Dissertations / Theses on the topic 'Aerofoil'
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1
Kingan, Michael Joseph. "Aeroacoustic noise produced by an aerofoil." Thesis, University of Canterbury. Mechanical Engineering, 2005. http://hdl.handle.net/10092/6596.
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This thesis describes an investigation into the aeroacoustic noise produced by an aerofoil
using experimental, computational and theoretical methods. Several different types of
aeroacoustic noise generation mechanisms and the parameters which affect these
mechanisms were identified and investigated.
The aerofoils used' in this investigation all had chord lengths of 100mm and had a
maximum thickness between 18mm and 30mm. Experimental testing was undertaken in
the low noise wind tunnel in the Department of Mechanical Engineering at the University
of Canterbury with the aerofoils mounted at the exit of the tunnel. Airflow speeds from
10m/s to 40m/s and a range of angles of incidence were investigated. A number of
modifications were made to reduce the noise and improve the operation of the wind
tunnel. Different methods of measuring the aeroacoustic noise produced by an aerofoil
were also investigated.
The theory of aeroacoustic noise generation is described and the effect of a scattering
surface on the efficiency of these aeroacoustic noise sources was investigated. A number
of different mechanisms by which an aerofoil produces aeroacoustic noise were
identified. These mechanisms were divided into three main categories: (1) blunt trailing
edge aerofoil noise (2) sharp trailing edge aerofoil noise and (3) stalled aerofoil noise.
The effect of air temperature on the production of aeroacoustic noise was also
investigated. It was found that in most instances air temperature would have little effect
on aeroacoustic noise generation.
An extensive study of the aeroacoustic noise produced by a number of different aerofoils
was undertaken. Modelling of the airflow over the aerofoils was used to determine the
mechanism by which aeroacoustic noise is produced. Several different aeroacoustic noise
generation mechanisms were identified. Theoretical models were also used to model the
aeroacoustic noise produced by the aerofoils.
Several treatments to reduce the level of aeroacoustic noise produced by an aerofoil were
investigated. The treatments reduced the aeroacoustic noise produced by an aerofoil with
varying degrees of success.
A method for measuring the aeroacoustic noise produced by car roof racks mounted on
the roof of a vehicle using a relatively small wind tunnel was established. The noise level
produced by a roof rack installed 011 the roof of a vehicle measured using this technique
compared favourably with measurements made on a full vehicle in a large wind tunnel.
The method shows promise as a low cost method of accurately measuring the
aeroacoustic noise produced by roof racks installed on a vehicle roof.
2
Andrew, David Neil. "Flow in regenerative compressors with aerofoil blading." Thesis, University of Cambridge, 1987. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.235767.
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Regenerative compressors are low specific speed, tangential flow turbomachines in which the working fluid follows an approximately toroidal-helical path around the machine. Although regenerative turbomachines have been in use for over fifty years, the use of aerofoil blades (as opposed to the more usual straight vanes) is a relatively new innovation designed to improve the internal flow. However, little work has been done on the details of this flow. This dissertation concentrates on the flow in the main region of the compressor, away from the inlet and exit ports. A simple, analytical model of the flow is developed; and it is shown that, under certain conditions, a similar flow pattern can be set up in a stationary model. Measurements of the flow in a machine are reported and a comparison is made with results from such a stationary model. It is concluded that, although discrepancies may arise from wall friction, Coriolis and centrifugal effects, or the variation in density, the stationary model is nevertheless capable of providing useful information about the loss-producing mechanisms in the flow. The results of more detailed measurements on stationary models of two different designs are presented, on the basis of which suggestions are made for an improved design of machine. The analysis is extended to include Coriolis and centrifugal effects, and the influence which the position of the blades relative to the machine axis has on the performance of the machine is examined. The presence of a uniform component of vorticity normal to the blade-to-blade plane is a characteristic feature of regenerative turbomachines. Some effects of this are discussed, and a computer program written to calculate the two-dimensional, inviscid, incompressible flow through a cascade with such a uniform spanwise vorticity is described.
3
Khoo, Hilary. "Separated flow past wind turbine aerofoil sections." Thesis, Imperial College London, 1991. http://hdl.handle.net/10044/1/46866.
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Paruchuri, Chaitanya. "Aerofoil geometry effects on turbulence interaction noise." Thesis, University of Southampton, 2017. https://eprints.soton.ac.uk/415884/.
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Fan broadband is one of the dominant noise sources on an aircraft engine, particularly at approach. The dominant noise generation mechanism is due to turbulent- aerofoil interaction noise (TAI). This thesis investigates the effect of changes in 2D aerofoil geometry on TAI noise. The main focus of this thesis is to attempt to reduce it through the development of innovative leading edge geometries. The first two chapters of the thesis deals with an experimental and numerical investigation into the effect of aerofoil geometry on interaction noise on single aerofoils and on cascades. Consistent with previous work, they show that variations in aerofoil parameters, such as aerofoil thickness, leading edge nose radius and camber, produce only a small changes in broadband interaction noise at approach conditions. Subsequent chapters deal with the development of innovative leading edge serration profiles aimed at reducing interaction noise. Chapter 4 is a detailed study into the limitations of single-wavelength serrations in reducing interaction noise. The optimum profile is identified. Chapters 5, 6 and 7 all deal with the development of innovative profiles that can provide up to 10dB of additional noise reductions compared to single-wavelength serrations. For each of the profiles investigated a simple model is developed to aid the understanding of their interaction mechanism.
5
Crompton, Matthew John. "The thin aerofoil leading edge separation bubble." Thesis, University of Bristol, 2001. http://hdl.handle.net/1983/25312c88-4d89-4149-bee9-d56cf80d9735.
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6
Chew, Siou Chye. "Numerical simulations of oscillatory flapping aerofoil propulsion." Thesis, Imperial College London, 2011. http://hdl.handle.net/10044/1/8137.
