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Dissertations / Theses on the topic 'Airfoils'
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1
Mahmoud, Osama Mohamed Kamal Mohamed. "Experimental investigation of low speed flow over flapping airfoils and airfoil combinations." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2001. http://handle.dtic.mil/100.2/ADA406240.
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Dissertation (Ph.D. in Aeronautical and Astronautical Engineering)--Naval Postgraduate School, Sept. 2001.Dissertation supervisor: Platzer, Max F. "September 2001." Includes bibliographical references (p. 171-174). Also available in print.
2
Yeung, William Wai-Hung. "Modelling stalled airfoils." Thesis, University of British Columbia, 1990. http://hdl.handle.net/2429/31120.
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The thesis deals with some new applications of the wake source model, a two-dimensional incompressible potential flow model used for bodies experiencing flow separation. The body contour is conformally mapped to a circle, for which the flow problem is solved using source singularities to create free streamlines simulating the separating shear layers. In common with other inviscid theories, it generally requires the pressure in the separated flow region, and the location of separation if boundary-layer controlled.
Different mapping sequences and flow models have been constructed for the following five problems,
1. the trailing-edge stall for single element airfoils,
2. flat plates with separation bubbles,
3. separation bubbles upstream of spoilers with downstream wakes,
4. spoiler/slotted flap combinations, at which the spoiler inclination is arbitrary, and
5. two-element airfoils near (trailing-edge) stall.
Predictions of pressure distribution are compared with wind tunnel measurements, and good agreement is found in cases 1 and 5. The initial shape of the separation streamlines also appears to be satisfactory. Results in cases 2 and 3 are promising although more work is needed to improve the bubble shapes and their pressure distributions. Partial success has been achieved on spoiler/ slotted flap configurations, depending on the spoiler inclination. For strong wake effect on the flap (e. g. δ = 90° ), the model predicts a very high suction peak over it. Whereas the experimental data resemble a stalled distribution even though flow visualization indicates the flap to be unstalled. This may be related to a limitation of the method, also noted in the separation-bubble problems, that it cannot specify a complete boundary condition on a free streamline. This discrepancy diminishes as the spoiler angle becomes smaller (e. g. δ = 30° ) in the cases of higher incidences so that the wake boundary tugs away from the flap sooner.Applied Science, Faculty of
Mechanical Engineering, Department of
Graduate
3
Knappskog, Håvard. "Nonlinear control of Tethered Airfoils : Path-following control of Tethered Airfoils." Thesis, Norges teknisk-naturvitenskapelige universitet, Institutt for teknisk kybernetikk, 2011. http://urn.kb.se/resolve?urn=urn:nbn:no:ntnu:diva-13458.
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This master thesis contains modeling, analysis and control design for atethered airfoil. A path-following controller has been developed and provenlocally asymptotically stable. The guidance law is general, and applicableto other path-following systems. The closed loop system is demonstratedin simulations, where a certain level of robustness is concluded.
4
Pope, Orrin Dean. "Aerodynamic Centers of Arbitrary Airfoils." DigitalCommons@USU, 2017. https://digitalcommons.usu.edu/etd/6890.
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The study of designing stable aircraft has been widespread and ongoing since the early days of Orville and Wilbur Wright and their famous Wright Flyer airplane. All aircraft as they fly through the air are subject to minor changes in the forces acting on them. The field of aircraft stability seeks to understand and predict how aircraft will respond to these changes in forces and to design aircraft such that when these forces change the aircraft remains stable. The mathematical equations used to predict aircraft stability rely on knowledge of the location of the aerodynamic center, the point through which aerodynamic forces act on an aircraft. The aerodynamic center of an aircraft is a function of the aerodynamic centers of each individual wing, and the aerodynamic center of each wing is a function of the aerodynamic centers of the individual airfoils from which the wing is made. The ability to more accurately predict the location of the airfoil aerodynamic center corresponds directly to an increase in the accuracy of aircraft stability calculations.
The Aerolab at Utah State University has develop new analytic mathematical expressions to describe the location of the airfoil aerodynamic center. These new expressions do not suffer from any of the restrictions, or approximations found in traditional methods, and therefore result in more accurate predictions of airfoil aerodynamic centers and by extension, more accurate aircraft stability predictions.
5
Alexandris, Georgios. "Supersonic flow past two oscillating airfoils." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1998. http://handle.dtic.mil/100.2/ADA350226.
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Thesis (M.S. in Applied Physics) Naval Postgraduate School, June 1998."June 1998." Thesis advisor(s): Max F. Platzer, James H. Luscombe, S. Weber. Includes bibliographical references (p. 71-72). Also available online.
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Shrewsbury, George D. "Dynamic stall of circulation control airfoils." Diss., Georgia Institute of Technology, 1990. http://hdl.handle.net/1853/12397.
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Neculita, Catalin Silviu. "Unsteady compressible flows past oscillating airfoils." Thesis, McGill University, 2005. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=99002.
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This thesis presents new and efficient analytical solutions for the unsteady subsonic compressible flows past rigid and flexible airfoils executing low frequency oscillations. These solutions are obtained using an especially developed method based on velocity singularities associated with the airfoil leading edge and ridges, which define the changes in the airfoil boundary conditions. The velocity singularity method has been initially developed by Mateescu.Closed form solutions are presented for the unsteady lift and pitching moment coefficients and for the chordwise distribution of the unsteady pressure difference coefficient in the general case of rigid airfoils executing oscillatory pitching rotations and translations, as well as for flexible airfoils executing flexural oscillations.
For the case of incompressible flows, the present solutions were found in excellent agreement with the previous incompressible flow results obtained by Theodorsen, Postel & Leppert and by Mateescu & Abdo.
8
Garzón, Víctor E. 1972. "Probabilistic aerothermal design of compressor airfoils." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/16995.
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Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.Includes bibliographical references (p. 175-183).
This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Despite the generally accepted notion that geometric variability is undesirable in turbomachinery airfoils, little is known in detail about its impact on aerothermal compressor performance. In this work, statistical and probabilistic techniques were used to assess the impact of geometric and operating condition uncertainty on axial compressor performance. High-fidelity models of geometric variability were constructed from surface measurements of existing hardware using principal component analysis (PCA). A quasi-two-dimensional cascade analysis code, at the core of a parallel probabilistic analysis framework, was used to assess the impact of uncertainty on aerodynamic performance of compressor rotor airfoils. Three rotor blades with inlet relative Mach numbers of 0.82, 0.90 and 1.25 were considered. Discrepancies between nominal and mean loss (mean-shift) of up to 20% were observed. Loss and turning variability were found to grow linearly with geometric noise amplitude. A probabilistic, gradient-based approach to compressor blade optimization was presented. Probabilistic objectives, constraints and gradients are approximated using low-resolution Monte Carlo sampling. Test airfoils were optimized both deterministically and probabilistically and then analyzed probabilistically to account for geometric variability. Probabilistically redesigned airfoils exhibited reductions in mean loss of up to 25% and in loss variability of as much as 65% from corresponding values for deterministically redesigned airfoils.
