Journal articles on the topic 'Aerofoil'
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1
Liang, Changping, and Huaxing Li. "Aerofoil optimization for improving the power performance of a vertical axis wind turbine using multiple streamtube model and genetic algorithm." Royal Society Open Science 5, no. 7 (July 2018): 180540. http://dx.doi.org/10.1098/rsos.180540.
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This paper reports on the optimization of the NACA0015 aerofoil for improving the power performance of a vertical axis wind turbine (VAWT). The target range of the chord Re is 3 × 10 5 –10 6 , the tip speed ratio (TSR) is 2–6 and the solidity is 0.2–0.6. This aerofoil is widely applied in small-scale VAWTs. In the optimization process, in which the class and shape function transformation parametrization method was used to perturb the aerofoil geometry, the thickness and camber of the aerofoil were selected as the constraints and the value of the maximum tangential force coefficient was chosen as the objective function. The aerodynamic performance of the aerofoil was calculated by combining the XFOIL program and Viterna–Corrigan post-stall model, while the aerofoil's performance was validated with computational fluid dynamic simulations. The results illustrated that, compared to an unoptimized NACA0015 aerofoil, the optimized aerofoil's lift to drag ratio was improved over a wide range of attack angles and the stall performance was gentler. The maximum lift coefficient, the maximum lift to drag ratio and the maximum tangential force coefficient were increased by 7.5%, 9% and 8.87%, respectively. Finally, this paper predicted the rotor efficiency with both the unoptimized and optimized NACA0015 aerofoils for different TSRs and different solidities using the multiple streamtube model. The results showed that the rotor with the optimized aerofoil has a higher efficiency.
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Zhang, Hailang, Yu Hu, and Gengqi Wang. "The effect of aerofoil camber on cycloidal propellers." Aircraft Engineering and Aerospace Technology 90, no. 8 (November 5, 2018): 1156–67. http://dx.doi.org/10.1108/aeat-08-2016-0128.
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Purpose This paper aims to investigate the impact of aerofoil camber on the performance of micro-air-vehicle-scale cycloidal propellers. Design/methodology/approach First, experiments were conducted to validate the numerical methodology. After that, three turbulent models were compared to select the most accurate one. Then, 2D numerical simulation was carried out on 11 aerofoils with different cambers, including five cambered aerofoils, one symmetrical aerofoil and five inverse cambered aerofoils. The inverse cambered aerofoils are symmetrical about the chord line to the corresponding cambered ones. Findings The cycloidal propeller with large cambered aerofoil gives the lowest hovering efficiency, but with symmetrical aerofoil or small inverse cambered aerofoil shows the highest. Also, blades with large cambered aerofoil display high performance at the upper part of its trajectory, while with symmetrical aerofoil or the inverse cambered aerofoil have their best at the lower part. In addition, intensified downwash can be observed in the rotor cage for all cases. When a blade runs through the top-left part of its circle path, all cases display the feature of deep dynamic stall. When the blade travels through the nadir of its path, the actual angle of attack is close to zero due to the strong downwash. Furthermore, there exits intensified blade-vortex interaction induced by the preceding blade for large cambered aerofoils at the lower-right part of its trajectory. Practical implications This paper develops a new cycloidal propeller which is more efficient than the one already present. Originality/value This paper discovers that the aerofoil camber is a vital design parameter in the performance of cycloidal propeller, and the authors expect that the rotor with deformable aerofoil on camber would achieve much higher efficiency.
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Shen, Xiang, Theodosios Korakianitis, and Eldad Avital. "Numerical Investigation of Surface Curvature Effects on Aerofoil Aerodynamic Performance." Applied Mechanics and Materials 798 (October 2015): 589–95. http://dx.doi.org/10.4028/www.scientific.net/amm.798.589.
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The prescribed surface curvature distribution blade design (CIRCLE) method optimises aerofoils and blades by controlling curvature continuity and slope of curvature distribution along their surfaces. The symmetrical NACA0012 exhibits a surface curvature discontinuity at the leading edge point, and the non-symmetrical E387 exhibits slope-of-curvature discontinuities in the surface. The CIRCLE method is applied to both aerofoils to remove both surface curvature and slope-of-curvature discontinuities. Computational fluid dynamics analyses are used to investigate the curvature effects on aerodynamic performance of the original and modified aerofoils. These results are compared with experimental data obtained from tests on the original aerofoil geometry. The computed aerodynamic advantages of the modified aerofoil are analysed in different operating conditions. The leading edge singularity of NACA0012 is removed and it is shown that the surface curvature discontinuity affects the aerodynamic performance near the stalling angle of attack. The discontinuous slope-of-curvature distribution of E387 influences the size of the laminar separation bubble at lower Reynolds numbers, and it affects the inherent profile of the aerofoil at higher Reynolds numbers. It is concluded that the surface curvature distribution of aerofoils has a significant effect on aerofoil aerodynamic performance, which can be improved by redesigning the surface curvature distribution of the original aerofoil geometry.
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., Sundram, and Rahul Kumar. "Investigation of Airfoil Design and Analysis." International Journal for Research in Applied Science and Engineering Technology 10, no. 10 (October 31, 2022): 863–80. http://dx.doi.org/10.22214/ijraset.2022.47087.
