rates at hypersonic Mach numbers. If the jet has enough momentum it can push the bow shock forward, resulting in reduced drag. This will also reduce heat transfer rate over most part of the body except around the jet re-attachment region. A reattachment shock impinging on the blunt body invariably increases the local heat flux. At lower momentum fluxes the forward facing jet cannot push the bow shock ahead of the blunt body and spreads easily over the boundary layer, resulting in reduced heat transfer rates. While the film cooling performance improves with mass flow rate of the jet, higher momentum flow rates can lead to a stronger reattachment leading to higher heat transfer rate at the reattachment zone. If we are able to reduce the momentum flux of the coolant for the same mass flow rate, the gas coming out can easily spread over the boundary layer and it is possible to improve the film cooling performance. In all the reported literature, the mass flow rate and the momentum flux are not varied independently. This means, if the mass flow rate is increased, there is a corresponding increase in the momentum flux. This is because the injection (from a particular orifice and for a particular coolant gas) is controlled only by the total pressure of injection and free stream conditions. The present investigation is mainly aimed at demonstrating the effect of reduction in momentum of the coolant (injected opposing a hypersonic freestream from the stagnation point of a blunt cone), keeping the mass flow rate the same, on the film cooling performance. This is achieved by splitting a single jet into a number of smaller jets of same injection area (for same injection total pressure and same free stream conditions). To the best of our knowledge there is no report on the use of forward facing micro-jet array for film cooling at hypersonic Mach numbers. In this backdrop the main objectives of the present study are: • To experimentally demonstrate the effect of splitting a single jet into an array of closely spaced smaller micro-jets of same effective area of injection (injected opposite to a hypersonic freestream from the stagnation zone), on the reduction in surface heat transfer rates on a large angle blunt cone. · Identifying various parameters that affect the flow phenomenon and doing a systematic investigation of the effect of the different parameters on the surface heat transfer rates and drag. Experimental investigations are carried out in the IISc hypersonic shock tunnel on the film cooling effectiveness. Coolant gas (nitrogen and helium) is injected opposing hypersonic freestream as a single jet (diameter 2 mm and 0.9 mm), and as an array of iv micro jets (diameter 300 micron each) of same effective area (corresponding to the respective single jet). The coolant gas is injected from the stagnation zone of a blunt cone model (58o apex angle and nose radius of 35 mm). Experiments are performed at a flow freestream Mach number of 5.9 at 0o angle of attack, with a stagnation enthalpy of 1.84 MJ/Kg, with and without injections. The ratios of the jet stagnation pressure to the pitot pressure (stagnation pressure ratio) used in the present study are 1.2 and 1.45. Surface convective heat transfer measurements using platinum thin film sensors, time resolved schlieren flow visualization and aerodynamic drag measurements using accelerometer force balance are used as flow diagnostics in the present study. The theoretical stagnation point heat transfer rate without injection for the given freestream conditions for the test model is 79 W/cm2 and the corresponding aerodynamic drag from Newtonian theory is 143 N. The measured drag value without injection (125 N) shows a reasonable match with theory. As the injection is from stagnation zone it is not possible to measure the surface heat transfer rates at the stagnation point. The sensors thus are placed from the nearest possible location from the stagnation point (from 16 mm from stagnation point on the surface). The sensors near the stagnation point measures a heat transfer rate of 65 W/cm2 on an average without any injection. Some of the important conclusions from the study are: • Up to 40% reduction in surface heat transfer rate has been measured near the stagnation point with the array of micro jets, nitrogen being the coolant, while the corresponding reduction was up to 30% for helium injection. Considering the single jet injection, near the stagnation point there is either no reduction in heat transfer rate or a slight increase up to 10%. · Far away from stagnation point the reduction in heat transfer with array of micro-jets is only slightly higher than corresponding single jet for the same pressure ratio. Thus the cooling performance of the array of closely spaced micro jets is better than the corresponding single jet almost