(cont.) A probabilistic mean-line multi-stage axial compressor model was used to estimate the impact of geometric variability on overall compressor performance. Probabilistic loss and turning models were exercised on a six-stage compressor model. At realistic levels of geometric variability, the mean polytropic efficiency was found to be upwards of 1% lower than nominal. Compressor simulations using airfoils redesigned probabilistically for minimum loss variability exhibited reductions of 30 to 40% in polytropic efficiency variability and mean shift.
by Victor E. Garzon.
Ph.D.
9
Walker, William Paul. "Unsteady Aerodynamics of Deformable Thin Airfoils." Thesis, Virginia Tech, 2009. http://hdl.handle.net/10919/34620.
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Unsteady aerodynamic theories are essential in the analysis of bird and insect flight. The study of these types of locomotion is vital in the development of flapping wing aircraft. This paper uses potential flow aerodynamics to extend the unsteady aerodynamic theory of Theodorsen and Garrick (which is restricted to rigid airfoil motion) to deformable thin airfoils. Frequency-domain lift, pitching moment and thrust expressions are derived for an airfoil undergoing harmonic oscillations and deformation in the form of Chebychev polynomials. The results are validated against the time-domain unsteady aerodynamic theory of Peters. A case study is presented which analyzes several combinations of airfoil motion at different phases and identifies various possibilities for thrust generation using a deformable airfoil.Master of Science
10
Joyce, Richard Kirk. "A method of testing two-dimensional airfoils." Thesis, Monterey, California. Naval Postgraduate School, 1992. http://hdl.handle.net/10945/23721.
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Garcia, Oscar Mauricio Arias. "Numerical simulations of compressible flows over airfoils." Instituto Tecnológico de Aeronáutica, 2006. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=316.
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A computer code was developed from scratch to simulate the flow over the NACA 0012 airfoil at different Reynolds and Mach numbers. The domain was discretized in a structured-grid context. The equations were numerically solved by a finite-volume technique, using three different time-marching schemes. The Euler flow was initially modeled as well as a Reynolds-averaged Navier Stokes formulation was calculated. The Baldwin and Lomax turbulence model was employed to close the problem. The influence of a number of numerical parameters upon the computational solutions was investigated in the first phase of the work. The inviscid simulations were compared with other numerical results available in the literature. Each modification is thoroughly described and compared to the base-line case. Conclusions were drawn regarding how each of these chances affected the final result. The last Euler simulation was done using the Jameson, MacCormack and the Shu schemes in order to select the most appropriate one of the three to be employed to solve the Reynolds-averaged Navier Stokes equations. The viscous flow simulations started with the incompressible, laminar flow over a flat plate. The implementation of the viscous terms was validated calculating and comparing the results with the known Blasius analytical solution. Finally, the compressible, turbulent viscous flow over the NACA 0012 airfoil was numerically solved. The pressure coefficient distribution along the airfoil chord and the normal force coefficient were compared with experimental data due to Harris.
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Lang, Jr Joseph Reagle. "Characterization of Oscillatory Lift in MFC Airfoils." Thesis, Virginia Tech, 2014. http://hdl.handle.net/10919/50935.
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The purpose of this research is to characterize the response of an airfoil with an oscillatory morphing, Macro-fiber composite (MFC) trailing edge. Correlation of the airfoil lift with the oscillatory input is presented. Modal analysis of the test airfoil and apparatus is used to determine the frequency response function. The effects of static MFC inputs on the FRF are presented and compared to the unactuated airfoil.
The transfer function is then used to determine the lift component due to cambering and extract the inertial components from oscillating airfoil. Finally, empirical wind tunnel data is modeled and used to simulate the deflection of airfoil surfaces during dynamic testing conditions. This research serves to combine modal analysis, empirical modeling, and aerodynamic testing of MFC driven, oscillating lift to formulate a model of a dynamic, loaded morphing airfoil.Master of Science
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Turner, Kevin E. "Stiffness Characteristics of Airfoils Under Pulse Loading." The Ohio State University, 2009. http://rave.ohiolink.edu/etdc/view?acc_num=osu1259113516.
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Skare, Steven Edward. "An Inverse Design Method for Supersonic Airfoils." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/731.
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Airfoil design is one of the most important aspects of aircraft design. Slight changes in airfoil geometry can lead to significant changes in a wide variety of aircraft performance metrics. Inverse design methods offer an efficient alternative to standard direct methods. The key to this design problem is to derive a direct relationship between changes in airfoil geometry and changes in pressure or velocity distributions. This relationship is then used to modify an initial airfoil and its pressure distribution to match a target pressure distribution, which is specified by design parameters. At this point, the engineer now has a final airfoil based off of the design requirements.
This paper attempts to provide a quick and easy inverse design method for a wide variety of supersonic scenarios. This is accomplished by using the class-shape transformation technique to parameterize airfoils during an iterative process. The robustness of the method is demonstrated through several distinct design cases including supersonic airfoils, unique geometries, and a Sears-Haack body.
15
Tuncer, Ismail H. "Unsteady aerodynamics of oscillating and rapidly pitched airfoils." Diss., Georgia Institute of Technology, 1988. http://hdl.handle.net/1853/12522.
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Wu, Jiunn-Chi. "A study of unsteady turbulent flow past airfoils." Diss., Georgia Institute of Technology, 1988. http://hdl.handle.net/1853/13091.
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Zope, Anup Devidas. "Response surface analysis of trapped-vortex augmented airfoils." Thesis, Mississippi State University, 2015. http://pqdtopen.proquest.com/#viewpdf?dispub=1604198.
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In this study, the effect of a passive trapped-vortex cell on lift to drag (L/D) ratio of an FFA-W3-301 airfoil is studied. The upper surface of the airfoil was modified to incorporate a cavity defined by seven parameters. The L/D ratio of the airfoil is modeled using a radial basis function metamodel. This model is used to find the optimal design parameter values that give the highest L/D. The numerical results indicate that the L/D ratio is most sensitive to the position on an airfoil’s upper surface at which the cavity starts, the position of the end point of the cavity, and the vertical distance of the cavity end point relative to the airfoil surface. The L/D ratio can be improved by locating the cavity start point at the point of separation for a particular angle of attack. The optimal cavity shape (o19_aXX) is also tested for a NACA0024 airfoil.
