The short range dispersion in the atmosphere of buoyant gases, such as hot air or natural gas, may be hazardous and dangerous. The available methods for studying this problem were reviewed. Wind tunnel studies were considered to be the most suitable method for studying near-field dispersion, and methods for accurately modelling the nearfield behaviour of a buoyant plume of gas were examined. The experiments were performed in the Oxford University 4m x 2m low speed wind tunnel at a model scale of 1:200. The mean trajectory and rate of spread of a buoyant plume from a 60 m high (full-scale) stack were measured in the presence of a simulated natural wind. The exact similarity requirements were derived from dimensional analysis and from the equations of motion. In practice, it is not possible to match all the necessary dimensionless groups and exact scaling of the exit gas density ratio and the exit Reynolds number is often relaxed. A series of experiments was performed to examine the effect of these two groups on mean plume behaviour, with the intention of providing guidance for correct simulation of plume dispersion at reduced-scale. The exit density ratio was found to have little effect on the near-field plume behaviour, provided all the other dimensionless groups were matched. Plumes with low Reynolds number were found to rise significantly higher than plumes with higher 'turbulent' Reynolds numbers. This difference in trajectory could not be correlated with the plume exit momentum flux. The effect of the cross-flow on near-field dispersion was examined by performing experiments in four different simulations of the earth's atmospheric boundary-layer. The behaviour of the plume was found to be sensitive to both the velocity profile and the turbulence intensity of the cross-flow. To study dispersion in the wind tunnel, the cross-flow should be an accurate simulation of the velocity profile and turbulence intensity components of the natural wind.

The aerodynamics of bodies in compressible ground effect flowfields from low-subsonic to supersonic Mach numbers have been investigated numerically and experimentally. A study of existing literature indicated that compressible ground effect has been addressed sporadically in various contexts, without being researched in any comprehensive detail. One of the reasons for this is the difficulty involved in performing experiments which accurately simulate the flows in question with regards to ground boundary conditions. To maximise the relevance of the research to appropriate real-world scenarios, multiple bodies were examined within the confines of their own specific flow regimes. These were: an inverted T026 wing in the low-to-medium subsonic regime, a lifting RAE 2822 aerofoil and ONERA M6 wing in the transonic regime, and a NATO military projectile at supersonic Mach numbers. Two primary aims were pursued. Firstly, experimental issues surrounding compressible ground effect flows were addressed. Potential problems were found in the practice of matching incompressible Computational Fluid Dynamics (CFD) simulations to wind tunnel experiments for the inverted wing at low freestream Mach numbers (<0.3), where the inverted wing was found to experience significant compressible effects even at Mach 0.15. The approach of matching full-scale CFD simulations to scale model testing at an identical Reynolds number but higher Mach number was analysed and found to be prone to significant error. An exploration was also conducted of appropriate ways to conduct experimental tests at transonic and supersonic Mach numbers, resulting in the recommendation of a symmetry (image) method as an effective means of approximating a moving ground boundary in a small-scale blowdown wind tunnel. Issues of scale with regards to Reynolds number persisted in the transonic regime, but with careful use of CFD as a complement to experiments, discrepancies were quantified with confidence. The second primary aim was to use CFD to gain a broader understanding of the ways in which density changes in the flowfield affect the aerodynamic performance of the bodies in question, in particular when a shock wave reflects from the ground plane to interact again with the body or its wake. The numerical approach was extensively verified and validated against existing and new experimental data. The lifting aerofoil and wing were investigated over a range of mid-to-high subsonic Mach numbers (1>M???>0.5), ground clearances and angles of incidence. The presence of the ground was found to affect the critical Mach number, and the aerodynamic characteristics of the bodies across all Mach numbers and clearances proved to be highly sensitive to ground proximity, with a step change in any variable often causing a considerable change to the lift, moment and drag coefficients. At the lowest ground clearances in both two and three dimensional studies, the aerodynamic efficiency was generally found to be less than that of unbounded (no ground) flight for shock-dominated flowfields at freestream Mach numbers greater than 0.7. In the fully-supersonic regime, where shocks tend to be steady and oblique, a supersonic spinning NATO projectile travelling at Mach 2.4 was simulated at several ground clearances. The shocks produced by the body reflected from the ground plane and interacted with the far wake, the near wake, and/or the body itself depending on the ground clearance. The influence of these wave reflections on the three-dimensional flowfield, and their resultant effects on the aerodynamic coefficients, was determined. The normal and drag forces acting on the projectile increased in exponential fashion once the reflections impinged on the projectile body again one or more times (at a height/diameter ground clearance h/d<1). The pitching moment of the projectile changed sign as ground clearance was reduced, adding to the complexity of the trajectory which would ensue.