over the entire surface. • The time resolved flow visualization studies show no major change in the shock standoff distance with the low momentum gas injection, indicating no major changes in other aerodynamic aspects such as drag. · The drag measurements also indicate that there is virtually no change in the overall aerodynamic drag with gas injection from the micro-orifice array. · The spreading of the jets injected from the closely spaced micro-orifice array over the surface is also seen in the visualization, indicating the absence of a region of strong reattachment. · The reduction in momentum flux of the injected mass due to the interaction between individual jets in the case of closely spaced micro-jet array appears to be the main reason for better performance when compared to a single jet. The thesis is organized in six chapters. The importance of film cooling at hypersonic speeds and the objectives of the investigation are concisely presented in Chapter 1. From the knowledge of the flow field with counter-flow injection obtained from the literature, the important variables governing the flow phenomena are organized as non-dimensional parameters using dimensional analysis in Chapter 2. The description of the shock tunnel facility, diagnostics and the test model used in the present study is given in Chapter 3. Chapter 4 describes the results of drag measurements and flow visualization studies. The heat transfer measurements and the observed trends in heat transfer rates with and without coolant injection are then discussed in detail in Chapter 5. Based on the obtained results the possible physical picture of the flow field is discussed in Chapter 6, followed by the important conclusions of the investigation.

CAPES / A indústria petrolífera utiliza diversos tipos de bombas conforme a natureza da operação ou a fase do processamento considerada. Entretanto, devido às condições nas quais opera esse tipo de indústria, óleo e gás podem ser produzidos simultaneamente, o que constitui uma questão de escoamento bifásico. A presença de gás em canais de bombas centrífugas prejudica significativamente o seu funcionamento, degradando a capacidade de elevação. As pesquisas relacionadas ao escoamento bifásico em bombas centrífugas centram seus estudos na avaliação da influência de condições operacionais, tais como a fração de vazio, rotação do rotor e pressão de entrada, no desempenho global da bomba. Essa influência, entretanto, é condicionada ao padrão de escoamento bifásico no interior do equipamento. Nesse cenário, o presente trabalho tem como objetivo avaliar o escoamento bifásico nos canais de uma bomba centrífuga mediante o comportamento de uma bolha isolada no meio líquido em rotação. Para isso, foi construída uma bancada experimental no laboratório do NUEM. A carcaça da bomba e o rotor do primeiro estágio originais foram substituídos por outros de material transparente a fim de permitir a visualização das bolhas no interior do canal do rotor. Em particular, pretende-se seguir o movimento de bolhas isoladas nos canais do rotor, como forma de identificar seus caminhos preferenciais, para avaliar suas velocidades ponto a ponto. Além disso, foi desenvolvido um estudo numérico como forma de obter as velocidades do líquido e pressão estática ao longo do canal, utilizando as posições das bolhas obtidas experimentalmente. Um modelo algébrico utilizou os dados numéricos e experimentais para calcular o coeficiente de arrasto e a força de arrasto segundo diferentes condições operacionais. / Electrical Submersible Pump (ESP), are quite common in the oil industry. Due to the nature of the oil production operations, multiphase flow of oil, gas and other fluids are present in reservoirs, pipelines and equipment, including ESP’s. Two-phase flows inside pump diffusers and impellers decrease an ESP’s lift and efficiency. Literature presents a large number of studies about the influence of operational parameters such as void fraction, rotor speed and inlet pressure, on an ESP’s global efficiency. Nevertheless, this influence is strictly related to the two-phase flow pattern. The present study evaluates the two-phase flow inside a pump diffuser and impeller by means of a single bubble flowing through the liquid mass. An experimental apparatus was designed and built at the NUEM facilities. The original pump housing and impeller were replaced by transparent pieces so that bubbles flowing inside the pump could be visualized. The path of the bubbles were followed so as to both define preference paths and to calculate bubbles velocities. Additional CFD analyses provided liquid velocities and static pressure through the pump impeller. An algebraic model fed with numerical and experimental data evaluated the drag coefficient and the drag force according to different experimental conditions.