18
Joh, Changyeol. "Efficient and robust design optimization of transonic airfoils." Diss., Virginia Tech, 1991. http://hdl.handle.net/10919/39983.
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Numerical optimization procedures have been employed for the design of airfoils in transonic flow based on the transonic small-disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented for solving the design problem of lift maximization with wave drag and area constraints. A simple linear approximation was utilized for the approximation of the lift. Accurate approximations for sensitivity derivatives of the wave drag were obtained through the utilization of Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without recreating the grid. Our design procedures experienced convergence problems for some TSD solutions, where the wave drag was found not to vary smoothly with the design parameters and consequently create local optimum problems. A procedure interchanging the role of the objective function and constraint, initially minimizing drag with a constraint on the lift was found to be effective in producing converged designs, usually in approximately 10 global iterations. This procedure was also shown to be robust and efficient for cases where the drag varied smoothly, such as with the Euler solutions. The direct lift maximization with move limits which were fixed absolute values rather than fractions of the design variables, was also found to be a reliable and efficient procedure for designs based upon the Euler equations.Ph. D.
19
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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Chao, David Dai-Wei. "Drag prediction and decomposition for airfoils and wings /." For electronic version search Digital dissertations database. Restricted to UC campuses. Access is free to UC campus dissertations, 2004. http://uclibs.org/PID/11984.
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Neace, Kerry S. "A computational and experimental investigation of the propulsive and lifting characteristics of oscillating airfoils and airfoil combinations in incompressible flow." Thesis, Monterey, California. Naval Postgraduate School, 1992. http://hdl.handle.net/10945/23578.
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Approved for public release; distribution is unlimitedComputational and experimental methods have been used to systematically study one and two airfoils undergoing unsteady motion. first, a single airfoil analysis with the modified computer code, U2DIFF. thrust, efficiency, and phase relationships were computed and compared to existing theoretical results. Furthermore, to help understand the dynamic stall process, relationships were developed between steady and quasi-steady pressure distributions for an airfoil undergoing a ramp motion. Next, an unsteady analysis for two airfoils was done with the modified computer code USPOTF2. Again, thrust and efficiencies for interfering, harmonically oscillating airfoils were computer and compared to existing theoretical results. Furthermore, an analysis was completed on the effects of a harmonically oscillating airfoil on the pressure gradient of a stationary airfoil. Finally, flow visualization experiments were conducted using a low smoke speed tunnel at the Naval Postgraduate School (NPS). This experiment demonstrated the effects of a thrust producing, oscillating airfoil on the formation of the wake vortices. Furthermore, a flow visualization experiment was conducted in the NPS low speed wind tunnel, which demonstrated the beneficial influence of a secondary airfoil oscillating in the vicinity of a stationary airfoil at high-angle-of-attack.
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Young, John Aerospace Civil & Mechanical Engineering Australian Defence Force Academy UNSW. "Numerical simulation of the unsteady aerodynamics of flapping airfoils." Awarded by:University of New South Wales - Australian Defence Force Academy. School of Aerospace, Civil and Mechanical Engineering, 2005. http://handle.unsw.edu.au/1959.4/38656.
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There is currently a great deal of interest within the aviation community in the design of small, slow-flying but manoeuvrable uninhabited vehicles for reconnaissance, surveillance, and search and rescue operations in urban environments. Inspired by observation of birds, insects, fish and cetaceans, flapping wings are being actively studied in the hope that they may provide greater propulsive efficiencies than propellers and rotors at low Reynolds numbers for such Micro-Air Vehicles (MAVs). Researchers have posited the Strouhal number (combining flapping frequency, amplitude and forward speed) as the parameter controlling flapping wing aerodynamics in cruising flight, although there is conflicting evidence. This thesis explores the effect of flapping frequency and amplitude on forces and wake structures, as well as physical mechanisms leading to optimum propulsive efficiency. Two-dimensional rigid airfoils are considered at Reynolds number 2,000 ??? 40,000. A compressible Navier-Stokes simulation is combined with numerical and analytical potential flow techniques to isolate and evaluate the effect of viscosity, leading and trailing edge vortex separation, and wake vortex dynamics. The wake structures of a plunging airfoil are shown to be sensitive to the flapping frequency independent of the Strouhal number. For a given frequency, the wake of the airfoil exhibits ???vortex lock-in??? as the amplitude of motion is increased, in a manner analogous to an oscillating circular cylinder. This is caused by interaction between the flapping frequency and the ???bluff-body??? vortex shedding frequency apparent even for streamlined airfoils at low Reynolds number. The thrust and propulsive efficiency of a plunging airfoil are also shown to be sensitive to the flapping frequency independent of Strouhal number. This dependence is the result of vortex shedding from the leading edge, and an interaction between the flapping frequency and the time for vortex formation, separation and convection over the airfoil surface. The observed propulsive efficiency peak for a pitching and plunging airfoil is shown to be the result of leading edge vortex shedding at low flapping frequencies (low Strouhal numbers), and high power requirements at large flapping amplitudes (high Strouhal numbers). The efficiency peak is governed by flapping frequency and amplitude separately, rather than the Strouhal number directly.
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Tamai, Masatoshi. "Experimental investigations on biologically inspired airfoils for MAV applications." [Ames, Iowa : Iowa State University], 2007.
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Nunes, Bonaventure R. "Numerical Loss Prediction of high Pressure Steam Turbine airfoils." Thesis, Virginia Tech, 2013. http://hdl.handle.net/10919/51742.
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Steam turbines are widely used in various industrial applications, primarily for power extraction. However, deviation for operating design conditions is a frequent occurrence for such machines, and therefore, understanding their performance at off design conditions is critical to ensure that the needs of the power demanding systems are met as well as ensuring safe operation of the steam turbines. In this thesis, the aerodynamic performance of three different turbine airfoil sections ( baseline, mid radius and tip profile) as a function of angle of incidence and exit Mach numbers, is numerically computed at 0.3 axial chords downstream of the trailing edge. It was found that the average loss coefficient was low, owing to the fact that the flow over the airfoils was well behaved. The loss coefficient also showed a slight decrease with exit Mach number for all three profiles. The mid radius and tip profiles showed near identical performance due to similarity in their geometries. It was also found out that the baseline profile showed a trend of substantial increase in losses at positive incidences, due to the development of an adverse pressure zone on the blade suction side surface. The mid radius profile showed high insensitivity to angle of incidence as well as low exit flow angle deviation in comparison to the baseline blade.Master of Science
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Garmann, Daniel J. "High-Fidelity Simulations of Transitional Flow Over Pitching Airfoils." University of Cincinnati / OhioLINK, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1276955868.