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The primary objective of the current research is to develop a Computational Fluid Dynamics (CFD) model to investigate rigid and flexible aerofoil propulsive characteristics when the aerofoil IS subjected to oscillatory flapping motions. The study is also extended to rectangular wings. Flows past the flapping aero foils at moderate Reynolds numbers are simulated using the twodimensional incompressible Navier-Stokes equations. The Baldwin-Lomax algebraic turbulence model is incorporated to determine eddy viscosity for higher Reynolds numbers turbulent flow simulations. Flows past flapping wings are simulated using strip theory, which computes the flows in multiple two-dimensional planes located at intervals along the wing span. The flows are assumed locally. two-dimensional and the three-dimensional effects between each section are incorporated via the consideration of vortex lattice effects. The simulations are modelled using piecewise linear finite element approximation method on an unstructured triangular finite element mesh. A dynamic moving mesh is used to compute flexible aerofoils and wings. The mesh is remeshed at each fluid time step using the spring segment analogy method. A novel treatment of the near.-wall viscous grids ensures that the good orthogonal properties are maintained to facilitate the turbulence computations. A wide range of simulations is carried out for an oscillatory heaving NACA0012 aerofoil. Parametric studies of basic parameters like the amplitude of oscillation, its reduced frequency, and the flow freestream Reynolds numbers effects on aero foil performance are conducted. The influences of the flexural profile on the flexible aerofoil propulsive characteristics are also investigated. The rectangular wing, oflow aspect ratio 4 and NACA0012 aerofoil cross-sections, is also simulated in oscillatory heaving motion. The chordwise flexural effects of the heaving flexible wing on its propulsive characteristics are studied too.
7
Hustad, C. W. "The drag of a circulation controlled aerofoil." Thesis, University of Bath, 1986. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.370668.
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8
Phillips, Russell Leslie. "Development of a reciprocating aerofoil wind energy harvester." Thesis, Nelson Mandela Metropolitan University, 2008. http://hdl.handle.net/10948/899.
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Cross flow wind turbines are not unique. The performance of Savonius and Darrieus turbines is well documented. Both share the advantage of being able to accept fluid flow from any direction. The Savonius is drag based and hence has poor power output while the Darrieus is lift based. Due to the fact that the Darrieus has fixed blades the fluid flow through the rotor does not result in optimal lift being generated at all points in the rotation circle. A drawback of the Darrieus system is that it has to operate at a high tip-to wind-speed ratio to obtain reasonable performance with the fixed blades. Deviation from a small optimal range of tip speed ratios results in poor performance. The Darrieus also has poor starting torque. The research conducted in this project focused on overcoming the shortcomings of other turbines and developing an effective cross flow turbine capable of good performance. A number of different concepts were experimented with, however all were based on a symmetrical aerofoil presented to the actual relative airflow at an angle that would produce the highest lift force at all times. The lift force was then utilized to generate movement and to do work on an electrical generator. All concepts contemplated were researched to ascertain their appropriateness for the intended application. During development of the final experimental platform and after lodging of a provisional patent (RSA 2007/00927) it was ascertained that the design shared some similarities with an American patent 5503525 dated 28/4/1994. This patent employed complex electronic sensing and control equipment for control of blade angle. This was thought to be overly complex and costly, particularly for small scale wind energy generation applications and a simpler mechanical solution was sought in the design of the final experimental platform used in this project. The design of the mechanical control system was refined in an attempt to make it simpler, more durable and employ the least number of moving parts. Literature studies and patent searches conducted, suggested that the mechanical control system as developed for the final experimental platform was unique. The enormous variation in the power available from the wind at the different wind speeds likely to be encountered by the device necessitated some means of control. In high wind conditions control of the amount of wind power into the device was deemed to be the preferable means of control. A number of different concepts to achieve this were devised and tested. The final concept employed limited the tail angle deflection and hence the lift produced by the aerofoils. This resulted in a seamless “throttle” control allowing the device to be used in any wind strength by adjusting the control to a position that resulted in the device receiving a suitable amount of power from the wind. The outcome of performance tests conducted indicated that the device has the potential to be developed into a viable wind turbine for both small and large scale applications. The ability to control the power input from the wind to the machine from zero to a maximum is considered to be one of the most beneficial outcomes of this project and together with the quiet operation and low speed, are considered the main advantages of the device over existing wind turbine designs. The possibilities of using the device to compress air for energy storage are exciting avenues that warrant further research.
9
Gorman, J. "Vortex formation in the flow around a lifting aerofoil." Thesis, Swansea University, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.637079.
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This thesis presents an investigation of vortex formation in the flow around a lifting aerofoil both with and without acoustic forcing. The chord Reynolds number range covered is from 3000 to 300000. At low chord Reynolds number, (3000 to 20000), and low incidence, (0 to 8°), vortex formation in the flow around an aerofoil has been studied in a water tunnel. Stability theory and in particular the concepts of absolute and convective instability are applied to the aerofoil wake. It is found that for non zero incidence a region of absolutely unstable flow is associated with the separation bubble. Koch's (1985) global mode selection criteria is applied to the wake and is found to give good agreement with the experimentally measured vortex formation frequency. This implies the existence of a critical profile in the wake which, in conjunction with the aerofoil trailing edge, traps energy at the vortex formation frequency. At higher chord Reynolds numbers, (17000 to 300000), the vortex formation characteristics of the aerofoil are further examined both with and without the excitation of an acoustic resonance. Detailed phase averaged measurements of the aerofoil wake are made allowing examination of its structure during resonance. A novel scheme for the accurate measurement of vortex phase is introduced, based on comparison of vortical structure of the vortex formation region using two-dimensional cross-correlation. The results obtained support a modified version of the acoustic resonance excitation mechanism proposed by of Arbey and Bataille (1983). Vortex street structures considerably different to the classic Kármán vortex street are observed. For antisymmetric streets, the vortex spacing ratio h/l is shown to be widely different from the theoretical value of 0.281 obtained by von Kármán and Rubach.
10
Baker, David. "Design of a morphing aerofoil using compliant structure optimisation." Thesis, University of Bristol, 2009. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.529826.
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11
Ho, Pui Yin. "An interactive boundary layer method for subsonic aerofoil flows." Thesis, Imperial College London, 1989. http://hdl.handle.net/10044/1/47475.