O presente estudo consistiu em avaliar os parâmetros aerodinâmicos de um modelo de ônibus rodoviário, comparando os resultados obtidos de simulação computacional via CFD (Computational Fluid Dynamics) com aqueles obtidos na experimentação em túnel de vento. O ônibus estudado foi do tipo rodoviário de um fabricante local, modelo Paradiso 1200. O veículo foi modelado em um software CAD (SolidWorks®) em duas escalas: 1/42 e 1/24. Além disso, para obter a comparação com a literatura, foram analisados dois tamanhos diferentes de um modelo do corpo de Ahmed. Posteriormente, foram criadas as malhas com as geometrias 3D e realizados os testes computacionais no software ANSYS FLUENT® para os quatro modelos, com o intuito de identificar alguns parâmetros aerodinâmicos como o coeficiente de arrasto, coeficiente de pressão, entre outros. Para as análises com o corpo de Ahmed foram utilizados os modelos de turbulência Spalart – Allmaras, κ – ε Standard, κ – ε RNG, κ – ω Standard, κ – ω SST e SST. Para os modelos de ônibus foram simulados apenas o modelo κ – ε Standard. Para a realização dos experimentos foi empregado um túnel de vento de circuito aberto, onde foram realizados testes de distribuição de pressão e arrasto aerodinâmico, variando a altura do vão livre entre a mesa automobilística e a superfície inferior dos modelos. Nos ensaios dos modelos onde houve a variação da altura em relação à mesa automobilística, foi identificado um aumento de 4,5% no valor do coeficiente de arrasto (Cd) para o corpo de Ahmed menor e 6,1% para o ônibus em escala 1/42. Comparando-se os resultados obtidos nos ensaios experimentais com aqueles obtidos nas análises numéricas, também ocorreram variações no Cd para todos os modelos. Nos ensaios de pressão o coeficiente de pressão (Cp) foi praticamente o mesmo entre os valores obtidos na análise em CFD e os valores experimentais, para ambos os modelos. Foram também realizados ensaios de visualização usando tufts de lã distribuídos na superfície externa do modelo menor de Ahmed e do modelo maior do ônibus. Esses ensaios indicaram nitidamente as regiões de recirculação de ar nos modelos, o que em parte não foi possível observar na análise computacional. Diante disso, verifica-se que os resultados experimentais obtidos em túnel de vento ainda são os mais confiáveis e utilizados, apesar dos altos custos envolvidos na construção de modelos, na instrumentação de alta tecnologia hoje disponível, nos métodos de visualização e na energia consumida nos testes. / Submitted by Ana Guimarães Pereira (agpereir@ucs.br) on 2016-12-09T18:14:15Z No. of bitstreams: 1 Dissertacao Giovanni Matheus Rech.pdf: 20273449 bytes, checksum: 70c97b02eb18e1d8b69454f3e92fe23d (MD5) / Made available in DSpace on 2016-12-09T18:14:15Z (GMT). No. of bitstreams: 1 Dissertacao Giovanni Matheus Rech.pdf: 20273449 bytes, checksum: 70c97b02eb18e1d8b69454f3e92fe23d (MD5) Previous issue date: 2016-12-09 / The present study was to evaluate the aerodynamic parameters of a road bus model by comparing the results of computer simulation via CFD (Computational Fluid Dynamics) with those obtained in experiments in a wind tunnel. The bus studied was a road type from a local manufacturer, Paradiso 1200 model. The vehicle was modeled on a CAD software (SolidWorks®) on two scales: 1/42 and 1/24. Furthermore, for comparison with the literature, we analyzed two different sizes of Ahmed body model. Thereafter, the meshes were created from 3D geometry and the computational tests performed with FLUENT® ANSYS software for the four models in order to identify some aerodynamic parameters such as the drag coefficient, pressure coefficient, among others. For analysis of Ahmed bodies, the turbulence models Spalart - Allmaras, κ - ε Standard, κ - ε RNG, κ - ω Standard, κ - ω SST and SST were used. For bus models, the turbulence model κ - ε Standard was only used. For the experiments we used an open circuit wind tunnel, where tests of pressure distribution and aerodynamic drag were performed, varying the height of the clearance between the automotive table and the bottom surface of the models. In the model tests, in which there were the height variation relative to the automotive table, an increase of 4.5% in the value of the drag coefficient (Cd) for the lower Ahmed body, and 6.1% for the bus 1/42 scale were identified. In pressure tests, the pressure coefficients (Cp) were almost the same between the values obtained from the CFD analysis and experimental values for both models. Visualization tests using wool tufts distributed on the outer surface of the smaller Ahmed model and the higher bus model were also performed. These tests clearly indicated the air recirculation regions in models, which in part was not observed in the computational analysis. Thus, it appears that the experimental results are in wind tunnel still the most reliable and used despite the high costs involved in the building models, in the high-tech instrumentation available today, in the visualization methods and in the energy consumed in the tests.