Football is widely regarded as the most popular sport in the world involving over 270 million people from different countries and cultures. It can be argued that the football is one of most important aspects of the game and hence the flight of the ball, if unexpected, can alter the outcome of the game. This thesis provides an engineering perspective and contribution to the continued understanding and improvement of the in-flight performance of FIFA approved footballs. Skilful players will impart spin onto a ball to induce a curve in-flight to try and deceive opponents. This flight is generally smooth, although subtle variations in the orientation and spin rate may cause conditions that affect the path and final ball position, in a manner considered to be unpredictable due to aerodynamic effects. Ball designs and manufacturing techniques are evolving and certain seam configurations are known to induce asymmetric pressure distributions resulting in lateral movement during flight. Aerodynamic research of sport balls has primarily focused on drag and the effects of high spin rates. Studies have shown the introduction of surface roughness affects the boundary layer state compared to a smooth sphere. Surface roughness on a football takes many forms including seam configurations and micro surface textures. The influence of changing the density, distribution and dimensions of the surface roughness with respect to the aerodynamic behaviour has been researched. The principle focus of this thesis is concerned with the influence on the lateral component as a result of applying surface roughness to the outer surfaces. The influence of the surface roughness on the drag and lateral components were determined using established wind tunnel techniques. Real balls and full size prototypes were tested. A mathematical flight model was employed to simulate realistic multiple flight trajectories based on empirical aerodynamic data. Mathematical and statistical techniques, including R.M.S and AutoCorrelation Functions were used to analyse the data. The results from this research showed how small variations in surface texture affected the complex nature of the lateral forces. Trajectories varied significantly depending on initial orientation and slow spin rate sensitivities. In conclusion, ball characterisation techniques were developed that identified lateral deviation and shape measures and considered a gradient profiling approach. Application of these novel parameters through multiple trajectory analysis allowed for an in-flight performance measure of footballs designs.

Como o erro é inerente ao próprio processo de medida, e portanto nunca será completamente eliminado, deve-se buscar minimizá-lo, reduzindo ao máximo as fontes de erros grosseiros e sistemáticos. Devido à dificuldade de identificar com segurança as fontes desses erros e tendo o indício que, em grande parte, ocorrem devido às variações geométricas e às características construtivas dos modelos ensaiados, é apresentada, como proposta de ferramenta de auxílio à detecção desses erros, a construção de um modelo padrão aeronáutico. Devido às semelhanças geométricas com os modelos das aeronaves atualmente ensaiados no Túnel Aerodinâmico número dois (TA-2) do IAE, foi escolhido como modelo-padrão o M5 da série M do Office National DÉtudes et de Recherches Aérospatiales (ONERA) para ensaios de proficiência intra e interlaboratorial do TA-2 nesse tipo de ensaio. O contexto que possibilitou a recomendação da proposta está descrito na primeira parte do presente trabalho, que trata do controle e melhoria da infraestrutura. Quanto à infraestrutura, propõe-se a melhoria das instalações e o conhecimento e controle sobre o escoamento do túnel. Sobre controle na execução dos ensaios intralaboratoriais são obedecidos os quesitos da NBR 17025 e da NBR 15100. Esses ensaios intralaboratoriais fornecem a reprodutividade dos ensaios de curto termo, isto é, ensaios executados numa mesma campanha ou próximos que forneceram resultados comparáveis aos de grandes complexos de túneis do primeiro mundo, porém a reprodutividade de longo termo, quando observadas condições diferentes de equipamentos, montagem, entre outros, nota-se uma característica de erros sistemáticos. / As the error is inherent to the proper process of measure, and therefore never completely it will be eliminated, it must be searched to minimize it, reducing to the maximum the sources of gross errors and systematic. Due to difficulty to identify with security the sources of these errors and having the indication that, to a large extent, occurs due to the geometric variations and the constructive characteristics of the assayed models, it is presented, as proposal of tool of aid to the detention of these errors, the construction of a model aeronautical standard. Had to the geometric similarities with the models of the aircraft currently assayed in the Wind tunnel number two (TA-2) of the IAE, the M5 of M serie of the Office National DÉtudes et de Recherches Aérospatiales (ONERA) it was chosen as model-standard for proficiency assays intra and interlaboratorial of the TA-2 in this type of assay. The context that made possible the recommendation of the proposal is described in the first part of the present work, that deals with the control and improvement of the infrastructure. About to the infrastructure, it is considered improvement of the installations and the knowledge and control on the draining of the tunnel. On control in the execution of the intralaboratoriais assays the questions of NBR 17025 and NBR 15100 are obeyed. These intralaboratoriais assays supply the reprodutividade of the assays of short term, that is, assays executed in one same campaign or next that they had supplied resulted comparable to the ones of great complexes of tunnels of the first world, however the reprodutividade of long term, when observed different conditions of equipment, assembly, among others, a characteristic of bias quality controls is noticed.