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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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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
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Matouk, Rabea. "Calculation of Aerodynamic Noise of Wing Airfoils by Hybrid Methods." Doctoral thesis, Universite Libre de Bruxelles, 2016. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/240641.
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This research is situated in the field of Computational AeroAcoustics (CAA). The thesis focuses on the computation of the aerodynamic noise generated by turbulent flows around wing, fan, or propeller airfoils. The computation of the noise radiated from a device is the first step for designers to understand the acoustical characteristics and to determine the noise sources in order to modify the design toward having acoustically efficient products. As a case study, the broadband or trailing-edge noise emanating from a CD (Controlled-Diffusion) airfoil, belonging to a fan is studied. The hybrid methods of aeroacoustic are applied to simulate and predict the radiated noise. The necessary tools were researched and developed. The hybrid methods consist in two steps simulations, where the determination of the aerodynamic field is decoupled from the computation of the acoustic waves propagation to the far field, so the first part of this thesis is devoted to an aerodynamic study of the considered airfoil. In this part of the thesis, a complete aerodynamic study has been performed. Some aspects have been developed in the used in-house solver SFELES, including the implementation of a new SGS model, a new outlet boundary condition and a new transient format which is used to extract the noise sources to be exported to the acoustic solver, ACTRAN. The second part of this thesis is concerned with the aeroacoustic study where four methods have been applied, among them two are integral formulations and the two others are partial-differential equations. The first method applied is Amiet’s theory, implemented in Matlab, based on the wall-pressure spectrum extracted in a point near the trailing edge. The second method is Curle’s formulation. It is applied proposing two approaches; the first approach is the implementation of the volume and surface integrals in SFELES to be calculated simultaneously with the flow in order to avoid the storage of noise sources which requires a huge space. In the second approach, the fluctuating aerodynamic forces, already obtained during the aerodynamics simulation, are used to compute the noise considering just the surface sources. Finally, Lighthil and Möhring analogies have been applied via the acoustic solver ACTRAN using sources extracted via SFELES. Maps of the radiated noise are demonstrated for several frequencies. The refraction effects of the mean flow have been studied.Doctorat en Sciences de l'ingénieur et technologie
info:eu-repo/semantics/nonPublished
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Ko, Lok Sun. "An experimental investigation of airfoils with perforated Gurney-type flaps." Thesis, McGill University, 2011. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=96955.
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The Gurney flap can provide an increase in lift with a small drag penalty up to a flap height of approximately 3% of the chord. For flap height larger than 3% chord (herein referred to as Gurney-type flap), there is a significant increase in drag, causing a deteriorating lift-to-drag ratio. In this work, perforation was introduced in large Gurney-type flaps in an attempt to improve their aerodynamic performance. A study of different flap heights and perforation porosities were performed by using force balance, surface pressure and hot-wire measurements in conjunction with particle image velocimetry (PIV). Results show that flap porosity reduced lift in comparison to the solid flap due to the decrease in positive camber effects and decompression of the cavity flow up stream of the flap. The corresponding reduction in drag, however, outweighed the loss in lift and rendered an improved lift-to-drag ratio. Detailed PIV flow field behind the perforated flap revealed that the existence of perforation-generated jets are responsible for the observed differences in aerodynamic performance. If strong enough, perforation- generated jets could eliminate the vortex shedding process behind the flap. Furthermore, the near wake was found to be disrupted and narrowed, drastically suppressing fluctuation intensity.Le volet de Gurney peut augmenter la force de sustentation (portance) avec peu de pénalité sur la traînée aérodynamique jusqu'à un hauteur de volet d'environ 3% de la corde du profil. Pour les volets plus hauts que 3% de la corde (appelles "Gurney-type flaps" ici), il y a une croissance significative à la traînée aérodynamique, causant une détérioration de la finesse (rapport portance/traînée). Dans cette étude, la perforation a été introduite dans les grands volets de Gurney afin d'essayer d'améliorer la finesse. Différentes hauteurs de volet et porosités de perforation ont été étudiées par balance de force, pression superficielle, fil chaud, et vélocimétrie par image de particules (PIV). Les résultats démontrent que la perforation des volets réduit la portance comparés aux cas non-perforés à cause de la diminution d'effet de cambrure positive et de la décompression de l'écoulement de la cavité en amont du volet. La réduction correspondante de la traînée, cependant, était supérieure à la perte de la portance et a amélioré la finesse. Les champs d'écoulement détaillés obtenus par PIV en aval du volet ont également indiqué que l'existence de jets d'air générés par la perforation sont responsable des différences observées à la finesse. Les jets générés par la perforation, si assez forts, peuvent éliminer le processus de décollement de tourbillon en aval du volet. En outre, le sillage proche s'est avéré perturbé et rétréci, supprimant l'intensité de fluctuation.
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Jackson, Keith S. (Keith Stuart). "CAD-casting of gas turbine airfoils using three dimensional printing." Thesis, Massachusetts Institute of Technology, 1997. http://hdl.handle.net/1721.1/10518.
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Merchant, Ali A. (Ali Abbas). "Design and analysis of supercritical airfoils with boundary layer suction." Thesis, Massachusetts Institute of Technology, 1996. http://hdl.handle.net/1721.1/10987.
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Bilgen, Onur. "Aerodynamic and Electromechanical Design, Modeling and Implementation Of Piezocomposite Airfoils." Diss., Virginia Tech, 2010. http://hdl.handle.net/10919/28665.
Full textAbstract:
Piezoelectrics offer high actuation authority and sensing over a wide range of frequencies. A Macro-Fiber Composite is a type of piezoelectric device that offers structural flexibility and high actuation authority. A challenge with piezoelectric actuators is that they require high voltage input; however the low power consumption allows for relatively lightweight electronic components. Another challenge, for piezoelectric actuated aerodynamic surfaces, is found in operating a relatively compliant, thin structure (desirable for piezoceramic actuators) in situations where there are relatively high external (aerodynamic) forces. Establishing an aeroelastic configuration that is stiff enough to prevent flutter and divergence, but compliant enough to allow the range of available motion is the central challenge in developing a piezocomposite airfoil. The research proposed here is to analyze and implement novel electronic circuits and structural concepts that address these two challenges.