O presente estudo consistiu em avaliar os parâmetros aerodinâmicos de um modelo de ônibus rodoviário, comparando os resultados obtidos de simulação computacional via CFD (Computational Fluid Dynamics) com aqueles obtidos na experimentação em túnel de vento. O ônibus estudado foi do tipo rodoviário de um fabricante local, modelo Paradiso 1200. O veículo foi modelado em um software CAD (SolidWorks®) em duas escalas: 1/42 e 1/24. Além disso, para obter a comparação com a literatura, foram analisados dois tamanhos diferentes de um modelo do corpo de Ahmed. Posteriormente, foram criadas as malhas com as geometrias 3D e realizados os testes computacionais no software ANSYS FLUENT® para os quatro modelos, com o intuito de identificar alguns parâmetros aerodinâmicos como o coeficiente de arrasto, coeficiente de pressão, entre outros. Para as análises com o corpo de Ahmed foram utilizados os modelos de turbulência Spalart – Allmaras, κ – ε Standard, κ – ε RNG, κ – ω Standard, κ – ω SST e SST. Para os modelos de ônibus foram simulados apenas o modelo κ – ε Standard. Para a realização dos experimentos foi empregado um túnel de vento de circuito aberto, onde foram realizados testes de distribuição de pressão e arrasto aerodinâmico, variando a altura do vão livre entre a mesa automobilística e a superfície inferior dos modelos. Nos ensaios dos modelos onde houve a variação da altura em relação à mesa automobilística, foi identificado um aumento de 4,5% no valor do coeficiente de arrasto (Cd) para o corpo de Ahmed menor e 6,1% para o ônibus em escala 1/42. Comparando-se os resultados obtidos nos ensaios experimentais com aqueles obtidos nas análises numéricas, também ocorreram variações no Cd para todos os modelos. Nos ensaios de pressão o coeficiente de pressão (Cp) foi praticamente o mesmo entre os valores obtidos na análise em CFD e os valores experimentais, para ambos os modelos. Foram também realizados ensaios de visualização usando tufts de lã distribuídos na superfície externa do modelo menor de Ahmed e do modelo maior do ônibus. Esses ensaios indicaram nitidamente as regiões de recirculação de ar nos modelos, o que em parte não foi possível observar na análise computacional. Diante disso, verifica-se que os resultados experimentais obtidos em túnel de vento ainda são os mais confiáveis e utilizados, apesar dos altos custos envolvidos na construção de modelos, na instrumentação de alta tecnologia hoje disponível, nos métodos de visualização e na energia consumida nos testes. / The present study was to evaluate the aerodynamic parameters of a road bus model by comparing the results of computer simulation via CFD (Computational Fluid Dynamics) with those obtained in experiments in a wind tunnel. The bus studied was a road type from a local manufacturer, Paradiso 1200 model. The vehicle was modeled on a CAD software (SolidWorks®) on two scales: 1/42 and 1/24. Furthermore, for comparison with the literature, we analyzed two different sizes of Ahmed body model. Thereafter, the meshes were created from 3D geometry and the computational tests performed with FLUENT® ANSYS software for the four models in order to identify some aerodynamic parameters such as the drag coefficient, pressure coefficient, among others. For analysis of Ahmed bodies, the turbulence models Spalart - Allmaras, κ - ε Standard, κ - ε RNG, κ - ω Standard, κ - ω SST and SST were used. For bus models, the turbulence model κ - ε Standard was only used. For the experiments we used an open circuit wind tunnel, where tests of pressure distribution and aerodynamic drag were performed, varying the height of the clearance between the automotive table and the bottom surface of the models. In the model tests, in which there were the height variation relative to the automotive table, an increase of 4.5% in the value of the drag coefficient (Cd) for the lower Ahmed body, and 6.1% for the bus 1/42 scale were identified. In pressure tests, the pressure coefficients (Cp) were almost the same between the values obtained from the CFD analysis and experimental values for both models. Visualization tests using wool tufts distributed on the outer surface of the smaller Ahmed model and the higher bus model were also performed. These tests clearly indicated the air recirculation regions in models, which in part was not observed in the computational analysis. Thus, it appears that the experimental results are in wind tunnel still the most reliable and used despite the high costs involved in the building models, in the high-tech instrumentation available today, in the visualization methods and in the energy consumed in the tests.