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Abstract: In this project “Aerofoil Design and Analysis” an attempt has made to make a complete study on lift and drag coefficient of various aerofoil sections using CFD. The primary goal of this project is to learn and analyse the aerodynamic performance of wings. The objective of this study is to improve aerofoil design using the software CATIA, And Fluent Analysis using the software ANSYS. Aerofoil is a principal part of any airplane construction. How much lift force and drag force is sufficient to balance the weight of the plane is decided by the aerofoil. Aerofoils are basically divided into two categories - they are Asymmetrical and Symmetrical aerofoils. Based on their drag and lift coefficient’s variation with angle of attack, stall angle of attack and magnitude of the coefficients they are divided. Here the NACA aerofoil is modified by adding dimples on the upper half of the wing and compared with the simple one. The comparison is made on different speed and pressure on the wing and the coefficient of lift and drag.
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., Albi, M. Dev Anand, and G. M. Joselin Herbert. "Aerodynamic Analysis on Wind Turbine Aerofoil." International Journal of Engineering & Technology 7, no. 3.27 (August 15, 2018): 456. http://dx.doi.org/10.14419/ijet.v7i3.27.17997.
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The aerofoils of wind turbine blades have crucial influence on aerodynamic efficiency of wind turbine. There are numerous amounts of research being performed on aerofoils of wind turbines. Initially, I have done a brief literature survey on wind turbine aerofoil. This project involves the selection of a suitable aerofoil section for the proposed wind turbine blade. A comprehensive study of the aerofoil behaviour is implemented using 2D modelling. NACA 4412 aerofoil profile is considered for analysis of wind turbine blade. Geometry of this aerofoil is created using GAMBIT and CFD analysis is carried out using ANSYS FLUENT. Lift and Drag forces along with the angle of attack are the important parameters in a wind turbine system. These parameters decide the efficiency of the wind turbine. The lift force and drag force acting on aerofoil were determined with various angles of attacks ranging from 0° to 12° and wind speeds. The coefficient of lift and drag values are calculated for 1×105 Reynolds number. The pressure distributions as well as coefficient of lift to coefficient of drag ratio of this aerofoil were visualized. The CFD simulation results show close agreement with those of the experiments, thus suggesting a reliable alternative to experimental method in determining drag and lift.
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Zachariou, E., P. Wilmott, and A. D. Fitt. "A cavitating aerofoil with a Prandtl-Batchelor eddy." Aeronautical Journal 98, no. 975 (May 1994): 171–76. http://dx.doi.org/10.1017/s000192400004985x.
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Abstract A simple model is presented for an aerofoil with a recirculating Prandtl-Batchelor region behind a spoiler. Using thin aerofoil theory the model is posed as a pair of coupled nonlinear singular integrodifferential equations for the shape of the separating streamline and the distribution of vorticity along the aerofoil. These equations are solved numerically and results are presented. In particular, some conclusions are drawn regarding the lift on such aerofoils.
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Saeedi, M., F. Ajalli, and M. Mani. "A comprehensive numerical study of battle damage and repairs upon the aerodynamic characteristics of an aerofoil." Aeronautical Journal 114, no. 1158 (August 2010): 469–84. http://dx.doi.org/10.1017/s0001924000003961.
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Abstract A NACA 641-412 aerofoil with circle and star damage and also three repair configurations has been numerically investigated. Two different methods of mesh generation were employed: multi structured mesh for the star damaged aerofoil and unstructured mesh for the other aerofoils. The results show that the damage will cause a reduction in lift coefficient of the aerofoil and also a different stall angle relative to that of the undamaged aerofoil. Each kind of repair improves the aerodynamic characteristics of the aerofoil considerably. The flow Field inside the damage hole and the cavity caused by the repair sheets was also investigated. Finally, the numerical solution was qualitatively and quantitatively validated using the available experimental results.
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Campanile, L. F., and G. Thwapiah. "A non-linear theory of torsional divergence." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 223, no. 11 (September 11, 2009): 2707–11. http://dx.doi.org/10.1243/09544062jmes1843.
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In recent years, research on aerofoil morphing is increasingly focusing on innovative ideas such as the use of compliant systems and the exploitation of aeroelastic servo-effects. If brought to their limit, these concepts would allow operating aerofoils in aeroelastically marginally stable or even unstable conditions. In this view, a non-linear approach to aeroelastic torsional divergence becomes relevant. This article presents an extension of the well-known linear theory of divergence, which takes into account non-linear effects of structural as well as aerodynamic nature. The non-linear theory is applied to the case of a thin aerofoil and the pre-critical as well as post-critical response is computed for selected values of the flow parameters. Instability curves are also included, which show the aerofoil's torsional deformation as a function of the dynamic pressure, for selected values of an imposed disturbance.
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Haselbach, Frank, Heinz-Peter Schiffer, Manfred Horsman, Stefan Dressen, Neil Harvey, and Simon Read. "The Application of Ultra High Lift Blading in the BR715 LP Turbine." Journal of Turbomachinery 124, no. 1 (February 1, 2001): 45–51. http://dx.doi.org/10.1115/1.1415737.
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The original LP turbine of the BR715 engine featured “High Lift” blading, which achieved a 20-percent reduction in aerofoil numbers compared to blading with conventional levels of lift, reported in Cobley et al. (1997). This paper describes the design and test of a re-bladed LP turbine with new “Ultra High Lift” aerofoils, achieving a further reduction of approximately 11 percent in aerofoil count and significant reductions in turbine weight. The design is based on the successful cascade experiments of Howell et al. (2000) and Brunner et al. (2000). Unsteady wake-boundary layer interaction on these low-Reynolds-number aerofoils is of particular importance in their successful application. Test results show the LP turbine performance to be in line with expectation. Measured aerofoil pressure distributions are presented and compared with the design intent. Changes in the turbine characteristics relative to the original design are interpreted by making reference to the detailed differences in the two aerofoil design styles.
