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1
Reid, Michael R. "Thin/cambered/reflexed airfoil development for micro-air vehicles at Reynolds numbers of 60,000 to 150,000 /." Electronic version of thesis, 2006. https://ritdml.rit.edu/dspace/handle/1850/2607.
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Huang, Liang. "Optimization of blowing and suction control on NACA0012 airfoil using genetic algoirthm with diversity control." Lexington, Ky. : [University of Kentucky Libraries], 2004. http://lib.uky.edu/ETD/ukymeen2004d00153/LiangDis.pdf.
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Thesis (M.S.)--University of Kentucky, 2004.Title from document title page (viewed Oct. 12, 2004). Document formatted into pages; contains xii, 113 p. : ill. Includes abstract and vita. Includes bibliographical references (p. 102-112).
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Janjua, Zaid Ayaz. "Ice accretion on aerofoils." Thesis, University of Nottingham, 2017. http://eprints.nottingham.ac.uk/45409/.
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Ice accretion on aerofoils is a problematic phenomenon affecting power lines, ships and aircraft wings. This work thus undertakes an experimental and computational investigation into the formation and adhesion of ice on aerofoils. An experimental setup to test the adhesion strength of ice was designed and tested for repeatability and the effect of temperature on it. It was found that the ambient temperature has a profound effect on the adhesion strength, possibly due to dependence on the heat transfer mechanism through an amorphous liquid-like layer between ice and substrate. The tests were expanded to determine the effect of contact angle parameters on the icephobicity of 14 nanocoatings. It was found that the surface should possess high receding contact angle and low CAH to reduce adhesion thereby reducing the ice-substrate contact points. Hydrophobicity and icephobicity may not necessarily be dual characteristics of a surface unless the aforementioned criteria is satisfied. Anti-icing tests on the same coatings showed that the freezing time of a droplet on the surface reduces with an increase in static contact angle. To understand the role of mixed ice, a one dimensional model is introduced to measure the accretion of mixed, rime and glaze ice on an aerofoil. This process occurs in four distinct stages and the effect of atmospheric parameters on the transition time between different growth types and height is determined. This mode was developed further to include a convective term to determine the profile of ice when rime grows above glaze/mixed with water flowing inside. This is a first step towards understanding the links between porous structures, ice structures and runback water that can generate interesting icy structures. This work forms part of the ICECOAT project funded by the EU Framework 7 CleanSky programme under grant award JTI-CS-2012-02-SFWA-01-051.
4
Arbos, Torrent Sara. "Aeromechanical performance of compliant aerofoils." Thesis, Imperial College London, 2013. http://hdl.handle.net/10044/1/28105.
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The aeromechanics of compliant aerofoils are studied. Several experimental techniques including hot-wire anemometry, particle image velocimetry, high speed photogrammetry and strain gauge force measurements are used. Tests are performed at a chord based Reynolds number of Re= 4x10^4 and angles of attack between 0° and 25°. They explore the impact of the geometry of the leading- and trailing-edge supports as well as the rigid- ity of the aerofoil on the aeromechanics and aerodynamics of membrane aerofoils. Tests on latex membrane wings subjected by four different types of supports are performed. Firstly, the study focuses on the structural performance by evaluating detailed measure- ments of membrane deflections and lift and drag forces. It will be shown that the use of lower bending stiffness supports results in noticeable deformations, both static and dynamic, especially at mid-to-high incidences. Moreover, the conjunction of hot-wire results with photogrammetry imagery of the membrane deformation indicates that the membrane vibration is coupled with the vortex shedding. This, when coupled with a low-stiffness rectangular cross-section leading- and trailing-edge, results in large amplitude vibrations affecting the membrane, the support and the wake. Hence, a more detailed study of the vortex shedding and the wake attributes is presented. The findings indicate that for low angles of attack the wake characteristics are highly affected by the leading- and trailing- edge geometry; as incidence increases the wake characteristics become less dependant on the support's geometry, eventually reaching a point in which they are fully independent of it and closely resembling a fully stalled rigid aerofoil. Finally, the effects of the aero- foil rigidity are analysed. Tests of varying thickness but constant Young's modulus on unidirectional carbon fibre composite plates are performed. Results show that the Weber number is a crucial parameter when defining the properties and performance of the wing. Furthermore, the study will show that lift and drag forces are higher for membrane wings than for composite plates and that the dynamic motions of the composite plates increase as the plate thickness is decreased resulting in earlier wing stall and worse post-stall be- haviour than membrane wings. The results of this study should provide valuable insight for future use of membrane wings in micro air vehicles.
5
Nash, Emma Clare. "Boundary layer instability noise on aerofoils." Thesis, University of Bristol, 1996. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.337698.
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Tse, Man-Chun. "Overall effects of separation on thin aerofoils." Thesis, McGill University, 1991. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=74592.