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12
Fernie, Robert Mark. "Low frequency shock motion on a NACA 0012 aerofoil." Thesis, University of Cambridge, 2005. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.614936.
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13
Pope, Martin Peter. "Mathematical modelling of unsteady problems in thin aerofoil theory." Thesis, University of Southampton, 1999. https://eprints.soton.ac.uk/192771/.
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The de-icing of aircraft wings by the injection of fluid through a slot in the leading edge of the wing is analysed. A review of current de-icing methods is presented and the semi-infinite slot-injection equation derived, which is a singular partial integro-differential equation. The Stefan condition is used to close the system. A discretisation of the equation is presented and the subsequent numerical results are analysed. The model is then revised to account for the retraction of the ice layer away from the slot. An asymptotic result for the thin ice layers is also presented. The problem of describing the motion of a thin, flexible membrane fixed at both ends (a 'sail') is then considered. The steady sail is analysed for the case of an inextensible sail and previous work on this topic is extended by using a discretisation of the singular integro-differential equation that is pertinent to the later analysis of the unsteady sail. An asymptotic expression for the eigenvalues of the system, defined as the values of the tension parameter for which the sail generates zero lift, is also presented. The problem is then extended to that of an extensible sail and numerical results are presented for both the sail with excess length and the membrane without slack. The case where the angle of incidence of the sail to the free stream is a prescribed function of time is then analysed. Previous work on this subject is extended to include the extensible sail and numerical results are presented. A linear stability analysis is then undertaken for both the extensible and elastic sails; the resulting quadratic eigenvalue problem is solved numerically and is in agreement with the numerical experiments. The trailing edge of the membrane is now permitted to move freely and thus the motion of a 'flag' is analysed. The inclusion of bending stiffness is found to be crucial to the stability properties of the flag. The steady equation of motion is numerically approximated for both a hinged flag and a flag that is clamped at the leading edge. The unsteady flag equation is then discretised and numerical results are presented. A linear stability analysis is performed, the conclusions of which are consistent with the numerical approximations of the unsteady flag equation.
14
Ai, Qing. "Novel morphing structures for aerofoil flow and noise control purposes." Thesis, University of Bristol, 2016. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.720818.
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Lattimer, Timothy Richard Bislig. "Singular partial integro-differential equations arising in thin aerofoil theory." Thesis, University of Southampton, 1996. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.243192.
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16
Lewis, Mark Charles. "Aerofoil testing in a self-streamlining flexible walled wind tunnel." Thesis, University of Southampton, 1987. https://eprints.soton.ac.uk/52285/.
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Morin, Sandrine. "A combined numerical and experimental study of aerofoil separation bubbles." Thesis, University of Southampton, 2003. https://eprints.soton.ac.uk/47094/.
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Ou, Hengan. "Finite element analysis and experimental investigation of stiffness characteristics of forming presses and forging of turbine aerofoil components." Thesis, University of Strathclyde, 2001. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.366750.
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Gibson, Thomas Mark. "The passive control of shock-wave/boundary-layer interactions." Thesis, University of Cambridge, 1997. https://www.repository.cam.ac.uk/handle/1810/272691.
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Cameron, Lee Stewart William. "Adaptive multi-objective aerofoil optimisation : impact of surface contamination and degradation." Thesis, Queen's University Belfast, 2013. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.602451.
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The preliminary design of aerofoils for use in versatile air vehicles is considered, with a view to addressing the sensitivity of the boundary layer transition to surface imperfections arising through surface contamination and degradation. An efficient multi-objective optimisation framework has been developed, enabling optimal trade-offs between key flight conditions to be quantified which include the impact of surface contaminants on the evolution of the design space. The impact of surface contamination and degradation on the design of optimal profiles is assessed through the reduction in the critical amplification factor of boundary layer disturbances at the point of transition onset and the resulting Pareto front compared with comparative 'clean' surface cases. A meta-model assisted global search alleviates the expense of coupling the algorithms with expensive CFD solvers, with an adaptive Kriging -based strategy adopted. The surrogate is iteratively refined in promising regions of the design space according to a probablistic adaptive sampling algorithm. Optimal solutions identified using the global multi-objective search are further improved using a novel local search and reparameterisation scheme. Additional degrees of freedom are iteratively introduced into the parameterisation resulting in an increasingly flexible geometry description. The proposed framework is benchmarked against traditional approaches and found to be superior in both accuracy and efficiency. The requirement that take-off be insensitive to contamination effects proves to be a key consideration and the physical mechanisms by which this robustness in design is achieved are discussed.
21
Andreou, Christodoulos. "Acoustic and aerodynamic properties of aerofoil leading-edge high-lift geometries." Thesis, University of Cambridge, 2009. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.611404.
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Lu, Bin. "3D Die Shape Optimisation For Net Shape Forging of Aerofoil Blades." Thesis, Queen's University Belfast, 2008. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.501609.
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McAlpine, Alan. "Generation of discrete frequency tones by the flow around an aerofoil." Thesis, University of Bristol, 1997. http://hdl.handle.net/1983/cbeb5264-28b8-4d23-b7bf-e869ea55a4a8.
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Tonal noise, the self-induced discrete frequency noise generated by aerofoils, is investigated. It is heard from an aerofoil placed in streams at low Mach number flows when inclined at a small angle to the stream. The tones are heard as a piercing whistle, commonly up to 30 dB above the background noise level. The work is motivated by the occurrence of tonal noise from rotors, fans and recently wind-turbines. Previous authors have attributed tonal noise to a feedback loop consisting of a coupling between laminar boundary-layer instability waves and sound waves propagating in the free stream. The frequency has been predicted by use of various methods based on this model. In this thesis a review of wind-tunnel results obtained by Dr. E.C. Nash at the University of Bristol is presented. Boundary-layer measurements show the presence of tonal noise is closely related to the existence of a region of separated flow close to the trailing edge of the aerofoil. Highly amplified boundary-layer instability waves were observed close to the trailing edge of the aerofoil at the frequency of the tone. A comprehensive analysis of the linear stability of the boundary-layer flow over the aerofoil is presented. The growth of boundary-layer instability waves over the aerofoil is calculated. The growth rates of the waves were obtained by solving the Orr? Sommerfeld problem at several stations on the aerofoil. The Falkner?Skan boundary layers were found to be a suitable form of velocity profiles to incorporate the adverse pressure gradients experienced by the flow over an aerofoil. The amplification of the instability waves is shown to be controlled almost entirely by the region of separated flow close to the trailing edge. The calculated frequency of the linear modes with maximum amplification over the aerofoil is found to be close to the observed frequency of the acoustic tone. A weakly nonlinear stability analysis was also performed and this appears to be a suitable description of the boundary-layer instability waves. The results indicate that the frequency of the tones may commonly be predicted to within 10% by using weakly nonlinear stability theory. The generation of sound by diffraction of the boundary-layer instability waves at the trailing edge of the aerofoil is also discussed as well as the proposed feedback models. A modified feedback model is proposed, being based on the experimental and theoretical results.