This project evaluates the suitability and practicality of the program Athena Vortex Lattice (AVL) by Mark Drela. A short user guide was written to make it easier (especially for students) to get started with the program AVL. AVL was applied to calculate the induced drag and the Oswald factor. In a first task, AVL was used to calculate simple wings of different aspect ratio A and taper ratio lambda. The Oswald factor was calculated as a function f(lambda) in the same way as shown by HOERNER. Compared to HOERNER's function, the error never exceed 7.5%. Surprisingly, the function f(lambda) was not independent of aspect ratio, as could be assumed from HOERNER. Variations of f(lambda) with aspect ratio were studied and general results found. In a second task, the box wing was investigated. Box wings of different h/b ratio: 0.31, 0.62, and 0.93 were calculated in AVL. The induced drag and Oswald factor in all these cases was calculated. An equation, generally used in the literature, describes the box wing's Oswald factor with parameters k1, k2, k3 and k4. These parameters were found from results obtained with AVL by means of the Excel Solver. In this way the curve k = f(h/b) was plotted. The curve was compared with curves with various theories and experiments conducted prior by other students. The curve built based on AVL fits very well with the curve from HOERNER, PRANDTL and a second experiment made in the wind tunnel at HAW Hamburg.

The interaction of a shock wave with a boundary layer is a classic viscous/inviscid interaction problem that occurs over a wide range of high speed aerodynamic flows. For example, on transonic wings, in supersonic air intakes, in propelling nozzles at offdesign conditions and on deflected controls at supersonic/transonic speeds, to name a few. The transonic interaction takes place at Mach numbers typically between 1.1 and 1.5. On an aerofoil, its existence can cause problems that range from a mild increase in section drag to flow separation and buffeting. In the absence of separation the drag increase is predominantly due to wave drag, caused by a rise in entropy through the interaction. The control of the turbulent interaction as applied to a transonic aerofoil is addressed in this thesis. However, the work can equally be applied to the control of interaction for numerous other occurrences where a shock meets a turbulent boundary layer. It is assumed that, for both swept normal shock and unswept normal shock interactions, as long as the Mach number normal to the shock is the same, then the interaction, and therefore its control, should be the same. Numerous schemes have been suggested to control such interaction. However, they have generally been marred by the drag reduction obtained being negated by the additional drag due to the power requirements, for example the pumping power in the case of mass transfer and the drag of the devices in the case of vortex generators. A system of piezoelectrically controlled flaps is presented for the control of the interaction. The flaps would aeroelastically deflect due to the pressure difference created by the pressure rise across the shock and by piezoelectrically induced strains. The amount of deflection, and hence the mass flow through the plenum chamber, would control the interaction. It is proposed that the flaps will delay separation of the boundary layer whilst reducing wave drag and overcome the disadvantages of previous control methods. Active control can be utilised to optimise the effects of the boundary layer shock wave interaction as it would allow the ability to control the position of the control region around the original shock position, mass transfer rate and distribution. A number of design options were considered for the integration of the piezoelectric ceramic into the flap structure. These included the use of unimorphs, bimorphs and polymorphs, with the latter capable of being directly employed as the flap. Unimorphs, with an aluminium substrate, produce less