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The separation bubble at the leading edge of a thin sharp-edged aerofoil in steady, incompressible and two-dimensional flow was studied. A simple method, using irrotational flow and source singularities, has been developed for predicting bubble reattachment length, drag and lift.For a flat-plate aerofoil predictions compare favourably with new experiments. The non-dimensional reattachment length $x sb R over rm c$ is proportional to the square of the incidence $( alpha)$ and the slope $x sb R over rm c alpha sp2$ depends on the growth of the outer part of the separated shear layer. The value of the term $x sb R over rm c alpha sp2$ was determined experimentally as $ pi over 0.08$. At incidences above 2$ sp circ$, the bubble drag becomes increasingly dominant when compared with the skin friction drag. Although the details of the bubble geometry are not simulated, the lift and stall are predicted fairly well.
The theory is extended to a circular-arc aerofoil. This part of the study is much less satisfactory. New experimental measurements do not appear to be sufficiently accurate to provide the empiricism to support the extended theory which must now account for regions of separated flow near the trailing edge.
7
Suddhoo, A. "Inviscid compressible flow past multi-element aerofoils." Thesis, University of Manchester, 1985. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.356714.
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Palmer, Nathaniel Thomas. "Surge-induced deflections of axial compressor aerofoils." Thesis, Cranfield University, 2004. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.442401.
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Scarbrough, William T. "NACA four-digit airfoil section generation using cubic parametric curve segments and the golden section /." Online version of thesis, 1992. http://hdl.handle.net/1850/11033.
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Chantharasenawong, Chawin. "Nonlinear aeroelastic behaviour of aerofoils under dynamic stall." Thesis, Imperial College London, 2007. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.440548.
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Khorrami, Ahmad Farid. "Hypersonic aerodynamics on flat plates and thin aerofoils." Thesis, University of Oxford, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.292584.
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Trevelyan, Conrad. "Application of circulation control aerofoils to wind turbines." Thesis, Loughborough University, 2002. https://dspace.lboro.ac.uk/2134/34575.
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Circulation control aerofoils potentially offer an additional means of load and power control for horizontal axis wind turbines by virtue of their rapid response time. Their suitability for these tasks has been assessed with respect to the power which they absorb, their interaction with aerofoils used on modern wind turbines, the infrastructure or hardware which they require and the degree to which they can affect the loads experienced by the turbine blades and other major components. It has been determined that the type of circulation control aerofoil most suited to use on wind turbine blades are those of the jet flap type and it has been realised that an ability to shed, as well as increase loads is advantageous in this application. To this end the behaviour of both negatively and positively deflected jets have been investigated with a two-dimensional computational fluid dynamics code, validated in the course of this work for such modelling. Particular emphasis has been placed on minimising the input power requirements of the circulation control aerofoils and in proposing an overall system that has the required level of robustness and reliability. A 2MW turbine has been modelled with a blade element momentum theory code in order to compare performance with and without circulation control aerofoils. These initial results show that there may be some positive benefits to be gained, but that the energy demands of the system place a hard limit on the degree to which circulation control aerofoils can determine the forces experienced by the turbine.
13
Cyr, Stéphane. "Prediction of flow separation over rigid and flexible aerofoils." Thesis, McGill University, 1996. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=40335.
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The separation of the boundary layer at the trailing edge of sails is considered in the present work. This problem was identified as one of the major reasons for the disagreement between existing theories and experimental results. The separation of the flow on flexible surfaces is a complex phenomenon because of the interdependence between the shape of the sail, the pressure distribution and the separation position. The problem is simplified by considering the flow over a two-dimensional sail set at design incidence to isolate the effect of the separation of the flow.A theoretical model predicting the pressure distribution around a rigid aerofoil with trailing edge separation has been developed. The model uses a point source in potential flow to simulate the effect of the separated region on the flow. The separation position is determined using a turbulent boundary layer development calculation based on an integral method. The source strength and its position are modified iteratively until a set of closure conditions is satisfied.
The model is first tested on circular-arc aerofoils and validated against experimental results of obtained on five different circular-arc aerofoils of camber ratio ranging from 10% to 27%. The theoretical results are also compared with other experimental results found in the literature. The agreement between the viscous theory and the experimental results is satisfactory up to camber ratios of approximately 23%.
The theoretical model is then extended to simulate the flexibility of a two-dimensional sail. The shape of a rigid aerofoil is modified iteratively using the pressure distribution obtained from the flow separation model and the sail equation until convergence.
The results of the viscous theory for two-dimensional sails are compared with a new set of experimental results obtained on four different two-dimensional sails of excess-length ratios ranging from 0.097 to 0.167. The predictions of the theory for a lower range of excess-length ratios are compared with experimental results from other authors. For the two-dimensional sails the viscous theory is generally in good agreement with experimental measurements up to excess-length ratios of 0.167.
The theory presented in the present thesis demonstrates that a single source can simulate the essential effect that separation of the boundary layer near the trailing edge has on a thin cambered aerofoil. The integration of the separation model in an iterative scheme made it possible to determine the effect of trailing edge separation on the aerodynamic performance of a two-dimensional sails at design incidence.
14
Gillan, Mark Andrew. "A computational analysis of viscous flow over porous aerofoils." Thesis, Queen's University Belfast, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.333813.
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Lam, Chun-Ming Gordon. "Nonlinear wake evolution of Joukowski aerofoils in severe maneuver." Thesis, Massachusetts Institute of Technology, 1989. http://hdl.handle.net/1721.1/14177.