Here, a detailed theoretical and experimental analysis of the aerodynamic and electromechanical systems that are necessary for a practical implementation of a piezocomposite airfoil is presented. First, the electromechanical response of Macro-Fiber Composite based unimorph and bimorph structures is analyzed. A distributed parameter electromechanical model is presented for interdigitated piezocomposite unimorph actuators. Necessary structural features that result in large electrically induced deformations are identified theoretically and verified experimentally. A novel, lightweight electrical circuitry is proposed and implemented to enable the peak-to-peak actuation of Macro-Fiber Composite bimorph devices with asymmetric voltage range.
Next, two novel concepts of supporting the piezoelectric material are proposed to form two types of variable-camber aerodynamic surfaces. The first concept, a simply-supported thin bimorph airfoil, can take advantage of aerodynamic loads to reduce control input moments and increase control effectiveness. The structural boundary conditions of the design are optimized by solving a coupled fluid-structure interaction problem by using a structural finite element method and a panel method based on the potential flow theory for fluids. The second concept is a variable-camber thick airfoil with two cascading bimorphs and a compliant box mechanism. Using the structural and aerodynamic theoretical analysis, both variable-camber airfoil concepts are fabricated and successfully implemented on an experimental ducted-fan vehicle. A custom, fully automated low-speed wind tunnel and a load balance is designed and fabricated for experimental validation. The airfoils are evaluated in the wind tunnel for their two-dimensional lift and drag coefficients at low Reynolds number flow. The effects of piezoelectric hysteresis are identified.
In addition to the shape control application, low Reynolds number flow control is examined using the cascading bimorph variable-camber airfoil. Unimorph type actuators are proposed for flow control in two unique concepts. Several electromechanical excitation modes are identified that result in the delay of laminar separation bubble and improvement of lift. Periodic excitation to the flow near the leading edge of the airfoil is used as the flow control method. The effects of amplitude, frequency and spanwise distribution of excitation are determined experimentally using the wind tunnel setup.
Finally, the effects of piezoelectric hysteresis nonlinearity are identified for Macro-Fiber Composite bimorphs. The hysteresis is modeled for open-loop response using a phenomenological classical Preisach model. The classical Preisach model is capable of predicting the hysteresis observed in 1) two cantilevered bimorph beams, 2) the simply-supported thin airfoil, and 3) the cascading bimorph thick airfoil.Ph. D.
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Loo, Felipe Manuel. "Numerical Study of Limit Cycle Oscillation Using Conventional and Supercritical Airfoils." Scholarly Repository, 2008. http://scholarlyrepository.miami.edu/oa_theses/176.
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Limit Cycle Oscillation is a type of aircraft wing structural vibration caused by the non-linearity of the system. The objective of this thesis is to provide a numerical study of this aeroelastic behavior. A CFD solver is used to simulate airfoils displaying such an aeroelastic behavior under certain airflow conditions. Two types of airfoils are used for this numerical study, including the NACA64a010 airfoil, and the supercritical NLR 7301 airfoil. The CFD simulation of limit cycle oscillation (LCO) can be obtained by using published flow and structural parameters. Final results from the CFD solver capture LCO, as well as flutter, behaviors for both wings. These CFD results can be obtained by using two different solution schemes, including the Roe and Zha scheme. The pressure coefficient and skin friction coefficient distributions are computed using the CFD results for LCO and flutter simulations of these two airfoils, and they provide a physical understanding of these aeroelastic behaviors.
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Kaya, Mustafa. "Path Optimization Of Flapping Airfoils Based On Unsteady Viscous Flow Solutions." Phd thesis, METU, 2008. http://etd.lib.metu.edu.tr/upload/2/12609349/index.pdf.
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The flapping path of a single airfoil and dual airfoils in a biplane configuration is optimized for maximum thrust and/or propulsive efficiency. Unsteady, low speed viscous flows are computed using a Navier-Stokes solver in a parallel computing environment. A gradient based algorithm and Response Surface Methodology (RSM) are employed for optimization. The evaluation of gradient vector components and the design of experiments for RSM, which require unsteady solutions, are also carried out in parallel. Parallel computations are performed using Parallel Virtual Machine (PVM) library. First, a single airfoil undergoing a combined sinusoidal or non-sinusoidal pitching and plunging motion is studied. The non-sinusoidal flapping motion is described using an elliptic curve or Non-Uniform Rational B-Splines
(NURBS). It is shown that the thrust generation may significantly
be increased in comparison to the sinusoidal flapping motion. For
a high thrust, the airfoil stays at high effective angle of attack
values during the upstroke and the downstroke, and the effective pitching occurs at minimum and maximum plunge positions. Secondly, the optimization of sinusoidal and non-sinusoidal flapping paths of dual airfoils is considered. Moving and deforming overset grids are used for computations. The deforming overset grids remove the restrictions on the flapping motion, and improve the optimization results obtained earlier. At low flapping frequencies, an airfoil in a biplane configuration produces more thrust than a single airfoil. Yet, at high frequencies the airfoil in biplane configuration produces less thrust at a significantly lower efficiency than the single airfoil.
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Thuné, Sebastian, and Torstein Soland. "Investigation Of Different Airfoils on Outer Sections of Large Rotor Blades." Thesis, Mälardalens högskola, Akademin för innovation, design och teknik, 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:mdh:diva-15293.
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Wind power counts for roughly 3 % of the global electricity production. In the chase to produce greener power, much attention lies on getting more electricity from the wind, extraction of kinetic energy, with help of wind turbines. Wind turbines have been used for electricity production since 1887 and have since then developed into more efficient designs and become significantly bigger and with a higher efficiency. The operational conditions change considerably over the rotor length. Inner sections are typically exposed to more complex operational conditions than the outer sections. However, the outer blade sections have a much larger impact on the power and load generation. Especially here the demand for good aerodynamic performance is large. Airfoils have to be identified and investigated on mid/outer sections of a 7.0 MW rotor with 165 m diameter. Blade performance criteria were determined and investigations like sensitivity analysis were made. With the use of XFLR5 (XFoil) and Qblade, the airfoils were made into a blade and tested with the blade element momentum theory. This simulation gave detailed information regarding performance and operational loads depending on the different airfoils used. These results were then validated in a professional aero-elastic code (Flex5), simulating steady state, turbulent and wind shear conditions. The best airfoils to use from this reports airfoil catalogue are the NACA 63-6XX and NACA 64-6XX. With the implementation of these airfoils, blade design 2 and 3 have a very high performance coefficient compared to large commercial HAWT rotors.
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Abdo, Mohammed. "Analysis of steady and unsteady flows past fixed and oscillating airfoils." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 2000. http://www.collectionscanada.ca/obj/s4/f2/dsk1/tape4/PQDD_0032/MQ64209.pdf.
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Abdo, Mohammed. "Theoretical and computational analysis of airfoils in steady and unsteady flows." Thesis, McGill University, 2004. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=84871.