24
Huang, X. "Active control of aerodynamic instabilities." Thesis, University of Cambridge, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.237877.
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Cyr, Stéphane. "A theoretical model for flow about a circular-arc aerofoil with separation /." Thesis, McGill University, 1992. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=61186.
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The aim of the present research is to develop a theoretical model that could predict the trailing-edge separation over a circular-arc aerofoil and its effect on the pressure distribution, and on the lift. The model uses sources, in a potential flow, to simulate the effect of the separated region on the pressure distribution. An integral method, turbulent boundary-layer development, combined with the irrotational pressure distribution on the circular-arc predicts the point of separation. To make sure that trailing-edge separation is present alone and mostly to avoid any difficulties in predicting the boundary-layer flow after reattachment, the aerofoil is studied at design incidence.Two different circular-arc aerofoils were tested; one of 10% and one of 18% camber. The calculation for the 10% aerofoil did not predict any separation at design incidence, which was confirmed experimentally. The theoretical model predicted the right position of separation for the 18% aerofoil. It also provided a good simulation of the pressure distribution, including the right value for the base pressure and a good prediction of the lift. It follows that the pressure distribution when integrated, gives a good estimate of the pressure or form drag. (Abstract shortened by UMI.)
26
Fox, M. D. "A study of bond coat cracking in TBCs for turbine aerofoil applications." Thesis, University of Cambridge, 2001. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.599161.
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This dissertation describes work on the thermomechanical fatigue (TMF) behaviour of coated single crystal, nickel-based superalloys. During the course of study, the TMF facility at Cambridge was developed to enhance the reliability of the test technique and to enable the testing of thin ceramic shell TBC test pieces. Benchmark TMF tests were carried out under strain control on a series of different coatings to evaluate the performance of two new coat systems developed at Rolls-Royce that incorporate a platinum surface modification. Cyclic life to failure suggested that the surface modification had minimal effect on TMF performance. The results compared favourably with data from different research centres as well as the lifting model used at Rolls-Royce to predict high temperature coating mechanical integrity (the Tcrit model). From this model, the effect of different TMF cycle directions was considered. Experimental tests on aluminide and overlay coated specimens did not support the Tcrit predictions in terms of cyclic life to failure, although crack density analysis would seem to provide a more accurate correlation between the TMF cycle experienced and the coating performance. A separate finite element (FE) approach to predict the stress/strain history in an overlay coated specimen was developed in order overcome some of the limitations of the Tcrit method. Incorporating developments in the experimental technique, TMF tests on the modified overlay bond coat with and without a thin shell ceramic top coat, were performed. The FE model was extended to include this ceramic in order to investigate the constraint of the bond coat. The level of constraint generated does not seem to support the effect that a thin shell ceramic has on suppressing the surface undulations and crack densities observed in TMF testing.
27
Tattersall, P. "Evaluation and reduction of numerical diffusion effects in viscous aerofoil flow calculations." Thesis, Loughborough University, 1993. https://dspace.lboro.ac.uk/2134/6755.
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The Reynolds-averaged Navier-Stokes (RANS) equations form the most accurate model of viscous flow which can currently be solved computationally on a routine basis for practical engineering problems, given the size and cost of present-day computers. Before RANS solution methods can be used with confidence for the design of aircraft components, a number of areas related to solution accuracy must be investigated, one of which is numerical diffusion. Numerical diffusion, arising from the discrete solution method employed, is necessary to ensure numerical stability, but if too much is included the ability to predict physical phenomena (particularly diffusive ones) accurately can be seriously impaired, with obvious implications for the rational assessmenot f turbulence models. The amount of numerical diffusion in solutions of the RANS equations is evaluated in the present work using two currently popular algorithms, for aerofoil flow test cases. The effect of the numerical diffusion on the prediction of physical processes is investigated, as is the behaviour of the numerical diffusion and corresponding solution when grid quality and algorithm smoothing parameters are varied. Results are presented in two ways, line diagrams giving detailed information along individual grid lines, and contour plots (showing a quantity called the Numerical Diffusion Ratio, NDR) giving overall information on accuracy of the solution throughout the field. The level of numerical diffusion in certain parts of the solution is shown to be unacceptably high in a number of cases. Methods for modifying the NDR are investigated, with the aim of making it suitable for use as a "weighting function" for guiding automatic grid adaptation, to improve solution accuracy. It is shown that some of the modified forms of NDR can be used successfully in this manner. The advantages and disadvantages of using such a solution-accuracy measure (as opposed to the usual solution-activity measures) are discussed and some conclusions and recommendations are made.
28
Mitchell, Robert David. "A computational study of heat transfer on transonic flow over an aerofoil." Thesis, Queen's University Belfast, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.238992.
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Obiga, Otuami. "Investigation of the performance of a slotted aerofoil at low Reynolds numbers." Thesis, University of Nottingham, 2018. http://eprints.nottingham.ac.uk/46841/.