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Ockfen, Alex Earle. "Viscous modeling of ground effect aerodynamics of airfoil and jet." Pullman, Wash. : Washington State University, 2008. http://www.dissertations.wsu.edu/Thesis/Fall2008/a_ockfen_112408.pdf.
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Thesis (M.S. in mechanical engineering)--Washington State University, December 2008.Title from PDF title page (viewed on Dec. 31, 2008). "School of Mechanical and Materials Engineering." Includes bibliographical references (p. 149-154).
17
Mokhtarian, Farzad. "Fluid dynamics of airfoils with moving surface boundary-layer control." Thesis, University of British Columbia, 1988. http://hdl.handle.net/2429/29026.
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The concept of moving surface boundary-layer control, as applied to the Joukowsky and NACA airfoils, is investigated through a planned experimental program complemented by theoretical and flow visualization studies. The moving surface was provided by one or two rotating cylinders located at the leading edge, the trailing edge, or the top surface of the airfoil. Three carefully designed two-dimensional models, which provided a wide range of single and twin cylinder configurations, were tested at a subcritical Reynolds number (Re = 4.62 x 10⁴ or Re — 2.31 x 10⁵) in a laminar-flow tunnel over a range of angles of attack and cylinder rotational speeds. The test results suggest that the concept is indeed quite promising and can provide a substantial increase in lift and a delay in stall.
The leading-edge rotating cylinder effectively extends the lift curve without substantially
affecting its slope. When used in conjunction with a second cylinder on the upper surface, further improvements in the maximum lift and stall angle are possible. The maximum coefficient of lift realized was around 2.22, approximately 2.6 times that of the base airfoil. The maximum delay in stall was to around 45°. In general, the performance improves with an increase in the ratio of cylinder surface speed (Uc) to the free stream speed (U). However, the additional benefit derived progressively diminishes with an increase in Uc/U and becomes virtually negligible for Uc/U > 5.
There appears to be an optimum location for the leading-edge-cylinder. Tests with the cylinder at the upper side of the leading edge gave quite promising results. Although the CLmax obtained was a little lower than the two-cylinder configuration (1.95 against 2.22), it offers a major advantage in terms of mechanical simplicity. Performance of the leading-edge-cylinder also depends on its geometry. A scooped configuration appears to improve performance at lower values of Uc/U (Uc/U ≤ 1). However, at higher rates of rotation the free stream is insensitive to the cylinder geometry and there is no particular advantage in using the scooped geometry.
A rotating trailing-edge-cylinder affects the airfoil characteristics in a fundamentally different manner. In contrast to the leading-edge-cylinder, it acts as a flap by shifting the CL vs. α plots to the left thus increasing the lift coefficient at smaller angles of attack before stall. For example, at α = 4°, it changed the lift coefficient from 0.35 to 1.5, an increase of 330%. Thus in conjunction with the leading-edge- cylinder, it can provide significant improvements in lift over the entire range of small to moderately high angles of incidence (α ≤ 18°).
On the theoretical side, to start with, the simple conformal transformation approach
is used to obtain a closed form potential-flow solution for the leading-edge-cylinder configuration. Though highly approximate, the solution does predict correct trends and can be used at a relatively small angle of attack. This is followed by an extensive numerical study of the problem using:
• the surface singularity approach including wall confinement and separated flow effects;
• a finite-difference boundary-layer scheme to account for viscous corrections; and
• an iteration procedure to construct an equivalent airfoil, in accordance with the local displacement thickness of the boundary layer, and to arrive at an estimate for the pressure distribution.
Effect of the cylinder is considered either through the concept of slip velocity or a pair of counter-rotating vortices located below the leading edge. This significantly
improves the correlation. However, discrepancies between experimental and numerical results do remain. Although the numerical model generally predicts CLmax with a reasonable accuracy, the stall estimate is often off because of an error in the slope of the lift curve. This is partly attributed to the spanwise flow at the model during the wind tunnel tests due to gaps in the tunnel floor and ceiling required for the connections to the externally located model support and cylinder drive motor. However, the main reason is the complex character of the unsteady flow with separation and reattachment, resulting in a bubble, which the present numerical
procedure does not model adequately. It is expected that better modelling of the cylinder rotation with the slip velocity depending on a dissipation function, rotation, and angle of attack should considerably improve the situation.
Finally, a flow visualization study substantiates, rather spectacularly, effectiveness
of the moving surface boundary-layer control and qualitatively confirms complex
character of the flow as predicted by the experimental data.Applied Science, Faculty of
Mechanical Engineering, Department of
Graduate
18
Shaw, Scott. "Numerical study of the unsteady aerodynamics of helicopter rotor aerofoils." Thesis, Cranfield University, 1999. http://hdl.handle.net/1826/4187.
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A two-dimensional model of the aerodynamics of rotor blades in
forward flight is proposed in which the motion of the blade is
represented by periodical variations of the freestrearn velocity and
incidence. A novel implicit methodology for the solution of the
compressible Reynolds averaged Navier-Stokes equations and a twoequation
model of turbulence is developed. The spatial discretisation is
based upon Osher's approximate Riernann solver, while time integration
is performed using a Newton-Krylov method.