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Wibowo, Heri. "Pengaruh Sudut Serang Aerofoil Terhadap Distribusi Tekanan dan Gaya Angkat." JURNAL DINAMIKA VOKASIONAL TEKNIK MESIN 2, no. 2 (October 1, 2017): 148. http://dx.doi.org/10.21831/dinamika.v2i2.15999.
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The aerofoils are used to get the lifting force on the design of plane’s wings. The lifting force is caused by difference air velocity on upper and lower aerofoil which the magnitude depend on attack angle of aerofoil and air velocity exist surrounding. This experiment aims to show the force vector (pressure distribute) on the aerofoil. The aerofoil is attached by air with konstant velocity. The research procedure is done by change the attack angle of aerofoil on five formation. The surface of aerofoil is connected with pressure gage which disperse at 11 point of measurement. The result shows that magnitude of force vector is depended on attack angle of aerofoil. Increasing angle of aerofoil until boundary angle will be followed by increasing air velocity on the point of measurement and finally increase force vector. Upper boundary angle will be followed by decreasing air velocity and finally decrease force vector.Aerofoil digunakan untuk mendapatkan gaya angkat pada desain sayap pesawat. Gaya angkat disebabkan oleh perbedaan kecepatan udara pada aerofoil atas dan bawah yang besarnya tergantung pada sudut serang aerofoil dan kecepatan udara yang ada disekitarnya. Eksperimen ini bertujuan untuk menampilkan vektor gaya (distribusi tekanan) pada aerofoil. Aerofoil dilekatkan pada udara dengan kecepatan konstan. Prosedur penelitian dilakukan dengan cara mengubah sudut serang aerofoil pada lima formasi. Permukaan aerofoil dihubungkan dengan alat ukur tekanan yang tersebar pada 11 titik pengukuran. Hasil menunjukkan bahwa besarnya vektor gaya bergantung pada sudut serang aerofoil. Sudut aerofoil yang meningkat hingga sudut batas akan diikuti dengan peningkatan kecepatan udara pada titik pengukuran dan akhirnya meningkatkan vektor gaya. Sudut batas atas akan diikuti oleh penurunan kecepatan udara pada titik pengukuran dan akhirnya menurunkan vektor gaya.
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Baddoo, P. J., and L. J. Ayton. "Potential flow through a cascade of aerofoils: direct and inverse problems." Proceedings of the Royal Society A: Mathematical, Physical and Engineering Sciences 474, no. 2217 (September 2018): 20180065. http://dx.doi.org/10.1098/rspa.2018.0065.
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The potential flow through an infinite cascade of aerofoils is considered as both a direct and inverse problem. In each case, a perturbation expansion about a background uniform flow is assumed where the size of the perturbation is comparable to the aspect ratio of the aerofoils. This perturbation must decay far upstream and also satisfy particular edge conditions, including the Kutta condition at each trailing edge. In the direct problem, the flow field through a cascade of aerofoils of known geometry is calculated. This is solved analytically by recasting the situation as a Riemann–Hilbert problem with only imaginary values prescribed on the chords. As the distance between aerofoils is taken to infinity, the solution is seen to converge to a known analytic expression for a single aerofoil. Analytic expressions for the surface velocity, lift and deflection angle are presented as functions of aerofoil geometry, angle of attack and stagger angle; these show good agreement with numerical results. In the inverse problem, the aerofoil geometry is calculated from a prescribed tangential surface velocity along the chords and upstream angle of attack. This is found via the solution of a singular integral equation prescribed on the chords of the aerofoils.
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Sri, Ram Deepak Akella, Srinivas Baswanth Pappula Sashendra, Charan Jalli Satish, and B. S. V. Ramarao. "Aerodynamics analysis of RAF family aerofoil using computational fluid dynamics." i-manager's Journal on Mechanical Engineering 12, no. 3 (2022): 47. http://dx.doi.org/10.26634/jme.12.3.18673.
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Aerofoil is a design that offers a gold standard ratio between the coefficient of drag to the coefficient of lift. Aerofoil are the standard cross-sectional structure of the plane wing. In this paper, a famous aerofoil family which is acknowledged to be the Royal Aircraft Factory (RAF) is regarded with all the accessible aerofoil models in that family. The pinnacle 5 aerofoils are considered and inspected for all drift editions over the viewed models with various flow transforming from 400, 420, 440, 460, 480 and 500 knots. The current research is to determine the high-quality aerofoil sketch over their family, for the quality aerofoil format by using plan amendment, and the float variation is carried out by varying the angle of assault with 5 stages interval from 0-25 ranges and the pressure, velocity, turbulence length, temperature, Mach number and the pressure performing in x & y-direction are recorded. All these parameters in consideration determined the most fulfilling model for the RAF family.
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Hassan, Mohd Roshdi, Yong Thian Haw, and Mohd Nasrisyam Asri. "Design Analysis of Two Ways Shape Memory Alloy (SMA) Actuated Aerofoil." Applied Mechanics and Materials 564 (June 2014): 340–45. http://dx.doi.org/10.4028/www.scientific.net/amm.564.340.
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This paper describes the design analysis of the behavior of a shape memory alloy (SMA) plate embedded into an aerofoil. Experimentation and simulation were done to fulfill this purpose. The aerofoil is made of silicone rubber material. The SMA plate which was embedded into the maximum chamber of aerofoil during the fabrication process was measured at approximately 175mm, 63mm and 3mm in length, width and thickness respectively. Experimentation was conducted to show that the SMA plate is able to produce two-way shape memory effect. Simulation was executed by using Abaqus 6.9-1 (finite element analysis software). The aerofoil profile was changed by the movement of SMA plate, which has subsequently changed the angle of aerofoil’s trailing edge. The result from the experiment shows that the aerofoil’s trailing edge has undergone a certain amount of displacement after heated. Upon cooling, the aerofoil’s trailing edge did not return to its initial position. Based on this analysis, it is clear that the simulation results are in agreement with the findings of experimental results.