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Slotted aerofoils have been suggested by numerous researchers as an effective means of controlling boundary layer flow separation, and improving aerodynamic performance. Numerous slot designs have been studied at high Reynolds number, but there is scarcity of study of such slots effect on aerofoil performance in low Reynolds number scenarios. In the present work, wind tunnel and numerical investigation of the effect of a unique slot configuration and its geometric parameters on the aerodynamic performance of a NACA0018 aerofoil at low Reynolds number was executed. The aim of this work is to ascertain if the unique slot configuration on the NACA0018 can improve the aerodynamic performance compared to a plain NACA0018, and if the slotted NACA0018 could be applied as rotors on a Darrieus-style vertical axis micro wind turbine for small scale energy conversion at low wind speeds. Four aerofoils were initially fabricated for the wind tunnel tests, each conforming to the NACA0018 profile; a plain aerofoil and three other slotted aerofoils, each with a span–length slot positioned at X=15%, X=45% and X=70% from the leading edge. The, chord length (c), span, slot slope (ψ) and slot width of the slotted aerofoils were 0.25m, 0.3m, 55° and 0.02c respectively. A 2D wind tunnel set up was used in testing the four aerofoils at Reynolds numbers of 92x103 138x103, 184x103 and 230x103, within 0° to 20° range of incidence. Comparing the slotted and plain aerofoils, the aerodynamic force data shows that the presence of the slots was detrimental to aerodynamic performance especially when the slot location is closer to the leading edge. Therefore, a 2D numerical parametric study of slot width and slope was carried out using ANSYS FLUENT 16.0 with the intention of improving the lift–to–drag (L/D) ratio of the span–length slotted aerofoils. Furthermore, a final slot configuration consisting of segmented slot pattern which incorporated the results of the parametric study was fabricated and tested in a wind tunnel. The aerodynamic force analysis shows a 50% increase in L/D ratio of the slotted aerofoil with slot position at X=70%, but its aerodynamic performance was still less than the Plain NACA0018. Thus this work proves that the suggested slot layout did not improve the aerodynamic performance of the NACA0018 aerofoil and as a result, it cannot be recommended to be used as a vertical axis wind turbine rotor. Finally, in order to improve the NACA0018 aerofoil performance, it was suggested that a new slot layout with slot slope on the pressure side inclined towards the leading edge should be designed and studied.
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Margot, Xandra Marcelle. "A physically guided zonal approach for Euler/Navier-Stokes predictions of aerofoil flows." Thesis, Imperial College London, 1993. http://hdl.handle.net/10044/1/11469.
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Adair, Desmond. "Turbulent flow in the vicinity of the trailing-edge of an aerofoil flap." Thesis, Imperial College London, 1986. http://hdl.handle.net/10044/1/37915.
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Breitfeld, Oliver. "Improving the performance of aerofoil sections using momentum transfer via a secondary flow." Thesis, University of South Wales, 2002. https://pure.southwales.ac.uk/en/studentthesis/improving-the-performance-of-aerofoil-sections-using-momentum-transfer-via-a-secondary-flow(a3627a17-2b56-4e3a-9c73-64e0e105f3d6).html.
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Aerodynamic flow control can improve aerofoil performance by influencing the natural growth of boundary layers, which develop on the surface of vehicles moving in viscous fluids. Many active and passive techniques have been developed to reduce drag and/or increase the lift of aerofoil sections. The work presented in this thesis is concerned with the active excitation of the boundary layer on the suction side of aerofoil sections through momentum transfer via a secondary flow. The secondary flow was achieved by air passing through an air breathing device (ABD) which was implemented in the aerofoil surface. This resulted in an almost tangential and uni-directional fluid interaction. Numerical and experimental work showed a beneficial influence of the secondary flow on the aerodynamic characteristics of the studied aerofoil sections. A Taguchi analysis was initially used to confirm findings from previous work on the use of an ABD on a NACA0012 aerofoil section. The resulting parameter ranking showed general agreement with previous data in that the most important parameters are the gap-size i.e. the length over which the two fluids are in contact and the velocity gradient between the two fluids. However, it also raised questions that required an additional in-depth analysis of the parameters governing the flow control process. Due to the greater importance to the modern aviation industry of the NACA65-415 aerofoil section this particular cambered aerofoil section was used for further investigations. This study highlighted the importance of the velocity gradient between the main and secondary flows as well as the location of interaction of the ABD. In addition the gap-size is also important. Consideration of the power requirements for the ABD indicated that this may limit exploitation of the device. An evolutionary search strategy based on genetic algorithms, was employed to optimize the air breathing geometry. This optimisation produced non-intuitive geometries which revealed the importance of promoting an inner fluid recirculation in the device. Finally experimental data in a closed loop wind-tunnel showed trends which were in general agreement with the numerical predictions. However, the measurements indicated significantly greater enhancements of lift forces than those predicted by thenumerical investigation.
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De, Silva R. Suraj I. "The flow field induced by a flat plate aerofoil with simulated battle damage." Thesis, Loughborough University, 2009. https://dspace.lboro.ac.uk/2134/34300.
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This investigation looks into the induced surface pressure field and flow field of a 2-D flat plate aerofoil with simulated battle damage. Enhancing survivability is a major design criteria for military aircraft designers, and the aerodynamic effects of battle damage has recently come to the fore as an area that warrants further attention. The current investigation is in effect a continuation of the battle damage research initiated at Loughborough University by Irwin, with the ultimate aim of enabling quick and accurate CFD studies to assess the aerodynamic impact of battle damage on a lifting surface.
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Rosti, Marco. "Direct numerical simulation of an aerofoil at high angle of attack and its control." Thesis, City, University of London, 2016. http://openaccess.city.ac.uk/15843/.