The method is employed to calculate the steady transonic
aerodynamics of two supercritical aerofoils and the unsteady
aerodynamics of pitching aerofoils. Comparison with experiment and
independent calculations for these test cases is satisfactory. Further
calculations are performed for the self-excited periodic flow around a biconvex
aerofoil. Comparison of quasi-steady and unsteady calculations
suggests that the flow instability responsible for the self-excited flow is
due to the presence of a shock induced separation bubble in the
corresponding steady flow.
Finally the method is used to predict the aerodynamics of aerofoils
performing inplane and combined inplane-pitching motions. Results
show that quasi-steady aerodynamic models are unsuitable at conditions
representative of high-speed forward flight. For shock free flows, the
unsteady effects of freestrearn oscillations can be represented by a
simple phase lag. For transonic flows the influence of unsteadiness on
shock wave dynamics is shown to be complex. Calculations for indicial
motion show that the unsteady behaviour of the flow is related to the
finite time taken by disturbance waves to travel to the shock wave from
the leading and trailing edges of the aerofoil.
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Berci, Marco. "Multidisciplinary multifidelity optimisation of flexible wing aerofoils by passive adaptivity." Thesis, University of Leeds, 2011. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.558796.
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This thesis is directed at developing and assessing a multifidelity model-based methodology for the flight performance analysis and multidisciplinary optimisation of flexible wing aerofoils. Such a strategy is pursued because of the high computational cost involved of solving such optimisation problems via high-fidelity simulations only. The methodology is applied to the preliminary design of a small flexible winged Unmanned Air Vehicle (UAV), the likes of which are particularly susceptible to wind gusts. The strategy adopted is directed at optimising both the passively adaptive structure and the shape of the flexible UAV wing for aerodynamic performance (i.e., drag reduction), weight reduction and gust response alleviation, formulated as an unsteady coupled Fluid- Structure Interaction (FSI) problem. A metamodel of the high-fidelity model response, based on a tuned low-fidelity one, is built in order to verify and validate the approach for aeroelastic problems. Both models are based on solutions of the aeroelastic equations for the wing Typical Section and the low-fidelity response tuned accordingly as prescribed by suitable Design of Experiments (DOEs). Several levels of complexity and computational cost are employed for modelling aerodynamics and structural dynamics. The role of aerodynamic damping, structural nonlinearities and turbulent Computational Fluid Dynamics (CFD) is investigated. Good agreement between the high-fidelity results and corrected low-fidelity ones shows that the methodology is suitable for use m aeroelastic performance optimisation problems. Using the multifidelity strategy developed, the flexible wing of a small UAV is optimised for best flight efficiency under aero-structural constraints. The wing structure is assumed fully flexible and a semi-analytical model for the aeroelastic analysis and gust response of a flexible Typical Section developed. Having tuned the low-fidelity response, a Genetic Algorithm (GA) is employed to fmd the global optimum, showing that a flexible wing aerofoil is characterised by a higher aerodynamic efficiency than its rigid counterpart.
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Vezza, Marco. "Numerical methods for the design and unsteady analysis of aerofoils." Thesis, University of Glasgow, 1986. http://theses.gla.ac.uk/4885/.
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Liu, Xiao. "Aerodynamic and wake development of aerofoils with trailing-edge serrations." Thesis, University of Bristol, 2018. http://hdl.handle.net/1983/b43f0a81-47ae-445e-84b6-4a4a49ff9c3e.
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A comprehensive study of the aerodynamic performance and wake development of aerofoils fitted with different types of trailing-edge serrations is provided. A symmetric NACA 0012 and a cambered NACA 65(12)-10 aerofoils have been studied experimentally. The aerodynamic force measurements have shown that the use of trailing-edge serrations can lead to significant reduction of the lift coefficient at low angles of attack for cambered aerofoils, while remaining negligible for symmetric aerofoils. The force measurements have also shown that the stall characteristics of the aerofoils do not change greatly as a result of the implementation of the trailing-edge serrations. The wake flow characteristics have been investigated using Particle Image Velocimetry (PIV), Laser Doppler Velocimetry (LDV) and hot-wire anemometer to improve the understanding of the wake development and the energy-frequency content of the wake turbulent structures associated with serrated trailing-edge. At relatively high angles of attack, where maximum lift-to-drag can be obtained, the turbulent kinetic energy and Reynolds stress results for the cambered aerofoils have shown that the wake flow turbulence can be greatly reduced in the aerofoil near wake region by using trailing-edge serrations, while that for the symmetric aerofoil were found to be less effective. The reduction of the wake turbulence for serrated cambered aerofoils is believed to be due to a complex interaction between the flow field over the tip and root planes and three-dimensional turbulent structures in the near-wake. The implementation of serrations has also shown a significant reduction of the wake flow energy over a wide range of frequencies. The reduction of turbulence within the near wake offers a new possibility in reducing noise generated by wake-aerofoil interaction. This will have significant implications for industrial applications involving multiple rows of aerofoils such as contra-rotating propellers, rotor-stator configuration, canard-wing body configuration, etc.
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Serrano, Galiano Sonia. "Fluid-structure interaction of membrane aerofoils at low Reynolds numbers." Thesis, University of Southampton, 2016. https://eprints.soton.ac.uk/414110/.