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Abu-Abdou, K., and M. F. Zedan. "Performance of improved thin aerofoil theory for modern aerofoil sections." Aeronautical Journal 95, no. 942 (February 1991): 64–70. http://dx.doi.org/10.1017/s0001924000023526.
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AbstractThe improved thin aerofoil method, which features extended expressions for lift and moment coefficients, is considered for further investigation and validation. The procedure to calculate the singularity coefficients is improved by using all aerofoil coordinates as control points in a least squares scheme. The classical NACA 0012 and NACA 65012 sections, the modern aviation aerofoil LS(1)—0417 and the extremely thick Kennedy-Marsden aerofoil are validated in place of the previously cited Karman-Trefftz aerofoil. This selection covers thickness ratios of up to 27·9%, camber ratios up to 7·69% and incidence up to 16·7°. Comparisons of velocity (or pressure) distributions and aerodynamic coefficients are made with two panel methods and with exact solution or experimental results whichever is available. Results indicated that the accuracy of the extended method is much better than expected and compares well with panel methods except for the extremely thick aerofoil. Additional results in the form of a systematic investigation of a weighted global error in the pressure distribution for the Karman-Trefftz aerofoils used in the previous study, are also included. Such an error shows similar trends and in many cases comparable magnitude to the errors generated by panel methods.
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Laratro, Alex, Maziar Arjomandi, Benjamin Cazzolato, and Richard Kelso. "Self-noise of NACA 0012 and NACA 0021 aerofoils at the onset of stall." International Journal of Aeroacoustics 16, no. 3 (April 2017): 181–95. http://dx.doi.org/10.1177/1475472x17709929.
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The aerodynamic noise of a NACA 0012 and NACA 0021 aerofoil is measured and compared in order to determine whether there are differences in their noise signatures with a focus on the onset of stall. Measurements of the self-noise of each aerofoil are measured in an open-jet Anechoic Wind Tunnel at Reynolds numbers of 64,000 and 96,000, at geometric angles of attack from −5° through 40° at a resolution of 1°. Further measurements are taken at Re = 96,000 at geometric angles of attack from −5 through 16° at a resolution of 0.5°. Results show that while the noise generated far into the stall regime is quite similar for both aerofoils the change in noise level at the onset of stall is significantly different between the two aerofoils with the NACA 0021 exhibiting a much sharper increase in noise levels below a chord-based Strouhal number of Stc = 1.1. This behaviour is consistent with the changes in lift of these aerofoils as well as the rate of collapse of the suction peak of a NACA 0012 aerofoil under these flow conditions.
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Radha Krishnan, Naren Shankar, and Shiva Prasad Uppu. "A novel approach for noise prediction using Neural network trained with an efficient optimization technique." International Journal for Simulation and Multidisciplinary Design Optimization 14 (2023): 3. http://dx.doi.org/10.1051/smdo/2023002.
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Aerofoil noise as self-noise is detrimental to system performance, in this paper NACA 0012 optimization parameters are presented for reduction in noise. Designing an aerofoil with little noise is a fundamental objective of designing an aircraft that physically and functionally meets the requirements. Aerofoil self-noise is the noise created by aerofoils interacting with their boundary layers. Using neural networks, the suggested method predicts aerofoil self-noise. For parameter optimization, the quasi-Newtonian method is utilised. The input variables, such as angle of attack and chord length, are used as training parameters for neural networks. The output of a neural network is the sound pressure level, and the Quasi Newton method further optimises these parameters. When compared to the results of regression analysis, the values produced after training a neural network are enhanced.
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Traub, L. W. "Estimating aerofoil lift from flow angle." Aeronautical Journal 119, no. 1219 (September 2015): 1167–73. http://dx.doi.org/10.1017/s0001924000011180.
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Estimation of the lift of an aerofoil is one of the fundamental measurements of fluid mechanics. Lift is commonly measured using a load cell or a force balance. Non-intrusive methods to measure lift are usually pressure based. Aerofoils may be pressure tapped where small surface orifices are connected via tubing to a pressure measurement system, either a multi-tube manometre or an electronic system. Both measurement options add cost and complication, especially in an educational setting. Pressure tapping small aerofoils can also be difficult, especially if the models are rapid prototyped (RP). Low model surface resolution (from RP manufacture) and confined geometry complicate model assembly and finishing. Boundary-layer transition caused by poorly implemented tappings (too large a diametre or poorly aligned, i.e. straight aft) can also alter results. Wall pressure tappings may also be used and have the benefit of being non-intrusive. To implement, the test section roof and floor is tapped with a streamwise row of ports that facilitate measurement of the wall pressure signature. Integration of the pressure differential then relates to the lift produced. This measurement methodology still requires a multi-channel pressure acquisition system and modification of the wind tunnel. In Refs 4,5 methods are presented that facilitate calculation of the instantaneous forces acting on a body through flow field measurements determined using particle image velocimetry. However, the required flow field measurements encompass those surrounding the body, and are not a simple point measurement. In Ref. 6 a method is presented to estimate the lift of an aerofoil using two Pitot-static tubes that are used to measure the velocity above and below the aerofoil’s quarter chord. Wall corrections are required to yield an accurate lift estimate.
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Koreanschi, A., O. Sugar-Gabor, and R. M. Botez. "Drag optimisation of a wing equipped with a morphing upper surface." Aeronautical Journal 120, no. 1225 (March 2016): 473–93. http://dx.doi.org/10.1017/aer.2016.6.