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This dissertation studies three aspects of airfoil flows: (i) second-order theoretical solutions of airfoils in steady flows; (ii) unsteady solutions for oscillating flexible airfoils; and (iii) numerical analysis of airfoil flows at low Reynolds numbers.The first part presents simple and efficient analytical solutions in closed form for the velocity and pressure distributions on airfoils of arbitrary shapes in steady flows, which are obtained using special singularities in the expression of the fluid velocity. A second-order accurate method is first developed for airfoils in inviscid incompressible flows to simultaneously solve the symmetric and anti-symmetric flow components defined by coupled boundary conditions. Then, the method is extended to take into account the viscous and compressibility effects on the pressure distribution. The resulting solutions were found to be in very good agreement with the available exact solutions (for specific airfoils), and with numerical and experimental results at various Mach and Reynolds numbers and moderate angles of attack.
The second part presents a new method of solution for the analysis of unsteady incompressible flows past oscillating rigid and flexible airfoils. The method has been successfully validated by comparison with the results obtained by Theodorsen and by Postel and Leppert for rigid airfoil and aileron oscillations in translation and rotation. The aerodynamic stiffness, damping and virtual mass contributions are specifically determined, as required in the aeroelastic studies. In all cases studied, this method led to very efficient and simple analytical solutions in closed form.
The third part presents an efficient numerical method for the incompressible flows past airfoils at low Reynolds numbers, which are of interest for micro-aircraft applications. The present analysis is based on a pseudo-time integration method using artificial compressibility to accurately solve the Navier-Stokes equations. Solutions are obtained with this method for airfoils at various incidences and very low Reynolds numbers between 400 and 6000. A detailed analysis is presented for the influence of the Reynolds number, incidence and airfoil shape on the pressure distribution, lift and drag coefficients. The flow separation is especially studied; the separation and reattachment positions are compared for various airfoil shapes, incidences and Reynolds numbers.
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Roy, Valentin. "Numerical analysis of airfoils with gurney flaps at low reynolds number." Thesis, McGill University, 2013. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=119634.
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A two-dimensional numerical investigation was performed to determine the effect of a Gurney flap on symmetric and cambered airfoils at very low Reynolds number. A Gurney flap is a device consisting of a small tab attached at the trailing edge and perpendicular to the chord line of the airfoil. Originally installed on a race car wing, a Gurney flap was proven to have a positive influence on the lift to drag ratio, and leading therefore to better efficiency and performance. First studies of Gurney flap showed that the optimal length of such a device is in the order of 1-4% of the chord length. The increase in lift comes primarily from the increased effective camber of the airfoil without disturbing the upper surface flow. The flow field around the airfoil is numerically studied in this thesis using an efficient numerical analysis based on a pseudo-time integration method using artificial compressibility. Gurney flap of sizes ranging from 1 to 4% of the airfoil chord were studied. The numerical solutions show the pressure distribution along the airfoil and flap as well as the lift and drag coefficients. The streamline contours illustrating the flow separation regions have also been generated. The influence of the flap length and the angle of attack on the aerodynamic coefficients was thoroughly studied.Une etude numérique bi-dimensionelle a été menée afin de determiner l'influence de la présence d'un aileron de type Gurney sur des profils d'aile symétriques et cambrés pour de très faibles nombres de Reynolds. Un rabat Gurney consiste en une petite pièce droite fixée au bord de fuite de l'aile et perpendiculaire à la corde de celle-ci. A l'origine installé sur l'aileron d'une voiture de grand tourisme, il a été observe que le rabat Gurney pouvait augmenter le coefficient de portance sans toutefois déteriorer le rapport portance/trainée, resultant ainsi en de meilleures performances. Les premières etudes faisant intervenir le rabat Gurney ont montrées que la longueur optimale de ce dernier est de l'ordre de 1 a 4% de la corde. L'augmentation de la portance vient premièrement du fait de l'augmentation effective de la cambrure de l'aile, sans toutefois venir perturber l'écoulement extrado. Le champ aérodynamique autour de l'aile a été calculé numériquement par le biais d'un schema numérique base sur une méthode d'intégration en pseudo-temps incluant le concept de compressibilité artificielle. Des volets dont la taille varie de 1à 4% de la corde de l'aile ont été etudiés. Des visualisations incluant les lignes de courant ainsi que les zones de recirculation ont également été génerées Cette etude détaille l'influence des parameters nombre de Reynolds, configuration de l'aile, angle d'attaque, taille du rabat sur l'écoulement.
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Reiss, Hans Claudius. "Experimental study on film cooling of gas turbine airfoils using shaped holes /." Lausanne : EPFL, 2000. http://library.epfl.ch/theses/?nr=2209.
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Gunaydinoglu, Erkan. "Low Reynolds Number Aerodynamics Of Flapping Airfoils In Hover And Forward Flight." Master's thesis, METU, 2010. http://etd.lib.metu.edu.tr/upload/12612397/index.pdf.
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The scope of the thesis is to numerically investigate the aerodynamics of flapping airfoils in hover and forward flight. The flowfields around flapping airfoils are computed by solving the governing equations on moving and/or deforming grids. The effects of Reynolds number, reduced frequency and airfoil geometry on unsteady aerodynamics of flapping airfoils undergoing pure plunge and combined pitch-plunge motions in forward flight are investigated. It is observed that dynamic stall of the airfoil is the main mechanism of lift augmentation for both motions at all Reynolds numbers ranging from 10000 to 60000. However, the strength and duration of the leading edge vortex vary with airfoil geometry and reduced frequency. It is also observed that more favorable force characteristics are achieved at higher reduced frequencies and low plunging amplitudes while keeping the Strouhal number constant. The computed flowfields are compared with the wide range of experimental studies and high fidelity simulations thus it is concluded that the present approach is applicable for investigating the flapping wing aerodynamics in forward flight. The effects of vertical translation amplitude and Reynolds number on flapping airfoils in hover are also studied. As the vertical translation amplitude increases, the vortices become stronger and the formation of leading edge vortex is pushed towards the midstroke of the motion. The instantaneous aerodynamic forces for a given figure-of-eight motion do not alter significantly for Reynolds numbers ranging from 500 to 5500.
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Liggett, Nicholas Dwayne. "Numerical investigation of static and dynamic stall of single and flapped airfoils." Diss., Georgia Institute of Technology, 2012. http://hdl.handle.net/1853/45834.