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This thesis investigates the fluid-structure interaction (FSI) problem of elastic membrane aerofoils at low Reynolds numbers. The dynamics of the fluid and membrane coupled system is studied via direct numerical simulation (DNS) using a newly developed computational framework whose characteristics and validation are included in this report. A set of two-dimensional DNS were performed for varying Reynolds number, membrane elasticity and aerofoil geometry in order to investigate the effect of these relevant fluid and structural parameters on the behaviour of fluid-structure coupled system. Static and dynamic features of the system, and their effect in aerodynamic properties, are described and compared for the different parameter combinations. The case with highest Reynolds number, Re = 10; 000, and intermediate elasticity was chosen as a base case to further study the fluid-structure coupling mechanism, particularly at low angle of attack conditions. The dynamic behaviour was characterised via spectral analysis in the frequency and wavenumber-frequency domains, which allowed the propagating wave nature of the membrane vibrations and their effect on the surrounding pressure field fluctuations to be clarified. The membrane vibrations are found to introduce upstream-propagating pressure waves that seem to be responsible for a loss in aerodynamic efficiency compared to a rigid aerofoil. Stability aspects of the FSI problem are also investigated by performing numerical experiments to analyse the response of the system to initial flow perturbations. The solutions of the 2D DNS are used as initial conditions for three-dimensional simulations, upon which initial perturbations in spanwise velocity are added. As the simulation is advanced in time the evolution of the perturbations is studied to determine the stability characteristics of the flow. Amplifications of the perturbations are found for Re > 10; 000. The coupling of the fully three-dimensional developed flow and the elastic aerofoil is also analysed with spectral techniques. Comparison of two- and three-dimensional results reveals that the three-dimensional flow development causes a decrease in the amplitude of the system fluctuations, but the same coupling mechanism found in the two-dimensional approach is also present in the three-dimensional case.
23
Rodriguez, Carlos G. "Viscous-inviscid interaction for incompressible flows over airfoils." Thesis, This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-09192009-040542/.
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Yeung, William Wai-Hung. "A mathematical model for airfoils with spoilers or split flaps." Thesis, University of British Columbia, 1985. http://hdl.handle.net/2429/25124.
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A flow model for a Joukowsky airfoil with an inclined spoiler or split flap is constructed based on the early work by Parkinson and Jandali. No restriction is imposed on the airfoil camber, the inclination and length of the spoiler or split flap, and the angle of incidence. The flow is assumed to be steady, two-dimensional, inviscid and incompressible. A sequence of conformal transformations is developed to deform the contour of the airfoil and the spoiler (split flap) onto the circumference of the unit circle over which the flow problem is solved. The partially separated flow region behind these bluff bodies is simulated by superimposing suitable singularities in the transform plane. The trailing edge, the tip of the spoiler (flap) are made critical points in the mappings so that Kutta conditions are satisfied there. The pressures at these critical points are matched to the pressure inside the wake, the only empirical input to the model. Some studies of an additional boundary condition for solving the flow problem were carried out with considerable success. The chordwise pressure distributions and the overall lift force variations are compared with experiments. Good agreement in general is achieved. The model can be extended readily to airfoils of arbitrary profile with the application of the Theodorsen transformation.Applied Science, Faculty of
Mechanical Engineering, Department of
Graduate
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Innes, Fraser. "An experimental investigation into the use of vortex generators to improve the performance of a high lift system." Thesis, City University London, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.307878.
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Zhu, Yu Ping. "Computational study of shock control at transonic speed." Thesis, Cranfield University, 2000. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.323930.
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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.
28
Moriarty, Julie Ann. "A nonlinear theory for thin aerofoils with non-thin trailing edges /." Title page, contents and summary only, 1987. http://web4.library.adelaide.edu.au/theses/09SM/09smm854.pdf.
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Dunn, Rob. "An aerodynamic analysis of aerofoils using a spline-velocity singularity method /." Thesis, McGill University, 1993. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=69712.
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This thesis presents a new approach based on a spline formulation for the analysis of thin aerofoils using the velocity singularity method. The method of velocity singularities was originally developed by Mateescu and Newman in conjunction with a polynomial representation of the normal perturbation velocities. The present method uses a cubic spline representation of the aerofoil contour, which led to improvements in the accuracy and stability of the solution, especially in the case of the jet-flapped aerofoils.This method has been first validated for the cases of rigid and flexible aerofoils. The pressure distributions obtained with the spline formulation have proven to be in good agreement with the previous solutions based on conformal transformation, or obtained by Thwaites, Nielsen, and by Mateescu and Newman.
The spline-velocity singularity method has been used for the jet flapped aerofoils and then extended to analyze the aerofoils with multiple sections, such as aerofoils with a flap. The solutions for these problems have been found to be in good agreement with the results obtained theoretically or experimentally by Spence, and Dimmock, and by Seebohm and Newman, based on a surface vortex method.
The spline-velocity singularity method displayed a better accuracy and an enhanced stability of the solution, in comparison with the polynomial formulation, especially in the cases when the aerofoil contour is not known a priori, such as for the flexible or jet-flapped aerofoils.
30
Hackenberg, Petra. "Numerical optimization of the suction distribution for laminar flow control aerofoils." Thesis, University of Southampton, 1994. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.241170.
