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Aerodynamic devices, such as wings, are used in higher levels of motorsport (Formula-1 etc.) to increase the contact force between the road and tyres (i.e. to generate downforce). This in turn increases the performance envelope of the race car. However the extra downforce increases aerodynamic drag which (apart from when braking) is generally detrimental to lap-times. The drag acts to slow the vehicle, and hinders the effect of available drive power and reduces fuel economy. Wings, in automotive use, are not constrained by the same parameters as aircraft, and thus higher angles of attack can be safely reached, although at a higher cost in drag. Variable geometry aerodynamic devices have been used in many forms of motorsport in the past offering the ability to change the relative values of downforce and drag. These have invariably been banned, generally due to safety reasons. The use of active aerodynamics is currently legal in both Formula SAE (engineering compet ition for university students to design, build and race an open-wheel race car) and production vehicles. A number of passenger car companies are beginning to incorporate active aerodynamic devices in their designs. In this research the effect of ground proximity on the lift, drag and moment coefficients of inverted, two-dimensional aerofoils was investigated. The purpose of the study was to examine the effect ground proximity on aerofoils post stall, in an effort to evaluate the use of active aerodynamics to increase the performance of a race car. The aerofoils were tested at angles of attack ranging from 0° - 135°. The tests were performed at a Reynolds number of 2.16 x 105 based on chord length. Forces were calculated via the use of pressure taps along the centreline of the aerofoils. The RMIT Industrial Wind Tunnel (IWT) was used for the testing. Normally 3m wide and 2m high, an extra contraction was installed and the section was reduced to form a width of 295mm. The wing was mounted between walls to simulate 2-D flow. The IWT was chosen as it would allow enough height to reduce blockage effect caused by the aerofoils when at high angles of incidence. The walls of the tunnel were pressure tapped to allow monitoring of the pressure gradient along the tunnel. The results show a delay in the stall of the aerofoils tested with reduced ground clearance. Two of the aerofoils tested showed a decrease in Cl with decreasing ground clearance; the third showed an increase. The Cd of the aerofoils post-stall decreased with reduced ground clearance. Decreasing ground clearance was found to reduce pitch moment variation of the aerofoils with varied angle of attack. The results were used in a simulation of a typical Formula SAE race car.

Vortices are present in many fluid flows and depending on the context they may be either beneficial or harmful for different systems or processes. Planar particle image velocimetry was used to examine the vortex evolution and its decay under different turbulence intensities and vortex circulation. The vortex decayed faster in the presence of high turbulence intensity. Vortex trajectories were impacted by turbulence intensity and vortex strength. Trajectories with no turbulence intensity had less variation. The vortex wandering amplitude decreased with growth of vortex strength. The vortex decay was confined to the core of the vortex, with the tangential velocity at large radial distances from the vortex center being relatively constant in time. The vortex core radius had a greater rate of growth with the low turbulence intensity and lower angle of attack. The amplitude of fluctuation of the core circulation increased for the higher turbulence intensity and weaker vortex.

The quality of the flow over a 75 degree-sweep delta wing was documented for steady angles of attack and during dynamic maneuvers with and without the use of two control surfaces. The three-dimensional velocity field over a delta wing at a steady angle of attack of 38 degrees and Reynolds number of 72,000 was mapped out using laser-Doppler velocimetry over one side of the wing. The three-dimensional streamline and vortex line distributions were visualized. Isosurfaces of vorticity, planar distributions of helicity and all three vorticity components, and the indicator of the stability of the core were studied and compared to see which indicated breakdown first. Visualization of the streamlines and vortex lines near the core of the vortex indicate that the core has a strong inviscid character, and hence Reynolds number independence, upstream of breakdown, with viscous effects becoming more important downstream of the breakdown location. The effect of cavity flaps on the flow over a delta wing was documented for steady angles of attack in the range 28 degrees to 42 degrees by flow visualization and surface pressure measurements at a Reynolds number of 470,000 and 1,000,000, respectfully. It was found that the cavity flaps postpone the occurrence of vortex breakdown to higher angles of attack than can be realized by the basic delta wing. The effect of continuously deployed cavity flaps during a dynamic pitch-up maneuver of a delta wing on the surface pressure distribution were recorded for a reduced frequency of 0.0089 and a Reynolds number of 1,300,000. The effect of deploying a set of cavity flaps <u>during</u> a dynamic pitch-up maneuver on the surface pressure distribution was recorded for a reduced frequency of 0.0089 and a Reynolds number of 1,300,000 and 187,000. The active deployment of the cavity flaps was shown to have a short-lived beneficial effect on the surface pressure distribution. The effect on the surface pressure distribution of the varying the reduced frequency at constant Reynolds number for a plain delta wing was documented in the reduced frequency range of 0.0089 to 0.0267. The effect of the active deployment of an apex flap <u>during</u> a pitch-up maneuver on the surface pressure distribution at Reynolds numbers of 532,000, 1,000,000, and 1,390,000 were documented with reduced frequencies of 0.0053 to 0.0114 with flap deployment locations in the range of 21° to 36° . The apex flap deployment was found to have a beneficial effect on the surface pressure distribution during the maneuver and in the post-stall regime after the maneuver is completed. / Ph. D.