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Separated flows about single and multi-element airfoils are featured in many scenarios of practical interest, including: stall of fixed wing aircraft, dynamic stall of rotorcraft blades, and stall of compressor and turbine elements within jet engines. In each case, static and/or dynamic stall can lead to losses in performance. More importantly, modeling and analysis tools for stalled flows are relatively poorly evolved and designs must completely avoid stall due to a lack of understanding. The underlying argument is that advancements are necessary to facilitate understanding of and applications involving static and dynamic stall.
The state-of-the-art in modeling stall involves numerical solutions to the governing equations of fluids. These tools often either lack fidelity or are prohibitively expensive. Ever-increasing computational power will likely lead to increased application of numerical solutions. The focus of this thesis is improvements in numerical modeling of stall, the need of which arises from poorly evolved analysis tools and the spread of numerical approaches. Technical barriers have included ensuring unsteady flow field and vorticity reproduction, transition modeling, non-linear effects such as viscosity, and convergence of predictions.
Contributions to static and dynamic stall analysis have been been made. A hybrid Reynolds-Averaged Navier-Stokes/Large-Eddy-Simulation turbulence technique was demonstrated to predict the unsteadiness and acoustics within a cavity with accuracy approaching Large-Eddy-Simulation. Practices to model separated flows were developed and applied to stalled airfoils. Convergence was characterized to allow computational resources to be focused only as needed. Techniques were established for estimation of integrated coefficients, onset of stall, and reattachment from unconverged data. Separation and stall onset were governed by turbulent transport, while the location of reattachment depended on the mean flow. Application of these methodologies to oscillating flapped airfoils revealed flow through the gap was dominated by the flap angle for low angles of attack. Lag between the aerodynamic response and input flap scheduling was associated with increased oscillation frequency and airfoil/flap gap size. Massively separated flow structures were also examined.
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Wo, Andrew Man-Chung. "Characteristics of airfoils in an oscillating external flow at low Reynolds numbers." Thesis, Massachusetts Institute of Technology, 1990. http://hdl.handle.net/1721.1/42212.
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Soueid, Houssam. "Optimisation of the kinematics of flapping airfoils for micro air vehicle applications." Toulouse 3, 2008. http://thesesups.ups-tlse.fr/791/.
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L'optimisation du battement d'un profil d'aile est réalisée par simulations numériques. L'objectif principal est d'identifier les paramètres de battement capables d'assurer à un micro-drone de bonnes performances aérodynamiques en termes de forces de poussée, de portance et de rendement propulsif. Les variables de contrôle sont les amplitudes des oscillations verticales et de tangage, leur déphasage, leur fréquence, et l'angle d'incidence moyen. Une fois déterminées, les cinématiques optimales sont confrontées aux valeurs observées dans la nature. La méthodologie est basée sur la résolution de l'écoulement autour d'un profil NACA0012 battant. Divers fonctionnels coûts sont minimisés en fonction de la mission du véhicule. Leurs gradients par rapport aux paramètres cinématiques sont calcules par la technique des sensibilités et la méthode du pas complexe. Un algorithme de quasi-Newton permet d'actualiser avec les gradients le mouvement du profil vers un optimum de la fonctionnelle. Les résultats montrent la capacité d'un profil optimisé d'aile battante à produire de grandes forces de poussée et de portance avec un rendement propulsif acceptable. Pour ce faire, il faut que les oscillations de tangage précèdent celles de translation verticale par un déphasage de l'ordre de 90º et que le nombre de Strouhal soit dans l'intervalle [0. 2, 0. 4]. Augmenter le nombre de Reynolds a pour effet de réduire les pertes d'énergie liées à la viscosité et d'augmenter le rendement propulsif. Ce rendement a aussi été accru pour les cas de grandes forces de poussée en incluant des harmoniques d'ordre supérieur dans les oscillations verticalesThe optimisation of the kinematics of a two-dimensional flapping airfoil is carried out by numerical simulation to identify flapping parameters capable to ensure high aerodynamic performances for a micro air vehicle application. Attention is focused on the lift and thrust and on the propulsive efficiency. The control variables are the amplitudes of the combined motions of heaving and pitching and the phase angle between them. Additionally, the flapping frequency and the mean angle of attack are considered. The optimal kinematics are compared to observations in nature on birds, insects and fish. The methodology is based on the solution of the flow field around a NACA0012 airfoil submitted to translational and angular oscillations. Then, the gradients of the cost functional with respect to the control parameters are evaluated thanks to the sensitivity technique and the complex step derivative method. The gradients are subsequently used in a quasi-Newton update algorithm to direct the solution towards its optimal value. Results show the ability of optimised flapping airfoils to produce large thrust and lift, with an acceptable efficiency. This is ensured when the pitching oscillations lead the heaving ones by a phase angle close to 90º and when the Strouhal number of oscillations is in the range [0. 2, 0. 4]. It is expected that the optimal efficiency would increase for higher Reynolds number, without much variation in the optimal kinematics. Furthermore, the drop in efficiency at high thrust forces may be limited by including higher order terms in the expression of heaving
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Mish, Patrick Francis. "An Experimental Investigation of Unsteady Surface Pressure on Single and Multiple Airfoils." Diss., Virginia Tech, 2003. http://hdl.handle.net/10919/26567.
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This dissertation presents measurements of unsteady surface pressure on airfoils encountering flow disturbances. Analysis of measurements made on an airfoil immersed in turbulence and comparisons with inviscid theory are presented with the goal of determining the effect of angle of attack on an airfoils inviscid response. Unsteady measurements made on the surface of a linear cascade immersed in periodic flow are presented and analyzed to determine the relationship between the blades inviscid response and tip leakage vortex strength.
Measurements of fluctuating surface pressure were made on a NACA 0015 airfoil immersed in grid generated turbulence. The airfoil model has a 2' chord and spans the 6' Virginia Tech Stability Wind Tunnel test section. Two grids were used to investigate the effects of turbulence length scale on the surface pressure response. A large grid which produced turbulence with an integral scale 13% of the chord and a smaller grid which produced turbulence with an integral scale 1.3% of the chord. Measurements were performed at angles of attack from 0º to 20º. An array of microphones mounted subsurface was used to measure the unsteady surface pressure. The goal of this measurement was to characterize the effects of angle of attack on the inviscid response.
Lift spectra calculated from pressure measurements at each angle of attack revealed two distinct interaction regions; for reduced frequencies < 10 a reduction in unsteady lift of up to 7 decibels (dB) occurs while an increase occurs for reduced frequencies > 10 as the angle of attack is increased. The reduction in unsteady lift at low reduced frequencies with increasing angle of attack is a result that has never before been shown either experimentally or theoretically. The source of the reduction in lift spectral level appears to be closely related to the distortion of inflow turbulence based on analysis of surface pressure spanwise correlation length scales. Furthermore, while the distortion of the inflow appears to be critical in this experiment, this effect does not seem to be significant in larger integral scale (relative to the chord) flows based on the previous experimental work of McKeough (1976) suggesting the airfoils size relative to the inflow integral scale is critical in defining how the airfoil will respond under variation of angle of attack.