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ABSTRACTThe drag coefficient and the laminar-to-turbulent transition for the aerofoil component of a wing model are optimised using an adaptive upper surface with two actuation points. The effects of the new shaped aerofoils on the global drag coefficient of the wing model are also studied. The aerofoil was optimised with an ‘in-house’ genetic algorithm program coupled with a cubic spline aerofoil shape reconstruction and XFoil 6.96 open-source aerodynamic solver. The wing model analysis was performed with the open-source solver XFLR5 and the 3D Panel Method was used for the aerodynamic calculation. The results of the aerofoil optimisation indicate improvements of both the drag coefficient and transition delay of 2% to 4%. These improvements in the aerofoil characteristics affect the global drag of the wing model, reducing it by up to 2%. The analyses were conducted for a single Reynolds number and speed over a range of angles of attack. The same cases will also be used in the experimental testing of the manufactured morphing wing model.
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Chaitanya, P., P. Joseph, S. Narayanan, C. Vanderwel, J. Turner, J. W. Kim, and B. Ganapathisubramani. "Performance and mechanism of sinusoidal leading edge serrations for the reduction of turbulence–aerofoil interaction noise." Journal of Fluid Mechanics 818 (April 4, 2017): 435–64. http://dx.doi.org/10.1017/jfm.2017.141.
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This paper presents the results of a detailed experimental investigation into the effectiveness of sinusoidal leading edge serrations on aerofoils for the reduction of the noise generated by the interaction with turbulent flow. A detailed parametric study is performed to investigate the sensitivity of the noise reductions to the serration amplitude and wavelength. The study is primarily performed on flat plates in an idealized turbulent flow, which we demonstrate captures the same behaviour as when identical serrations are introduced onto three-dimensional aerofoils. The influence on the noise reduction of the turbulence integral length scale is also studied. An optimum serration wavelength is identified whereby maximum noise reductions are obtained, corresponding to when the transverse integral length scale is approximately one-fourth the serration wavelength. This paper proves that, at the optimum serration wavelength, adjacent valley sources are excited incoherently. One of the most important findings of this paper is that, at the optimum serration wavelength, the sound power radiation from the serrated aerofoil varies inversely proportional to the Strouhal number $St_{h}=fh/U$, where $f$, $h$ and $U$ are frequency, serration amplitude and flow speed, respectively. A simple model is proposed to explain this behaviour. Noise reductions are observed to generally increase with increasing frequency until the frequency at which aerofoil self-noise dominates the interaction noise. Leading edge serrations are also shown to reduce aerofoil self-noise. The mechanism for this phenomenon is explored through particle image velocimetry measurements. Finally, the lift and drag of the serrated aerofoil are obtained through direct measurement and compared against the straight edge baseline aerofoil. It is shown that aerodynamic performance is not substantially degraded by the introduction of the leading edge serrations on the aerofoil.
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Suddhoo, A., and I. M. Hall. "Test cases for the plane potential flow past multi-element aerofoils." Aeronautical Journal 89, no. 890 (December 1985): 403–14. http://dx.doi.org/10.1017/s0001924000096767.
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SummaryThe method of images is used to calculate the analytical solution for an inviscid incompressible flow past four arbitrary circles which are then conformally mapped onto aerofoil sections. The analytical solution of the corresponding flow past the system of aerofoils is obtained and the pressure distributions about a particular three-element and a particular fourelement aerofoil are presented. These results are intended to be used to assess the accuracy of numerical methods.
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He, W., and X. Liu. "Improved aerofoil parameterisation based on class/shape function transformation." Aeronautical Journal 123, no. 1261 (March 2019): 310–39. http://dx.doi.org/10.1017/aer.2018.165.
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ABSTRACTA new aerofoil parameterisation method is put forward to represent an aerofoil by combining the leading edge modification class/shape function transformation (LEM CST) method and improved Hicks–Henne bump function’s method. The new class/shape function transformation (NEW CST) method has two additional basis functions comparing the original CST method. In order to confirm these two basis functions, the radial basis functions neural network (RBF) model is trained by some samples which are generated by the Latin hypercube design (LHD) method and Genetic Algorithm (GA) is proposed to achieve the basis functions of the NEW CST method. The NEW CST method has been evaluated in fitting precision of 1,545 aerofoils by comparison with the LEM CST method and the original CST method. And the improved ability of the NEW CST at the leading edge and trailing edge is verified by a series of complex aerofoil case studies within 1,545 aerofoils. The results indicate that the NEW CST method can represent the whole aerofoils and possesses the intuitive property as well as the original CST. Moreover, the number of control parameters (NCP) to parameterise aerofoils is the fewest among these three methods. Furthermore, when the NCP of the NEW CST and LEM CST is the same, the NEW CST method has the higher accuracy and smaller root mean square errors (RMSE) especially at the leading edge and trailing edge.
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Watt, G. D., and G. V. Parkinson. "On the application of linearised theory to multi-element aerofoils Part II: Effects of thickness, camber and stagger." Aeronautical Journal 90, no. 899 (November 1986): 339–56. http://dx.doi.org/10.1017/s0001924000015943.
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SummaryA linearised two-dimensional incompressible potential flow theory for two-element aerofoil sections is developed. It is capable of predicting the effects of angle of attack, flap deflection, camber, thickness, stagger and overlap of aerofoil elements on the forces. These effects are summarised in integrals which are analogous to the one-element thin aerofoil theory Munk integrals. Analytical expressions for the forces on tandem NACA 23012 aerofoils have been derived and results are presented. Comparisons are also made with a realistic slotted flap configuration with overlap for which experimental data is available. The linearised theory is seen to correlate with these real flow results as well as and better than it does in the one-element regime.