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Detailed analysis of the flow around a NACA0020 aerofoil at moderate low chord Reynolds number (Rec = 2×104) in completely stalled conditions has been carried out by means of Direct Numerical Simulations. The stalled condition is either a steady configuration at a fixed angle of attack (α = 20o) or it is reached via a ramp-up manoeuvre, increasing the angle of attack from 0o to 20o. Concerning this last case, new insights on the vorticity dynamics leading to the lift overshoot, lift crisis and the damped oscillatory cycle that gradually matches the steady condition, are discussed using a number of post-processing techniques. These include a detailed analysis of the flow ensemble average statistics and coherent structures identification that has been carried out using the Q-criterion and the Finite-Time Lyapunov Exponent technique. Based on the fundamental knowledge achieved in studying the static and the dynamic stall, we introduced a biomimetic passive control technique to mitigate the aerodynamic performance degradation typical of such flow conditions. In particular, the envisaged control technique has been inspired by the dorsal feathers that are used by almost all birds to adapt their wing characteristics to delay stall or to moderate its adverse effects (e.g., during landing or sudden increase in angle of attack due to gusts). Some of the feathers are believed to pop up as a consequence of flow separation and to interact with the flow producing beneficial modifications of the unsteady vorticity field. The adoption of self adaptive flaplets in aircrafts, inspired by birds feathers, requires the understanding of the physical mechanisms leading to their aerodynamic benefits and the determination of the characteristics of optimal flaps including their size, positioning and ideal fabrication material. In this framework, we have used numerical simulation to study the effects of this passive control technique in both steady and dynamic stall. In particular, for the static case, we have defined an optimal condition as the one that delivers the highest lift coefficient CL, preserving or improving the aerodynamic efficiency E = CL/CD. To achieve a condition close to optimality we started by considering a simplified scenario, to determine the main characteristics of the flap (i.e., variations of its length, position and natural frequency). Later on, a detailed direct numerical simulation analysis is used to understand the origin of the aerodynamic benefits introduced by the pop-up of the optimal flaplet. It is found that an optimal flap can deliver a mean lift increase of about 20% on a NACA0020 aerofoil at an incidence of 20o degrees. The analysis of direct numerical simulation data of the flow field around the aerofoil equipped with the optimal flap allowed to elucidate the main mechanism that promotes the aerodynamic improvements. In particular, it is found that the flaplet movement, induced by the transit of a large recirculation bubble on the aerofoil suction side, displaces the trailing edge vortices further downstream, away from the wing. The downstream displacement of the trailing edge generated vortices, limits the downforce generated by those vortices also regularising the shedding cycle that appears to be much more organised when the flaplet is activated. A similar study has also been carried out for the dynamic case. We have analysed the effects produced by the presence of an elastically mounted flap on the transient behaviour of the flow fields. For a specific ramp-up manoeuvre characterised by a reduced frequency slower the shedding one, it is found that it is possible to design flaps that limit the severity of the dynamic stall breakdown. In particular, it is possible to increase the value of the lift overshoot and to smooth its abrupt decay in time. A detailed analysis on the modification of the unsteady vorticity field due to the flap-flow interaction during the ramp-up motion is also provided to explain the physical mechanism that lead to more benign aerodynamic response.
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Yeung, Matthew. "The sound field due to a pair of vortices moving past a thin aerofoil." Thesis, Imperial College London, 2003. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.404471.
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Jones, Gareth. "Control of flow around an aerofoil at low Reynolds numbers using periodic surface morphing." Thesis, Imperial College London, 2016. http://hdl.handle.net/10044/1/51101.
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This thesis combines experimental and computational methods to investigate the low Reynolds number flow (Re = 50,000) around a NACA 4415 aerofoil, and its control using periodic surface motion. A physical model was fabricated and tested in a closed-loop wind tunnel and a good comparison between the experiments and computations was achieved. Time-resolved measurements of the surface reveal that the peak-to-peak displacement is a function of both the amplitude and frequency of the input voltage signal but the addition of aerodynamic forces does not cause significant changes in surface behaviour. The vibration mode shape exhibits a single peak and is uniform in the spanwise direction at frequencies below 80Hz, above which a change in the vibration mode occurs. The flow around the actuated aerofoil was compared with the baseline (i.e. unactuated) flow. The latter exhibits a large separation region and, as a result, produces relatively high drag and low lift forces. By analysing the experimental and computational data, the large separation zone was found to be the result of laminar separation without reattachment. Transition to turbulence does occur but too close to the trailing edge, and far from the wall, for sufficient pressure recovery to take place for reattachment. When actuated at 70 Hz, the frequency spectra in the vicinity of the trailing edge and near-wake was found to be dominated by the actuation frequency. Sharp peaks suggest the production of Large Coherent Structures at this frequency. In agreement with the experiments, the computations revealed that the vortex shedding from the shear layer was 'locked-on' to the surface motion and spanwise coherent vortices were produced during each actuation cycle. The increased momentum entrainment associated with them enabled a large suppression of the separated region, which was seen in both the experiments and computations. The result was a simultaneous increase in Lift and decrease in Drag and therefore a large increase in the L/D ratio.
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Song, Longfei. "Leading-edge flow separation control over a NACA 0012 aerofoil with DBD plasma actuators." Thesis, University of Nottingham, 2018. http://eprints.nottingham.ac.uk/49841/.
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An experimental investigation has been conducted in a low-speed wind tunnel at the University of Nottingham to study the flow separation control capability of a wall-normal plasma jet by DBD plasma actuator over a NACA 0012 aerofoil. As an active flow separation control technique, DBD plasma actuators could be applied when required to manipulate a flow. They are surface-mounted and require no moving parts, ducts, holes or cavities, so no profile drag penalty will be caused. Moreover, they are fast responding since they are purely electrical devices and could be operated at a higher frequency relative to other flow control techniques. DBD plasma actuators are easy to manufacture, low in weight, low energy consuming and can be easily fitted to aerofoils. Therefore, they are ideal tools to control the flow separation around aerofoil. Up to date, wall plasma jet was used to add momentum to flow directly so that flow becomes more energetic and capable of withstanding adverse pressure gradient. In this study, a wall-normal plasma jet by steady actuation of plasma actuator was investigated and PIV results show that it has the capability of controlling the separation around aerofoil at post-stall angles of attack. The wall-normal jet is bent towards freestream direction and some small-scale vortical structures are created due to the interaction between the wall-normal plasma jet and freestream. These vortical structures could promote mixing and transport high-momentum fluids into the boundary layer, which affects the flow above the suction surface significantly. Moreover, unsteady actuation of plasma actuator was also utilised to control the flow separation around aerofoil. It was found that it has a stronger ability to control flow separation even at a much lower energy consumption than steady actuation of plasma actuator. PIV measurements demonstrate that separated flow could be reattached at post-stall angle of attack of 14° with only 10% of the energy consumption by steady actuation. Flow is well organized and a series of large-scale vortices are created with periodic activation of plasma actuator, these vortices enhance entrainment and the outwards transport of fluids from aerofoil surface leads to a favourable pressure gradient, resulting in a control of flow separation.