D.Ing. (Mechanical Engineering) / The most common data sought during a wind tunnel test program are the forces and moments acting on an airframe, (or any other test article). The most common source of this data is the internal strain gauge balance. Balances are six degree of freedom force transducers that are required to be of small size and of high strength and stiffness. They are required to deliver the highest possible levels of accuracy and reliability. There is a focus in both the USA and in Europe to improve the performance of balances through collaborative research. This effort is aimed at materials, design, sensors, electronics calibration systems and calibration analysis methods. Recent developments in the use of statistical methods, including modern design of experiments, have resulted in improved balance calibration models. Research focus on the calibration of six component balances has moved to the determination of the uncertainty of measurements obtained in the wind tunnel. The application of conventional statistically-based approaches to the determination of the uncertainty of a balance measurement is proving problematical, and to some extent an impasse has been reached. The impasse is caused by the rapid expansion of the problem size when standard uncertainty determination approaches are used in a six-degree of freedom system that includes multiple least squares regression and iterative matrix solutions. This thesis describes how the uncertainty of loads reported by a six component balance can be obtained by applying a direct simulation of the end-to-end data flow of a balance, from calibration through to installation, using a Monte Carlo Simulation. It is postulated that knowledge of the error propagated into the test environment through the balance will influence the choice of calibration model, and that an improved model, compared to that determined by statistical methods without this knowledge, will be obtained. Statistical approaches to the determination of a balance calibration model are driven by obtaining the best curve-fit statistics possible. This is done by adding as many coefficients to the modelling polynomial as can be statistically defended. This thesis shows that the propagated error will significantly influence the choice of polynomial coefficients. In order to do this a Performance Weighted Efficiency (PWE) parameter is defined. The PWE is a combination of the curve-fit statistic, (the back calculated error for the chosen polynomial), a value representing the overall prediction interval for the model(CI_rand), and a value representing the overall total propagated uncertainty of loads reported by the installed balance...

Much research and experiments have gone into studying idealized wing-body junction flows and their impact on horseshoe vortex and wake formation.  The vortices have been found to generate regions of high surface pressure fluctuations and turbulence that are detrimental to structural components and acoustics.  With the focus in the military and commercial industry on reducing the acoustical impact of aircraft and their engines, very little research has been done to examine the potential impact wing-body junctions may have on acoustics, especially for high lifting bodies such as propellers.  Two similar tests were conducted in the Virginia Tech Open Jet Wind Tunnel where boundary layer measurements, oil flow visualizations, acoustic linear array and surface pressure fluctuation measurements of a baseline Rood airfoil model and two novel junction fairing designs were all taken.  Boundary layer measurements were taken at four locations along the front half of the flat plate and the profiles were shown to be all turbulent despite the low Reynolds number of the flow, (test 1: Re_"<1400, test 2: Re_"<550).  Oil flow visualizations were taken and compared to those of previous researchers and the location of separation and line of low shear along with the maximum width of the wake and width of wake at the trailing edge all scaled relatively well with the Momentum Deficit Factor, defined for wing-body junction flows [Fleming, J. L., Simpson, R. L., Cowling, J. E. & Devenport, W. J., 1993. An Experimental Study of a Turbulent Wing-Body Junction and Wake Flow. Experiments in Fluids, Volume 14, pp. 366-378. ].  A linear microphone array was used to estimate the directivity of the facility acoustic background noise to be used to improve background subtraction methods for surface pressure fluctuation measurements.  Surface pressure fluctuation spectra were taken ahead of the leading edge of the plate and along the surface of the models.  These showed that the fairings reduced pressure fluctuations along the plate upstream of the leading edge, with fairing 1 reducing them to clean tunnel flow levels.  On the surface of the models, the fairings tended to reduce low frequency (<1000Hz) pressure fluctuation peaks when compared to the baseline model and increase the pressure fluctuations in the high frequency range.  Simple scaling arguments indicate that this spectral change may be more beneficial than detrimental as low frequency acoustics especially those between 800 Hz and 1200 Hz are the frequencies that humans perceive as the loudest noise levels.  Scaling the frequencies measured to those of full scale applications using Strouhal numbers show that frequencies below 1000 Hz in this experiment result in frequencies at the upper limit of the human hearing frequency range.  Low frequency acoustic waves also tend to travel farther and high frequency acoustic waves are more apt to be absorbed by the surrounding atmosphere. / Master of Science