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Fosas, De Pando Miguel Ángel. "Tonal noise generation in flows around aerofoils : a global stability analysis." Palaiseau, Ecole polytechnique, 2012. https://theses.hal.science/tel-00816987.
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La génération de fréquences discrètes dans l'écoulement autour d'un profil d'aile est étudiée dans cette thèse au moyen de simulations numériques non-linéaires et d'études de stabilité globale. À cette fin, un code numérique mettant en œuvre une nouvelle technique pour accéder à la dynamique linéaire, directe et adjointe, a d'abord été développé et ensuite appliqué à l'écoulement autour d'un profil. Les simulations non-linéaires confirment l'apparition de fréquences discrètes dans le spectre sonore, et pour le cas considéré, l'analyse de stabilité globale de la dynamique linéaire de l'écoulement moyen montre que celle-ci est stable. Cependant, la réponse de l'écoulement à des perturbations incidentes révèle de fortes croissances transitoires amenant à l'établissement de cycles de rétroaction aéroacoustique. Ces cycles comprennent la croissance des instabilités hydrodynamiques dans les couches limites sur l'intrados et l'extrados ainsi que leurs interactions avec les radiations acoustiques du bord de fuite. Les processus associés à l'apparition des fréquences discrètes dans le spectre sonore ainsi que les cycles de rétroaction sont ensuite mis en relation avec les modes globaux les moins stables: d'un coté, la structure spatiale des modes directs montre la croissance des instabilités hydrodynamiques sur l'extrados et la zone du sillage proche du bord de fuite; d'un autre coté, les modes adjoints associés présentent l'intrados comme la zone la plus réceptive à des perturbations externes. Finalement, l'analyse de la région dite du wavemaker indique, en accord avec les expériences, le rôle fondamental de la couche limite sur l'intradosThe generation of discrete acoustic tones in the compressible flow around an aerofoil is addressed in this thesis by means of nonlinear numerical simulations and global stability analyses. To this end, a nonlinear simulation code featuring a novel technique for gaining access to the linearized direct and adjoint dynamics has been developed and applied to the flow around an aerofoil. The nonlinear simulations confirm the appearance of discrete tones in the acoustic spectrum, and for the chosen flow case, the global stability analyses of the mean-flow dynamics reveal that the linearized operator is stable. However, the flow response to incoming disturbances exhibits important transient growth effects that culminate into the onset of aeroacoustic feedback loops, involving instability process on the suction- and pressure-surface boundary-layers together with their cross interaction by acoustic radiation at the trailing edge. The features of the aeroacoustic feedback loops and the appearance of discrete tones are then related to the features of the least stable modes in the global spectrum: on the one hand, the spatial structure of the direct modes display the growth of hydrodynamic instabilities on the suction surface and the near wake; on the other hand, the associated adjoint modes display increased receptivity of the flow on the pressure surface. Finally, the analysis of the wavemaker region highlights, in agreement with previous experimental investigations, the sensitivity of the flow to the pressure-surface boundary layer
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Linn, Anthony Blane. "Determination of average lift of a rapidly pitching airfoil." Link to electronic version, 2000. http://www.wpi.edu/Pubs/ETD/Available/etd-0512100-095442.
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Kim, Sangho. "Observation and measurements of flow structures in the stagnation region of a wing-body junction." Diss., This resource online, 1991. http://scholar.lib.vt.edu/theses/available/etd-08222008-063435/.
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Wilder, Michael C. "Airfoil-vortex interaction and the wake of an oscillating airfoil." Diss., This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-10022007-144516/.
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Santos, Luis Carlos de Castro. "A hybrid inverse optimization method for aerodynamic design of lifting surfaces." Diss., Georgia Institute of Technology, 1993. http://hdl.handle.net/1853/12105.
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Loretz, Yves Daniel. "Flow control on a NACA 4418 airfoil using streamwise synthetic jet actuators." Thesis, Georgia Institute of Technology, 2001. http://hdl.handle.net/1853/16377.
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Aves, Mark Antony. "Multigrid multiblock computation of steady compressible flows." Thesis, University of Reading, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.304393.
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Jeffrey, David Robert Michael. "An investigation into the aerodynamics of Gurney flaps." Thesis, University of Southampton, 1998. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.287300.
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Yeung, Chong Pui. "A study of three-dimensional interaction of wakes and boundary-layers." Thesis, University of Cambridge, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.319566.
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Khajeel, Hasan T. Abu. "Effect of humps on the stability of boundary layers over an airfoil /." This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-12042009-020055/.
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Luk, K. F. "Experiment on flow-induced vibration of an airfoil due to vortex shedding generated from upstream circular cylinder /." View Abstract or Full-Text, 2002. http://library.ust.hk/cgi/db/thesis.pl?MECH%202002%20LUK.
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Thesis (M. Phil.)--Hong Kong University of Science and Technology, 2002.Includes bibliographical references (leaves 54-57). Also available in electronic version. Access restricted to campus users.
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Starn, Andrew Douglas. "Doppler global velocimetry measurements in a wing flow field with tip blowing." Morgantown, W. Va. : [West Virginia University Libraries], 2004. https://etd.wvu.edu/etd/controller.jsp?moduleName=documentdata&jsp%5FetdId=24.