No description available.

No description available.

This thesis deals with Computational Fluid Dynamics (CFD)simulations of the flow around delta wings at high angles ofattack. These triangular wings, mainly used in militaryaircraft designs, experience the formation of two vortices ontheir lee-side at large angles of attack. The simulation ofthis vortical flow by solving the Navier-Stokes equations isthe subject of this thesis. The purpose of the work is toimprove the understanding of this flow and contribute to thedesign of such a wing by developing methods that enable moreaccurate and efficient CFD simulations. Simulations of the formation, burst and disappearance of thevortices while the angle of attack is changing are presented.The structured flow solver NSMB has been used to get thetime-dependent solutions of the flow. Both viscous and inviscidresults of a 70°-swept delta wing pitching in anoscillatory motion are reported. The creation of the dynamiclift and the hysteresis observed in the history of theaerodynamic forces are well reproduced. The second part of the thesis is focusing on automatic meshrefinement and its influence on simulations of the delta wingleading-edge vortices. All the simulations to assess the gridquality are inviscid computations performed with theunstructured flow solver EDGE. A first study reports on theeffects of refining thewake of the delta wing. A70°-swept delta wing at a Mach number of 0.2 and an angleof attack of 27° where vortex breakdown is present abovethe wing, is used as testcase. The results show a strongdependence on the refinement, particularly the vortex breakdownposition, which leads to the conclusion that the wake should berefined at least partly. Using this information, a grid for thewing in the wind tunnel is created in order to assess theinfluence of the tunnel walls. Three sensors for automatic meshrefinement of vortical flows are presented. Two are based onflow variables (production of entropy and ratio of totalpressures) while the third one requires an eigenvalue analysisof the tensor of the velocity gradients in order to capture theposition of the vortices in the flow. These three vortexsensors are successfully used for the simulation of the same70° delta wing at an angle of attack of 20°. Acomparison of the sensors reveals the more local property ofthe third one based on the eigenvalue analysis. This lattertechnique is applied to the simulation of the wake of a deltawing at an angle of attack of 20°. The simulations on ahighly refined mesh show that the vortex sheet shed from thetrailing-edge rolls up into a vortex that interacts with theleading-edge vortex. Finally the vortex-detection technique isused to refine the grid around a Saab Aerosystems UnmannedCombat Air Vehicle (UCAV) configuration and its flight dynamicscharacteristics are investigated. Key words:delta wing, high angle of attack, vortex,pitching, mesh refinement, UCAV, vortex sensor, tensor ofvelocity gradients.

The spatial stability of hypersonic conical boundary layers is investigated utilizing different numerical techniques. First, the development and verification of a Linearized Compressible Navier-Stokes solver (LinCS) is presented, followed by an investigation of different effects that affect the stability of the flow in free-flight/ground tests, such as: high-enthalpy effects, wall-temperature ratio, and three-dimensionality (i.e. angle-of-attack). A temporally/spatially high-order of accuracy parallelized Linearized Compressible Navier-Stokes solver in disturbance formulation was developed, verified and employed in stability investigations. Herein, the solver was applied and verified against LST, PSE and DNS, for different hypersonic boundary-layer flows over several geometries (e.g. flat plate - M=5.35 & 10; straight cone - M=5.32, 6 & 7.95; flared cone - M=6; straight cone at AoA = 6 deg - M=6). The stability of a high-enthalpy flow was investigated utilizing LST, LinCS and DNS of the experiments performed for a 5 deg sharp cone in the T5 tunnel at Caltech. The results from axisymmetric and 3D wave-packet investigations in the linear, weakly, and strongly nonlinear regimes using DNS are presented. High-order spectral analysis was employed in order to elucidate the presence of nonlinear couplings, and the fundamental breakdown of second mode waves was investigated using parametric studies. The three-dimensionality of the flow over the Purdue 7 deg sharp cone at M=6 and AoA =6 deg was also investigated. The development of the crossflow instability was investigated utilizing suction/blowing at the wall in the LinCS/DNS framework. Results show good agreement with previous computational investigations, and that the proper basic flow computation/formation of the vortices is very sensitive to grid resolution.