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Kadja, M. "Computation of recirculating flow in complex domains with algebraic Reynolds stress closure and body fitted meshes." Thesis, University of Manchester, 1987. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.384419.
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Greene, Robert Matthew. "Finite element simulation and die shape optimisation method applied to forging of 3D aerofoil blades." Thesis, Queen's University Belfast, 2008. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.485072.
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The dimensional tolerances are among the most important manufacturing criteria in the forging of aerofoil blades for aero-engine applications. These are usually classified in relation to aerofoil shape, thickness and twist along cross-sections and are. affected by many factors such as preform shape, temperature, die elasticity and component springback. The interaction of all these factors makes design for net-shape forging extremely difficult. The focus was on the final shape of the forged aerofoil blades, which was affected by temperature, die elasticity and springback. A thermaIly coupled 3D thermo-elasto-visco-plastic analysis was employed to . investigate the entire forging cycle including forging, removal of dies and cooling of the aerofoil section, so that the e{fect of forging and post forging conditions such as temperature can be taken into account. A die-shape compensation strategy is presented in this· thesis. A di~ cleaning technology has been dev.eloped to represent the dies as a series of sampled points. These points were subsequently utilized as spline control points in the definition of a 3D spline surface. Methods are presented to evaluate the net shape error and to use this error data iri order to redefine the die . surface for a successive iteration. An iterative procedure was applied and a significant reduction in deviation from nominal shape was achieved The various components in this study have been developed in parallel to aIlow the successful integration of CAD, CAE, FE simulation and die shape modification in the optimisation of aerofoil blade dies to achieve net shape forging.
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Makem, J. E. "Virtual Net-Shape Forging of Aerofoil Blades - Dimensional Inspection and Shape Sensitivity to Process Variables." Thesis, Queen's University Belfast, 2010. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.517543.
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Lau, Alex Siu Hong. "High-order computations on aerofoil-gust interaction noise and the effects of wavy leading edges." Thesis, University of Southampton, 2012. https://eprints.soton.ac.uk/355961/.
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High-order accurate numerical simulations are performed to investigate the effects of wavy leading edges on aerofoil gust interaction (AGI) noise. The present study is based on periodic velocity dis-turbances predominantly in streamwise (x-) and vertical (y-) directions that are mainly responsible for the surface pressure fluctuation of an aerofoil. The perturbed velocity components of the present gust model do not vary in the spanwise (z-) direction. In general, the present results show that wavy leading edges lead to reduced AGI noise. Under the current incident gusts, it is found that the ratio of the wavy leading-edge peak-to-peak amplitude (LEA) to the longitudinal wavelength of the incident gust (λg) is the most important factor for the reduction of AGI noise. It is observed that AGI noise reduces with increasing LEA/λg, and significant noise reduction can be achieved for LEA/λg≥0.3. The present results also suggest that any two different cases with the same LEA/λg lead to a strong similarity in their profiles of noise reduction relative to the straight leading-edge case. The wavelength of wavy leading edges (LEW), however, shows minor influence on the reduction of AGI noise under the present gust profiles used. Nevertheless, the present results show that a meaningful improvement in noise reduction may be achieved when 1.06LEW/λg 61.5. In addition, it is found that the beneficial effects of wavy leading edges are maintained for various angles of attack and aerofoil thicknesses. Also, wavy leading edges remain effective in reducing AGI noise for gust profiles containing multiple frequency components. It is discovered in the current research that wavy leading edges result in in-coherent response time to the incident gust across the span, which causes a decreased level of surface pressure fluctuations, hence a reduced level of AGI noise.
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Tsiachris, Fotios K. "Retreating blade stall control on a NACA 0015 aerofoil by means of a trailing edge flap." Thesis, University of Glasgow, 2005. http://theses.gla.ac.uk/5109/.
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Trailing edge flaps may provide a mechanism for alleviating retreating blade stall. In the present investigation numerical simulations were conducted involving a NACA 0015 aerofoil section fitted with a plain trailing edge (TE) flap. All simulations were conducted using DIVEX, a tool being developed at the University of Glasgow, Department of Aerospace Engineering. In summary, the code uses a surface shedding discrete vortex method (DVM) for the simulation of 2-D incompressible flows around pitching aerofoils. The aero-foil is oscillating in pitch about its quarter chord axis and the clap undergoes negative pitch inputs, i.e. upward. An interesting feature appears to be that the cause of the severe nose down pitching movement introduced during dynamic stall is due to the cortical pair of the DSV and TEV where it is shown that the former feeds the latte in the case of the clean aerofoil for the range of reduced frequencies varying between k = 0.128 and k = 0.180. This fact suggests that manipulation of the vorticity in the vicinity of the trailing edge may be a mechanism for modification of the dynamic stall vortex (DSV) trajectory. This was found to relieve the aerofoil from severe pitching moment undershoot occurring during dynamic stall under appropriately phased flap actuations. Results obtained so far encourage the employment of a flap with fairly small size, 15% of the aerofoil chord. A parametric study is described which identifies the proper aerodynamic and actuation parameters for the current problem. In addition a simple open loop control scheme is developed based purely on rotor and flap related quantities.
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Lewington, Neil. "Enhancing lift on a three element high lift aerofoil system by installing air jet vortex generators." Thesis, City University London, 2001. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.269315.
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Powell, Jonathan Edward. "The application of particle image velocimetry to vortical flow fields." Thesis, Heriot-Watt University, 2000. http://hdl.handle.net/10399/575.
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Walter, Daniel James, and Daniel james walter@gmail com. "Study of aerofoils at high angle of attack in ground effect." RMIT University. Aerospace, Mechanical and Manufacturing Engineering, 2007. http://adt.lib.rmit.edu.au/adt/public/adt-VIT20080110.145138.
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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.