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Thesis (M.S.)--West Virginia University, 2004.Title from document title page. Document formatted into pages; contains vi, 101 p. : ill. (some col.). Includes abstract. Includes bibliographical references (p. 41-45).
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Ayton, Lorna Jane. "Asymptotic approximations for the sound generated by aerofoils in unsteady subsonic flows." Thesis, University of Cambridge, 2014. https://www.repository.cam.ac.uk/handle/1810/246549.
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This thesis considers the sound generated by unsteady perturbations interacting with solid aerofoils in background steady flows, in an attempt to further develop analytic models for the noise generated by blades within turboengines. Specifically, high-frequency unsteady gust and sound wave perturbations are considered and asymptotic results are obtained for, primarily, the far-field noise. Previous analytic work has examined high-frequency gust-aerofoil interactions in steady uniform flows using rapid distortion theory, and has focused on aerofoils with simple geometries. We extend this to deal with aerofoils with more realistic geometries (by including camber, thickness, and angle of attack), as well as considering the new topic of sound-aerofoil interactions in steady uniform flows for aerofoils with realistic geometries. The assumption of a steady uniform flow is later relaxed and we investigate the sound generated by high-frequency gust-aerofoil interactions in steady shear flows. Throughout all of the aforementioned work, the key process involves identifying various asymptotic regions around the aerofoil where different sources dominate the generation of sound. Solutions are obtained in each region and matched using the asymptotic matching rule. The dominant regions producing noise are the local, “inner”, regions at the leading and trailing edges of the aerofoil. Approximations for the far-field noise (in the “outer” regions) are the principal results, however one can also extract approximations for the unsteady pressure generated on the surface of the aerofoil. The surface pressure generated by high-frequency gust-aerofoil interaction in uniform flow is found to contain a singularity at the leading-edge stagnation point, thus the final piece of work in this thesis focuses more closely on turbulent interactions with solid body stagnation points in uniform flow, eliminating this singularity.
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Watt, Robert McFarlane. "Effects of surface roughness on the boundary-layer characteristics of turbine aerofoils." Thesis, University of Oxford, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.330065.
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Lin, Hequan. "Prediction of separated flows around pitching aerofoils using a discrete vortex method." Thesis, University of Glasgow, 1997. http://theses.gla.ac.uk/3950/.
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A surface shedding discrete vortex method has been developed for stimulating incompressible flows around pitching aerofoils. The method is able to predict both attached and separated flows, the latter typified by the formation and transport of large vortices. The structures of dynamic stall flow are well captured without the need for other means to predetermine the separation points. In contrast to most other vortex methods, the method presented herein can perform quantitative analysis. Throughout a wide range of incidence, the pressure distributions are smooth and the normal force and pitching moment are in good agreement with experimental data. The method is also able to predict the flow with external constraints for simulating the effects of wind tunnel blockage. In this regard quantitative results and flow structures have been obtained which are consistent with those expected. Following the review of previous work presented in the introduction, the mathematical formulation of the method is expounded. A velocity expression is theoretically derived for flows with both a moving inner boundary (aerofoil) and fixed external constraints (wind tunnel walls). To maintain both no penetration and no slip conditions, it is concluded that an external constraint parallel to the free stream can be modelled by placement of a constant vortex sheet along the boundary, and the introduction of distributed vortices next to the constraint to represent the boundary layer. The vortex sheet strength is equivalent to the free stream velocity while the strength of the vortices can be calculated in the same manner as for the internal boundary. This conclusion avoids the necessity of employing mirror vortices and iteration techniques in traditional models. The aerodynamic loads are computed from the pressure distribution. For the computation of surface pressures, the relationship between the pressure gradient and the rate of vorticity creation on the surface has been developed for a moving boundary.
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Stolcis, Luca. "Computation of the turbulent flow development around single- and multi-element aerofoils." Thesis, University of Manchester, 1992. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.316504.
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Elmstrom, Michael E. "Numerical prediction of the impact of non-uniform leading edge coatings on the aerodynamic performance of compressor airfoils." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2004. http://library.nps.navy.mil/uhtbin/hyperion/04Jun%5FElmstrom.pdf.
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Spanrad, Sven Klaus. "Fatigue crack growth in laser shock peened aerofoils subjected to foreign object damage." Thesis, University of Portsmouth, 2011. https://researchportal.port.ac.uk/portal/en/theses/fatigue-crack-growth-in-laser-shock-peened-aerofoils-subjected-to-foreign-object-damage(b367cb9f-b746-4c27-9479-49cd48999519).html.