Desde o início da história da aviação, a cada novo projeto de aeronave, modificações são realizadas para se obter melhorias nas características aerodinâmicas como, por exemplo, diminuição do arrasto, aumento de sustentação, melhora na manobrabilidade em elevados ângulos de ataque, entre outras. Essas modificações podem ser feitas em diversas partes de uma aeronave como, por exemplo, no perfil aerodinâmico e enflechamento da asa, nas superfícies de controle e até em novas posições da entrada de ar do motor intake. Embora os conceitos de aeronaves com asas de enflechamento negativo, Canard e intake dorsal não sejam novos, ainda não existe uma aeronave que apresente essas três características juntas. Portanto, o presente trabalho teve como objetivo estudar uma aeronave não convencional com essas características. Para isso, testes foram realizados no túnel de vento do Laboratório de Aerodinâmica (LAE), do Departamento de Engenharia Aeronáutica da Escola de Engenharia de São Carlos, da Universidade de São Paulo. Três estudos foram realizados para diferentes ângulos de ataque da aeronave e diferentes ângulos de incidência de três pares de Canard projetados, que foram colocados em duas posições diferentes em relação à asa. No primeiro estudo, as forças aerodinâmicas atuantes na aeronave foram medidas por meio de uma balança aerodinâmica para se obter relações entre seus coeficientes e a eficiência conforme a troca dos pares de Canard. Posteriormente, o estudo do escoamento na entrada do intake dorsal englobou a relação entre a velocidade local no intake, a velocidade do escoamento livre e a recuperação de pressão total no mesmo. Por fim, um mapeamento da vorticidade no campo da asa foi realizado para se analisar a influência do Canard no escoamento raiz da asa. O modelo estudado obteve uma maior eficiência e menor arrasto em elevados ângulos de ataque para ângulos de incidência negativos de Canard. O Canard que se destacou nos dois primeiros estudos foi utilizado no estudo de mapeamento, porém, não apresentou uma influência no escoamento da raiz da asa como desejado. / Since the beginning of aviation history, in each new aircraft design, modifications are made to obtain improvements in aerodynamic characteristics such as drag reduction, increase of lift, improvement in maneuverability at high angles of attack. These modifications can be made in different parts of an aircraft, such as, aerodynamic profile, on the control surfaces and even in new intake positions. Although the concepts of aircraft with forward swept wing, Canard and dorsal intake are not new, there is still no aircraft that presents these three characteristics together. Therefore, the present work had as objective to study an unconventional aircraft with these characteristics. Wind tunnel tests were carried out in order to analyze the aerodynamic characteristics at the Aerodynamics Laboratory of the Department of Aeronautical Engineering of EESC-USP. Three studies were carried out for different angles of attack of the aircraft and different angles of incidence of three pairs of Canard, which were placed in two different positions in relation to the wing. In the first study, the aerodynamic forces acting on the aircraft were measured by an aerodynamic balance to obtain relations between their coefficients and the efficiency according to the change of the Canard pairs. Subsequently, the study of flow at the entrance of the dorsal intake included the relation between the local velocity at the intake and the velocity of the free flow and the total pressure recovery. Finally, a mapping of vorticity in the wing field was performed to analyze the influence of Canard on the boundary layer at the root of the wing. The model studied obtained higher efficiency and lower drag at high angles of attack at negative Canard angles of incidence. The Canard that stood out in the first two studies was used in the mapping study, however, it did not present an influence on the root of the wing as desired.

<p>This thesis deals with Computational Fluid Dynamics (CFD)simulations of the flow around delta wings at high angles ofattack. These triangular wings, mainly used in militaryaircraft designs, experience the formation of two vortices ontheir lee-side at large angles of attack. The simulation ofthis vortical flow by solving the Navier-Stokes equations isthe subject of this thesis. The purpose of the work is toimprove the understanding of this flow and contribute to thedesign of such a wing by developing methods that enable moreaccurate and efficient CFD simulations.</p><p>Simulations of the formation, burst and disappearance of thevortices while the angle of attack is changing are presented.The structured flow solver NSMB has been used to get thetime-dependent solutions of the flow. Both viscous and inviscidresults of a 70°-swept delta wing pitching in anoscillatory motion are reported. The creation of the dynamiclift and the hysteresis observed in the history of theaerodynamic forces are well reproduced.</p><p>The second part of the thesis is focusing on automatic meshrefinement and its influence on simulations of the delta wingleading-edge vortices. All the simulations to assess the gridquality are inviscid computations performed with theunstructured flow solver EDGE. A first study reports on theeffects of refining thewake of the delta wing. A70°-swept delta wing at a Mach number of 0.2 and an angleof attack of 27° where vortex breakdown is present abovethe wing, is used as testcase. The results show a strongdependence on the refinement, particularly the vortex breakdownposition, which leads to the conclusion that the wake should berefined at least partly. Using this information, a grid for thewing in the wind tunnel is created in order to assess theinfluence of the tunnel walls. Three sensors for automatic meshrefinement of vortical flows are presented. Two are based onflow variables (production of entropy and ratio of totalpressures) while the third one requires an eigenvalue analysisof the tensor of the velocity gradients in order to capture theposition of the vortices in the flow. These three vortexsensors are successfully used for the simulation of the same70° delta wing at an angle of attack of 20°. Acomparison of the sensors reveals the more local property ofthe third one based on the eigenvalue analysis. This lattertechnique is applied to the simulation of the wake of a deltawing at an angle of attack of 20°. The simulations on ahighly refined mesh show that the vortex sheet shed from thetrailing-edge rolls up into a vortex that interacts with theleading-edge vortex. Finally the vortex-detection technique isused to refine the grid around a Saab Aerosystems UnmannedCombat Air Vehicle (UCAV) configuration and its flight dynamicscharacteristics are investigated.</p><p><b>Key words:</b>delta wing, high angle of attack, vortex,pitching, mesh refinement, UCAV, vortex sensor, tensor ofvelocity gradients.</p>