A prediction scheme is developed that correctly accounts for the effects of distortion when the inflow integral scale is small relative to the airfoil chord. This scheme utilizes Rapid Distortion Theory to account for the distortion of the inflow with the distortion field modeled using a circular cylinder.
Measurement of the unsteady surface pressure response of a linear cascade in periodic disturbance is presented. Unsteady pressure was measured on the suction and pressure side of two cascade blades with an array of 24 microphones (12 per blade side) mounted subsurface. The periodic disturbance was generated using a pair of vortex generators attached to a moving end wall. Measurements were made for 8 tip gaps (t/c = 0.00825, 0.0165, 0.022, 0.033, 0.045, 0.057, 0.079, 0.129) and phased averaged with respect to the vortex generator pair position. This measurement was motivated by the results presented by Ma (2003). The work of Ma (2003) suggested that tip leakage vortex shedding in the presence of a periodic disturbance is heavily influenced by the inviscid response of the cascade blade. This conclusion was arrived at by Maâ s (2003) observation that as the tip gap is increased the amount of fluctuation in the tip leakage vortex circulation increases dramatically, in fact, many times the circulation in the inflow vortices.
Unsteady pressure measurements reveal that the blade response involves a complex interaction of both inviscid response and viscous phenomena. However, a close relationship between unsteady tip loading and tip leakage vortex circulation is revealed suggesting the inviscid response is significant in determining the tip leakage vortex circulation. Additionally, predictions using inviscid theory agree well with measured levels of unsteady tip loading. As such, inviscid theory may be useful for predicting the tip leakage circulation and perhaps, pressure fluctuations in the tip leakage vortex.Ph. D.
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Kaya, Mustafa. "Computation Of Viscous Flows Over Flapping Airfoils And Parallel Optimization Of Flapping Parameters." Master's thesis, METU, 2003. http://etd.lib.metu.edu.tr/upload/12612286/index.pdf.
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Airfoils °apping in pitch and plunge are studied, and the °
apping motion parameters are op- timized to maximize thrust generation and the e±
ciency of the thrust generation. Unsteady viscous °
ow¯
elds over °
apping airfoils are computed on overset grids using a Navier-Stokes solver. Computations are performed in parallel using Parallel Virtual Machine library routines in a computer cluster. A single °
apping airfoil and dual airfoils °
apping in a biplane con- ¯
guration are considered. A gradient based optimization algorithm is employed. The thrust production and the e±
ciency of the thrust production are optimized with respect to °
apping parameters
the plunging and pitching amplitudes, the °
apping frequency, and the phase shift between the pitch and plunge motions. It is observed that thrust generation of °
apping airfoils strongly depends on the phase shift and high thrust values may be obtained at the expense of reduced e±
ciency. For a high e±
ciency in thrust generation, the e®
ective angle of attack of the airfoil is reduced and large scale vortex formations at the leading edge are prevented. At a ¯
xed reduced °
apping frequency of 1, a single °
apping airfoil in pitch and plunge motion produces the maximum average thrust coe±
cient of 1:41 at the plunge amplitude of 1:60, the pitch amplitude of 23:5o, and the phase shift of 103:4o whereas the maximum e±
ciency of 67:5% is obtained at the plunge amplitude of 0:83, the pitch amplitude of 35:5o and the phase shift of 86:5o.
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Tveiterås, Vebjørn. "Numerical Study of the Interaction of Flow over Two Airfoils in Relative Motion." Thesis, Norges teknisk-naturvitenskapelige universitet, Institutt for fysikk, 2011. http://urn.kb.se/resolve?urn=urn:nbn:no:ntnu:diva-13652.
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Computational fluid dynamics (CFD) was be used to investigate aspects of interaction of flow over two airfoils in relatie motion in detail.%Problem statement2D tandem airfoil setups were studied, where the leading airfoil was performing an oscillating motion in the vertical direction while the trailing airfoil was kept stationary. The NACA 0012 and the S809 airfoils were considered.%ApproachAnsys Fluent v13.0 was used as the CFD solver, and Gambit v2.4.6 was employed for grid generation. All simulations were transient at a Reynolds number of either 2*10^4 (laminar flow) or 3*10^6 (turbulent flow). The Transition SST turbulence model was chosen to model turbulence, and Fluent's sliding grid technique was used to achieve the relative motion between the airfoils.%ResultsThe tandem setup was found able to outperform a single airfoil for similar conditions. The presence of a trailing airfoil did not significantly affect the leading airfoil's performance, whereas it did affect the wake structures significanlty. The suction peak near the nose of the airfoil was found to be the most important factor determining the airfoil's propulsive efficiency. Therefore, leading edge vortex (LEV) shedding was found to be of higher importance than trailing edge vortex (TEV) shedding when airfoil performance was concidered.The asymmetric S809 airfoil provided similar results as the symmetric NACA 0012 airfoil. However, the NACA 0012 airfoil achieved slightly higher propulsive efficiencies for the cases investigated, indicating that a symmetric airfoil is desired for flapping airfoil setups. For the tandem setup the highest propulsive efficienies were 0.766 and 0.742 for the NACA 0012 airfoil and the S809 airfoil, respectively. Both peaks were found at k = 0.3 and h = 0.6 (Sr = 0.11) for the leading airfoil. A maximum thrust coefficient of 2.32 was found for the tandem S809 airfoil setup at k = 1.5 and h = 0.5 (Sr = 0.48).The Strouhal number was found to be an important describing parameter, but additional information about the reduced frequency or the oscillating amplitude was needed in order to fully describe the setup.
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Geyer, Thomas F. [Verfasser], and Ennes [Akademischer Betreuer] Sarradj. "Trailing edge noise generation of porous airfoils / Thomas F. Geyer. Betreuer: Ennes Sarradj." Cottbus : Universitätsbibliothek der BTU Cottbus, 2011. http://d-nb.info/1016457812/34.
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Smith, Justin L. "Computational analysis of airfoils in ground effect for use as a design tool." Morgantown, W. Va. : [West Virginia University Libraries], 2007. https://eidr.wvu.edu/etd/documentdata.eTD?documentid=5291.
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Thesis (M.S.)--West Virginia University, 2007.Title from document title page. Document formatted into pages; contains viii, 59 p. : ill. (some col.). Includes abstract. Includes bibliographical references (p. 51-52).
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