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Gobbi, Giangiacomo. "Analysis and reconstruction of dynamic-stall data from nominally two-dimensional aerofoil tests in two different wind tunnels." Thesis, University of Glasgow, 2010. http://theses.gla.ac.uk/1362/.
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This work is a specific investigation into low speed aerofoils. The term “low speed” is normally used to indicate free stream velocity less than Mach = 0.5 and, here, not more than 0.2 M when considering dynamic-stall. This field of investigation, for the QinetiQ aerofoil, has been somewhat ignored till now to the advantage of higher speeds starting from 0.3 M. In order to improve the knowledge of the behaviour of aerofoils under M<0.2 conditions, the University of Glasgow, in cooperation with QinetiQ, carried out two-dimensional aerodynamic tests on a RAE9645 aerofoil in 2002. By the end of November, of the same year, high quality unsteady pressure measurements from dynamic-stall tests were available. The tests were conducted on two different RAE9645 aerofoil models in two different wind tunnels. The first of these data came from the aerofoil that was tested in the Department of Aerospace’s Handley Page Wind Tunnel. The second data set was from tests carried out by QinetiQ on an aerofoil in the Department of Aerospace’s Argyll Wind Tunnel. The objectives of this investigation are divided in three main topics. First part considers the analysis of the data. This means (a) the assessment of the aerodynamic coefficients and consequent analysis of the various features of the dynamic-stall including the critical angle, the pitching moment and stall onset. (b) A comparison of the overall aerodynamic coefficients and (c) the carry out of final analysis of the most important quantities such as Cp deviation, vortex development and convection speed and re-establishment of fully attached flow. The assessment of the all same quantities for the second aerofoil tested by QinetiQ and the comparison of them xxiii with the first model are the objectives of the second part of the project. Hence a most useful comparison of two data sets from two different wind tunnels will be achieved. The third part was to establish the coefficients for the Beddoes third generation dynamic-stall model for the clean aerofoil without any flow control, using both aerofoil data. The Beddoes third generation dynamic-stall model is the last version of a model which has been in constant development over thirty years and is known as the most popular semi empirical method for assessing unsteady airloads such as lift, drag and pitching moment. This applies both to helicopters and wind turbines. The simplicity and undergoing philosophy of this method is its strength, especially compared with the current solution of Navier-Stokes or Euler equations. At the completion of this work, all the coefficients and information necessary for running the Beddoes simulating dynamicstall model were obtained for the RAE9645 aerofoil. At the same time refinements, improvements and new guide lines were pursued in order to make the model easier and more powerful than before. Some of these changes are associated only to low Mach numbers. It has been concluded that the Beddoes’ model has been enhanced to better re-construct the RAE9645 aerofoils data of low Mach numbers.
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Carr, M. I. "The excitation of acoustic resonances in an axial flow compressor stage by vortex shedding from aerofoil section blading." Thesis, Swansea University, 1986. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.636209.
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In recent years continuing development of the axial flow compressor for use in the aero-engine has increased its susceptibility to unsteady flow phenomena which can cause severe blade vibration. A source which has emerged and become of considerable importance is excitation by acoustic resonances. An experimental investigation in a single stage axial flow compressor rig has been performed to ascertain whether acoustic resonances can be excited by vortex shedding from loaded aerofoil section blades. A further experimental programme, to study further the effect of inter-row spacing, was peformed in both an open jet facility and a wind tunnel facility with a tandem plate arrangement. Results showed that acoustic resonances could be excited in a compressor stage in which there was severe blade loading. The speed range over which the resonances were excited was demonstrated to be not only a function of the degree of loading but also the inter-row spacing. Vortex shedding will drive a resonance when the shedding is correlated by the resonant acoustic field and interaction between the vortices and the acoustic field in the vicinity of the blades may result in a net positive input of acoustic energy. As a result the phase of the acoustic field as vortices pass over the trailing edge of the shedding blades and the leading and trailing edges of the downstream blades, control the energy generation. The inter-row spacing controls the phase of the downstream blade interaction and therefore is a major factor influencing the resonant acoustic amplitude. As well as the fundamental acoustic mode, a resonance can also drive significant blade vibration in two other consequential frequency bands which are: a) Sum and Difference frequency bands due to acoustic non-linearity and b) Sidebands of the fundamental modes due to spatial modulation effects caused by flow distortions.
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Mulholland, Daniel J. "Distribution of aerofoil section lift from section pitching moment applied to a dynamic model for aircraft poststall departure." Thesis, This resource online, 1995. http://scholar.lib.vt.edu/theses/available/etd-02132009-172236/.
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El-Ibrahim, Salah Jamil Saleh. "Prediction of the effects of aerofoil surface irregularities at high subsonic speeds using the Viscous Garabedian and Korn (VKG) method." Thesis, University of Hertfordshire, 2000. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.365928.
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Irwin, Andrew J. "Investigation into the aerodynamic effects of simulated battle damage to a wing." Thesis, Loughborough University, 1999. https://dspace.lboro.ac.uk/2134/10735.
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A key stage in the design-cycle of a military aircraft is the assessment of its vulnerability to hostile threat mechanisms. Such mechanisms inflict battle-damage to the aircraft structure and systems. This experimental investigation considered the aerodynamic consequences of simulated battle-damage to a two-dimensional wing. Key assumptions and techniques were identified leading to the modelling of both gunfire and missile fragmentation damage. Wind tunnel balance measurements were undertaken, together with surface pressure measurements and flow-visualisation methods. Force and moment results indicated extensive changes in coefficient values, whilst both smoke and surface visualisation paint successfully indicated the flow mechanisms present. Using these techniques the influences of damage and experimental variables were investigated, including damage type, size, location and Reynolds Number. Studies were also made into cases of multiple gunfire holes and the influence of internal wing construction. Results indicated that damage at quarter and half-chord locations gave greater coefficient changes than those seen for either leading or trailing edge damage. This was primarily due to reductions in the upper surface pressure peak due to through-flow. Such reductions were seen to extend in both a chordwise and spanwise direction. The flow mechanism identified indicated both similarities and differences to those of flat-plate jets in crossflows. Analysis of both gunfire and missile damage data lead to the development of a set of empirical relationships, which related damage location and size to coefficient changes.