deflection than bimorphs and multimorphs. However, they can withstand and overcome the pressure loads associated with SBLI control. For the current experiments, it was found that near optimal control of the swept and unswept shock wave boundary layer interactions was attained with flap deflections between 1mm and 3mm. However, to obtain the deflection required for optimal performance in a full scale situation, a more powerful piezoelectric actuator material is required than currently available. A theoretical model is developed to predict the effect of unimorph flap deflection on the displacement thickness growth angles, the leading shock angle and the triple point height. It is shown that optimal deflection for SBLI control is a trade-off between reducing the total pressure losses, which is implied with increasing the triple point height, and minimising the frictional losses.

The interaction of a shock wave with a boundary layer is a classic viscous/inviscid interaction problem that occurs over a wide range of high speed aerodynamic flows. For example, on transonic wings, in supersonic air intakes, in propelling nozzles at offdesign conditions and on deflected controls at supersonic/transonic speeds, to name a few. The transonic interaction takes place at Mach numbers typically between 1.1 and 1.5. On an aerofoil, its existence can cause problems that range from a mild increase in section drag to flow separation and buffeting. In the absence of separation the drag increase is predominantly due to wave drag, caused by a rise in entropy through the interaction. The control of the turbulent interaction as applied to a transonic aerofoil is addressed in this thesis. However, the work can equally be applied to the control of interaction for numerous other occurrences where a shock meets a turbulent boundary layer. It is assumed that, for both swept normal shock and unswept normal shock interactions, as long as the Mach number normal to the shock is the same, then the interaction, and therefore its control, should be the same. Numerous schemes have been suggested to control such interaction. However, they have generally been marred by the drag reduction obtained being negated by the additional drag due to the power requirements, for example the pumping power in the case of mass transfer and the drag of the devices in the case of vortex generators. A system of piezoelectrically controlled flaps is presented for the control of the interaction. The flaps would aeroelastically deflect due to the pressure difference created by the pressure rise across the shock and by piezoelectrically induced strains. The amount of deflection, and hence the mass flow through the plenum chamber, would control the interaction. It is proposed that the flaps will delay separation of the boundary layer whilst reducing wave drag and overcome the disadvantages of previous control methods. Active control can be utilised to optimise the effects of the boundary layer shock wave interaction as it would allow the ability to control the position of the control region around the original shock position, mass transfer rate and distribution. A number of design options were considered for the integration of the piezoelectric ceramic into the flap structure. These included the use of unimorphs, bimorphs and polymorphs, with the latter capable of being directly employed as the flap. Unimorphs, with an aluminium substrate, produce less deflection than bimorphs and multimorphs. However, they can withstand and overcome the pressure loads associated with SBLI control. For the current experiments, it was found that near optimal control of the swept and unswept shock wave boundary layer interactions was attained with flap deflections between 1mm and 3mm. However, to obtain the deflection required for optimal performance in a full scale situation, a more powerful piezoelectric actuator material is required than currently available. A theoretical model is developed to predict the effect of unimorph flap deflection on the displacement thickness growth angles, the leading shock angle and the triple point height. It is shown that optimal deflection for SBLI control is a trade-off between reducing the total pressure losses, which is implied with increasing the triple point height, and minimising the frictional losses.