This project examined the effect of Gurney flaps on a 2D, 2-ft chord NACA 0036 airfoil in the Cal Poly 3’x4’ low speed wind tunnel at 25 m/s. It also covered the numerical simulation of the experiment in computational fluid dynamics (CFD). During the study, problems with the wind tunnel data were seen. After a careful diagnosis, the problem was traced to dirty flow conditioners which were subsequently replaced. Five Gurney flaps at 1, 2, 3, 4, and 5% of the chord were tested. The Gurney flaps had the effect of eliminating the lift reversal effect and lowering the profile drag at low angles of attack, ranging from 4-27%. The optimal Gurney flap appeared to be 2% of the chord. CFD modeling of the problem had limited success, with the best results coming from Mentor’s k-w SST turbulence model. This model reproduced the non-linear lift curve, and captured the trend in rising drag fairly well, but failed to predict the correct point of separation. Attempts to model the Gurney flap in CFD were unsuccessful.

An aircraft may experience inlight ice accretion and corresponding reductions in performance and control when the vehicle encounters clouds of super-cooled water droplets. In order to study anti-icing coatings, the EADS-IW Surface Engineering Group is building a refrigerated wind tunnel. Several variations of droplet delivery systems were explored to determine the most effective way to introduce mono-dispersed droplets into the wind tunnel. To investigate this flow, timeurate, unsteady viscous flow simulations were performed using the Loci/CHEM flow solver with a multi-scale hybrid RANS/LES turbulence model. A Lagrangian droplet model was employed to simulate the movement of water droplets in the wind tunnel. It was determined that the droplet delivery system required pressure relief to properly orient the flow inside the droplet delivery tube. Additionally, a streamlined drop tube cross-section was demonstrated to reduce turbulence in the wake and decrease the variability in droplet trajectories in the test section.

An experimental study of active control of fully separated flow over a symmetrical circular-arc airfoil at high angles of attack was performed. The experiments were carried out in a low-speed, open circuit wind tunnel. Angles of attack from 10 to 40 degrees were tested. Low-power input, unsteady excitation was applied to the leading or trailing edge shear layers. The actuation was provided by the periodic oscillation of a 4-percent-chord flap placed on the suction side of the airfoil and facing the sharp edge. Vortex-shedding frequencies were measured and harmonic combinations selected as the applied actuator frequencies. Pressure measurements over the airfoil show that the control increased the normal force coefficient by up to 70%. This supports the idea of vortex capture in the time-averaged sense, enhancing the lift on the airfoil by managing the shear layer roll up. The results indicate the viability of the control of large-scale flow fields by exploiting the natural amplification of disturbances triggered by small-scale actuators. The application of flow control on sharp-edged aircraft wings could lead to improved maneuverability, innovative flight control and weight reduction. These can be achieved by inexpensive, low-power, rugged actuators. / Master of Science

This thesis deals with the problems of scaling aerodynamic data from wind tunnel conditions to free flight. The main challenges when this scaling should be performed is how the model support, wall interference and the potentially lower Reynolds number in the wind tunnel should be corrected. Computational Fluid Dynamics (CFD) simulations have been performed on a modern transonic transport aircraft in order to reveal Reynolds number effects and how these should be scaled accurately. This investigation also examined how the European Transonic Wind tunnel (ETW) twin sting model support influences the flow over the aircraft. In order to further examine Reynolds number effects a MATLAB based code capable of extracting local boundary layer properties from structured and unstructured CFD calculations have been developed and validated against wind tunnel measurements. A general scaling methodology is presented. / QC 20101123