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Morris, Wallace J., and Zvi Rusak. "Stall onset on aerofoils at low to moderately high Reynolds number flows." Journal of Fluid Mechanics 733 (September 24, 2013): 439–72. http://dx.doi.org/10.1017/jfm.2013.440.
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AbstractThe inception of leading-edge stall on stationary, two-dimensional, smooth, thin aerofoils at low to moderately high chord Reynolds number flows is investigated by a reduced-order, multiscale model problem via numerical simulations. The asymptotic theory demonstrates that a subsonic flow about a thin aerofoil can be described in terms of an outer region, around most of the aerofoil’s chord, and an inner region, around the nose, that asymptotically match each other. The flow in the outer region is dominated by the classical thin aerofoil theory. Scaled (magnified) coordinates and a modified (smaller) Reynolds number $(R{e}_{M} )$ are used to correctly account for the nonlinear behaviour and extreme velocity changes in the inner region, where both the near-stagnation and high suction areas occur. It results in a model problem of a uniform, incompressible and viscous flow past a semi-infinite parabola with a far-field circulation governed by a parameter $\tilde {A} $ that is related to the aerofoil’s angle of attack, nose radius of curvature, thickness ratio, and camber. The model flow problem is solved for various values of $\tilde {A} $ through numerical simulations based on the unsteady Navier–Stokes equations. The value ${\tilde {A} }_{s} $ where a global separation zone first erupts in the nose flow, accompanied by loss of peak streamwise velocity ahead of it and change in shedding frequency behind it, is determined as a function of $R{e}_{M} $. These values indicate the stall onset on the aerofoil at various flow conditions. It is found that ${\tilde {A} }_{s} $ decreases with $R{e}_{M} $ until some limit $R{e}_{M} $ (${\sim }300$) and then increases with further increase of Reynolds number. At low values of $R{e}_{M} $ the flow is laminar and steady, even when stall occurs. The flow in this regime is dominated by the increasing effect of the adverse pressure gradient, which eventually overcomes the ability of the viscous stress to keep the boundary layer attached to the aerofoil. The change in the nature of stall at the limit $R{e}_{M} $ is attributed to the appearance of downstream travelling waves in the boundary layer that shed from the marginal separation zone and grow in size with either $\tilde {A} $ or $R{e}_{M} $. These unsteady, convective vortical structures relax the effect of the adverse pressure gradient on the viscous boundary layer to delay the onset of stall in the mean flow to higher values of ${\tilde {A} }_{s} $. Computed results show agreement with marginal separation theory at low $R{e}_{M} $ and with available experimental data at higher $R{e}_{M} $. This simplified approach provides a universal criterion to determine the stall angle of stationary thin aerofoils with a parabolic nose.
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Raghunathan, S., T. Setoguchi, and K. Kaneko. "Predictions of Aerodynamic Performance of Wells Turbines From Aerofoil Data." Journal of Turbomachinery 112, no. 4 (October 1, 1990): 792–95. http://dx.doi.org/10.1115/1.2927723.
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LEPPINGTON, F. G., and R. A. SISSON. "On the interaction of a moving hollow vortex with an aerofoil, with application to sound generation." Journal of Fluid Mechanics 345 (August 25, 1997): 203–26. http://dx.doi.org/10.1017/s0022112097006253.
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A hollow vortex in the form of a straight tube, parallel to the z-axis, and of radius a, moves in a uniform stream of fluid with velocity U in the x-direction, with U small compared with the sound speed c. This steady flow is disturbed by the presence of a thin symmetric fixed aerofoil. With a change of x-coordinate, the problem is equivalent to that of a moving aerofoil cutting through an initially fixed vortex in still fluid. The aim of this work is to determine the resulting perturbed flow, and to estimate the distant sound field. A detailed calculation is given for the perturbed velocity potential in the incompressible flow case, for the linearized equations in the limit of small aerofoil thickness. A formally exact solution involves a four-fold integral and an infinite sum over all mode numbers. For the important special case where the vortex tube has small radius a compared with the aerofoil width, the deformed vortex is characterized by a hypothetical vortex filament located at the ‘mean centre’ x¯(z, t), y¯(z, t) of the tube. Explicit results are given for x¯(z, t), y¯(z, t) for the case where the aerofoil has the elementary rectangular profile; results can then be obtained for more general and realistic cylindrical aerofoils by a single integral weighted with the derivative of the aerofoil thickness function. Finally the distant sound field is estimated, representing the aerofoil by a distribution of moving monopole sources and representing the effect of the deformed vortex in terms of compressible dipoles along the mean centre of the vortex.
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Crowdy, Darren. "Calculating the lift on a finite stack of cylindrical aerofoils." Proceedings of the Royal Society A: Mathematical, Physical and Engineering Sciences 462, no. 2069 (January 24, 2006): 1387–407. http://dx.doi.org/10.1098/rspa.2005.1631.
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The classic exact solution due to Lagally (Lagally, M. 1929 Die reibungslose strömung im aussengebiet zweier kreise. Z. Angew. Math. Mech . 9 , 299–305.) for streaming flow past two cylindrical aerofoils (or obstacles) is generalized to the case of an arbitrary finite number of cylindrical aerofoils. Given the geometry of the aerofoils, the speed and direction of the oncoming uniform flow and the individual round-aerofoil circulations, the complex potential associated with the flow is found in analytical form in a parametric pre-image region that can be conformally mapped to the fluid region. A complete determination of the flow then follows from knowledge of the conformal mapping between the two regions. In the special case where the aerofoils are all circular, the conformal mapping from the parametric pre-image region to the fluid domain is a Möbius mapping. The solution for the complex potential in such a case can then be used, in combination with the Blasius theorem, to compute the distribution of hydrodynamic forces on the multi-aerofoil configuration.