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Foreign Object Damage (FOD) is one of the main life limiting factors for aeroengine fan blades. The FOD impacts during takeoff and landing cause severe damage to aerofoils, resulting in reduced air safety and life time with an estimated annual cost of $4 billion for the aeroengine industry. Advanced surface treatments, such as Laser Shock Peening (LSP) have significantly improved the fatigue strength and crack growth resistance of critical components under FOD. However, it is not yet possible to predict the protective residual stresses and utilise their full potential for enhancing fatigue resistance and damage tolerance capacity in service. This research programme aims to utilise some of the established methods for fatigue tolerance assessment of critical components, based on fracture mechanics principles, to address the effects of complex residual stresses due to LSP and FOD on fatigue crack growth in aerofoils under simulated service loading conditions. The experimental study involved fatigue testing of LSPed and FODed specimens with a geometry representative of fan blades made from Ti-6Al-4V alloy. A four point bend fatigue test setup was designed and calibrated. A real-time computer-controlled crack growth monitoring system and optical crack monitoring techniques were developed. Scanning Electron Microscopy (SEM) and Back-Scatter Electron (BSE) were used to conduct metallographic and fractographic studies, including crack initiation, early fatigue crack growth and FOD damage characterisation. The fracture mechanics analyses used the weight function method and the finite element method to obtain a modified stress intensity factor considering residual stresses due to LSP and FOD. Fatigue crack growth data under low cycle fatigue(LCF), high cycle fatigue (HCF) and combined LCF and HCF loading conditions were correlated using a standard and the modified stress intensity factors. The influence of impact angles and loading conditions on fatigue crack growth behaviour was assessed, and the results were compared with those from untreated FODed specimens.
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Kurtulus, Berkin. "Development of a Tool for Inverse Aerodynamic Design and Optimisation of Turbomachinery Aerofoils." Thesis, KTH, Flygdynamik, 2021. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-293353.
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The automation of airfoil design process is an ongoing effort within the field of turbo-machinery design, with significant focus on developing new reliable and consistent methods that can meet the needs of the engineers. A wide variety of approaches has been in use for inverse airfoil design process which benefit from theoretical inverse design, statistical methods, empirical discoveries and many other ways to solve the design problem. This thesis work also develops a tool in Python to be used in airfoil aerodynamic design process that is simple, fast and accurate enough for initial design of turbo-machinery blades with focus on turbine airfoils used for operation in aircraft engines. To convey the decision-making process during development a simplified case is presented. The underlying considerations are discussed. Other available methods in the literature used for similar problems, are also evaluated and compared to demonstrate the advantages and limitations of the methods used within the tool. The inverse design problem is formulated as a multi-objective optimization problem to handle various different objectives that are relevant for aerodynamic design of turbo-machinery airfoils. Test runs are made and the results are discussed to assess how robust the tool is and how the current capabilities can be modified or extended. After the development process, the tool is verified to be a suitable option for real-life design optimization tasks and can be used as a building block for a much more comprehensive tool that may be developed in the future.Automatisering av processen för design av vingprofiler kräver fortlöpande insatser inom området turbomaskindesign, med stort fokus på att utveckla nya tillförlitliga och konsekventa metoder som kan tillgodose ingenjörernas behov. Ett stort antal olika tillvägagångssätt har provats för omvänd design av vingprofiler såsom teoretisk invers design, statistiska metoder, empiriska upptäckter och många andra sätt att lösa designproblemet. Detta avhandlingsarbete är också ett lyckat försök att utveckla ett verktyg i Python som ska användas i den aerodynamiska designprocessen; det är enkelt, snabbt och noggrant för den initiala designen av turbomaskinblad med fokus på turbinblad som för användning i flygmotorer. För att förmedla beslutsprocessen under utvecklingen presenteras ett förenklat fall. De underliggande övervägandena diskuteras. Andra tillgängliga metoder i litteraturen som används för liknande problem utvärderas och jämförs för att visa fördelarna och begränsningarna med de metoder som används i verktyget. Det omvända designproblemet formuleras som ett multi-objektivt optimeringsproblem för att hantera olika mål som är relevanta för aerodynamisk design av turbomaskiner. Testkörningar görs och resultaten diskuteras för att bedöma hur robust verktyget är och hur de nuvarande funktionerna kan modifieras eller utökas. Efter utvecklingsprocessen verifieras verktyget som ett lämpligt alternativ för verkliga designoptimeringsuppgifter och kan användas som en byggsten för ett mycket mer omfattande verktyg som kan utvecklas i framtiden.
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Allan, William D. E. "An experimental study of flow about an airfoil with slotted flap and spoiler using Joukowsky profiles." Thesis, University of British Columbia, 1988. http://hdl.handle.net/2429/28363.
Full textAbstract:
An experimental study has been carried out on an airfoil with slotted flap and spoiler using Joukowsky profiles. Pressure distributions
were measured as functions of angle of attack, flap deflection angle, spoiler size and inclination. The results are uncorrected for wind tunnel wall effects but the data base is available to carry out the corrections. The results will be used to compare with predictions of a theoretical model, yet to be worked out, which combines work previously done by Williams, Jandali, Parkinson and Yeung. This theory will involve the potential flow about a two-element 'near'-Joukowsky airfoil system. The secondary airfoil is a simulated slotted flap. Various size spoilers are introduced to the system at arbitrary angles of inclination using methods proposed by Parkinson and Yeung.
The experimental results are qualitatively reasonable and some interesting effects are observed. The behaviour of spoilers, when used with slotted flaps at various deflection angles, corresponds well with requirements of aircraft in approach or landing situations. Similarly, the use of slotted flaps alone provides the high lift at low angle of attack which is beneficial to aircraft taking off. Some recommendations are proposed for further testing with this equipment.Applied Science, Faculty of
Mechanical Engineering, Department of
Graduate