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Marimuthu, Siva, and Dhavamani Chinnathambi. "Computational analysis to enhance the compressible flow over an aerofoil surface." Aircraft Engineering and Aerospace Technology 93, no. 5 (July 5, 2021): 925–34. http://dx.doi.org/10.1108/aeat-06-2020-0122.
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Purpose Since the inception of aerospace engineering, reducing drag is of eternal importance. Over the years, researchers have been trying to improve the aerodynamics of National Advisory Committee for Aeronautics (NACA) aerofoils in many ways. It is proved that smooth-surfaced NACA 0012 aerofoil produces more drag in compressible flow. Recent research on shark-skin pattern warrants a feasible solution to many fluid-engineering problems. Several attempts were made by many researchers to implement the idea of shark skin in the form of coatings, texture and more. However, those ideas are at greater risk when it comes to wing maintenance. The purpose of this paper is to implement a relatively larger biomimetic pattern which would make way for easy maintenance of patterned wings with improved performance. Design/methodology/approach In this paper, two biomimetic aerofoils are designed by optimizing the surface pattern of shark skin and are tested at different angles of attack in the computational flow domain. Findings The results of the biomimetic aerofoils prove that viscous and total drag can be reduced up to 33.08% and 3.68%, respectively, at high subsonic speed when validated against a NACA 0012 aerofoil. With the ample effectiveness of patched shark-skin pattern, biomimetic aerofoil generates as high as 10.42% lift than NACA 0012. Originality/value In this study, a feasible shark-skin pattern is constructed for NACA 0012 in a transonic flow regime. Computational results achieved using the theoretical model agree with experimental data.
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Marimuthu, Siva, Samer Al-Rabeei, and Hithim Ahmed Boha. "Three-Dimensional Analysis of Biomimetic Aerofoil in Transonic Flow." Biomimetics 7, no. 1 (January 22, 2022): 20. http://dx.doi.org/10.3390/biomimetics7010020.
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Since the invention of the aircraft, there has been a need for better surface design to enhance performance. This thirst has driven many aerodynamicists to develop various types of aerofoils. Most researchers have strongly assumed that smooth surfaces would be more suitable for air transport vehicles. This ideology was shattered into pieces when biomimetics was introduced. Biomimetics emphasized the roughness of a surface instead of smoothness in a fluid flow regime. In this research, the most popular 0012 aerofoils of the National Advisory Committee for Aeronautics (NACA) are considered to improve them, with the help of a surface pattern derived from the biological environment. Original and biomimetic aerofoils were designed in three dimensions with the help of Solidworks software and analyzed in the computational flow domain using the commercial code ANSYS Fluent. The implemented biomimetic rough surface pattern upgraded the NACA 0012 aerofoil design in the transonic flow regime. Lift and viscous forces of the aerofoil improved up to 5.41% and 9.98%, respectively. This research has proved that a surface with a little roughness is better than a smooth surface.
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Wild, Graham. "Is that lift diagram correct? A visual study of flight education literature." Physics Education 58, no. 3 (March 15, 2023): 035018. http://dx.doi.org/10.1088/1361-6552/acbad8.
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Abstract With a complex topic such as aerodynamics, subtle points are critical. In this work images illustrating air flow around wings and aerofoils were studied to explore misunderstandings in aerodynamics education. While these images are common in textbooks and popular science media, this study was limited to the 135 physics education articles on the topic of lift, of which 49 contained illustrations of air flow around an aerofoil or wing. These 49 cases were included for qualitative comparison using visual semiotics. It was found that 28% of images did not include upwash, and only 44% included stagnation points. For the case of 2D flow around aerofoils 30% were illustrated correctly, while for wings 75% were correct. These results excluded the seven completely incorrect illustrations where common misconceptions were presented as facts. Most illustrations of flow around an aerofoil incorrectly depicted flow around a wing.
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Kozic, M. S., and D. Sredojevic. "Development of unstructured dynamic grids for solving unsteady two-dimensional Euler equations." Aeronautical Journal 102, no. 1014 (April 1998): 195–200. http://dx.doi.org/10.1017/s0001924000096305.
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AbstractA method for the solution of the time dependent Euler equations on unstructured grids is presented for unsteady flows about oscillating aerofoils. The flow solver involves a finite volume spatial discretisation and a Runge-Kutta time-stepping scheme. A dynamic mesh algorithm is used for problems where the aerofoil moves and/or deforms. Steady and unsteady results are obtained for single and multi-element aerofoils.
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Bhatia, D., Y. Zhao, D. Yadav, and J. Wang. "Drag Reduction Using Biomimetic Sharkskin Denticles." Engineering, Technology & Applied Science Research 11, no. 5 (October 12, 2021): 7665–72. http://dx.doi.org/10.48084/etasr.4347.
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This paper explores the use of sharkskin in improving the aerodynamic performance of aerofoils. A biomimetic analysis of the sharkskin denticles was conducted and the denticles were incorporated on the surface of a 2-Dimensional (2D) NACA0012 aerofoil. The aerodynamic performance including the drag reduction rate, lift enhancement rate, and Lift to Drag (L/D) enhancement rate for sharkskin denticles were calculated at different locations along the chord line of the aerofoil and at different Angles of Attack (AOAs) through Computational Fluid Dynamics (CFD). Two different denticle orientations were tested. Conditional results indicate that the denticle reduces drag by 4.3% and attains an L/D enhancement ratio of 3.6%.
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Hajian, Rozhin, and Justin W. Jaworski. "The steady aerodynamics of aerofoils with porosity gradients." Proceedings of the Royal Society A: Mathematical, Physical and Engineering Sciences 473, no. 2205 (September 2017): 20170266. http://dx.doi.org/10.1098/rspa.2017.0266.
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This theoretical study determines the aerodynamic loads on an aerofoil with a prescribed porosity distribution in a steady incompressible flow. A Darcy porosity condition on the aerofoil surface furnishes a Fredholm integral equation for the pressure distribution, which is solved exactly and generally as a Riemann–Hilbert problem provided that the porosity distribution is Hölder-continuous. The Hölder condition includes as a subset any continuously differentiable porosity distributions that may be of practical interest. This formal restriction on the analysis is examined by a class of differentiable porosity distributions that approach a piecewise, discontinuous function in a certain parametric limit. The Hölder-continuous solution is verified in this limit against analytical results for partially porous aerofoils in the literature. Finally, a comparison made between the new theoretical predictions and experimental measurements of SD7003 aerofoils presented in the literature. Results from this analysis may be integrated into a theoretical framework to optimize turbulence noise suppression with minimal impact to aerodynamic performance.
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Qin, N., Y. Zhu, and D. I. A. Poll. "Surface suction on aerofoil aerodynamic characteristics at transonic speeds." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 212, no. 5 (May 1, 1998): 339–51. http://dx.doi.org/10.1243/0954410981532324.
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This paper presents a numerical study of the effects of an active flow control through surface suction on shock boundary layer interactions over transonic aerofoils. Two different aerofoils were studied. Firstly, for the purpose of validation, an NACA64A010 aerofoil with a trailing edge flap was investigated and the numerical results were compared with experimental data with and without suction for surface pressure distributions and lift and drag coefficients. Grid sensitivity has also been studied regarding the numerical accuracy. The second geometry was an RAE9647 aerofoil, which was designed for superior aerodynamic performance when applied to a helicopter rotor blade. An active surface was used to prevent or alleviate shock-induced separation. The suction strength and location were varied to determine the effect on aerodynamic performance and to provide an effective means of suppressing undesirable flow features. In both cases, increases in both lift and drag were observed when surface suction was applied. However, the benefit of suction appeared in the form of a substantial increase in the lift-drag ratio. It was also found that the shock location and strength are very sensitive to the suction location and strength. Two different mechanisms for active flow control over transonic aerofoils are discussed.
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Cameron, L., J. Early, R. McRoberts, and M. Price. "Constrained multi-objective aerofoil design using a multi-level optimisation strategy." Aeronautical Journal 119, no. 1217 (July 2015): 833–54. http://dx.doi.org/10.1017/s0001924000010940.
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AbstractA novel approach for the multi-objective design optimisation of aerofoil profiles is presented. The proposed method aims to exploit the relative strengths of global and local optimisation algorithms, whilst using surrogate models to limit the number of computationally expensive CFD simulations required. The local search stage utilises a re-parameterisation scheme that increases the flexibility of the geometry description by iteratively increasing the number of design variables, enabling superior designs to be generated with minimal user intervention. Capability of the algorithm is demonstrated via the conceptual design of aerofoil sections for use on a lightweight laminar flow business jet. The design case is formulated to account for take-off performance while reducing sensitivity to leading edge contamination. The algorithm successfully manipulates boundary layer transition location to provide a potential set of aerofoils that represent the trade-offs between drag at cruise and climb conditions in the presence of a challenging constraint set. Variations in the underlying flow physics between Pareto-optimal aerofoils are examined to aid understanding of the mechanisms that drive the trade-offs in objective functions.
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Lewthwaite, Matthew Thomas, and Chiemela Victor Amaechi. "Numerical Investigation of Winglet Aerodynamics and Dimple Effect of NACA 0017 Airfoil for a Freight Aircraft." Inventions 7, no. 1 (March 7, 2022): 31. http://dx.doi.org/10.3390/inventions7010031.
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Drag reduction is an ever-present challenge within the aeronautical engineering industry. This paper presents two substantial wing modifications: the addition of a winglet of a freighter aircraft and a dimpled wing on the NACA 0017 aerofoils. Studies on nine (9) different geometries of dimpled aerofoils were performed against a control model of an aerofoil without any dimple. Computational fluid dynamics (CFD) analysis was performed using two (2) commercial CFD platforms. This paper also explored two novel solutions of aircraft optimisation to mitigate the effects of drag and leading-edge pressure, while increasing the effect of lift. The optimised performance model of a freighter aircraft increased its aerodynamic efficiency. The study found that at take-off velocity of 82 m/s, winglets decreased pressure on the wing by 16.31%, through flow redirection and better flow integration into aerofoils wake. The study also analysed the separation layer and its effect through the appropriate use of the dimple effect. Increased lift effects were observed on a NACA 0017 aerofoil. Despite the low increase in drag of 6% from the modifications, the resultant L/D ratio was highly increased. This study also faced some challenges with validating the model. Hence some validation approaches were taken, and some other approaches suggested for future studies.
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Setyawan K, Catur, and Djoko Sardjadi. "PENINGKATAN KOEFISIEN GAYA ANGKAT AEROFOIL KENNEDY-MARSDEN DENGAN ZAP FLAP." Jurnal Konversi Energi dan Manufaktur 1, no. 1 (October 24, 2013): 14–21. http://dx.doi.org/10.21009/jkem.1.1.2.
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Makalah ini membahas peningkatan prestasi aerodinamika aerofoil dengan menggunakan zap flap. Penggunaan zap flap ini memungkinkan dipertahankannya bentuk asli aerofoil pada saat flap tidak digunakan. Sedangkan pada saat zap flap dijulurkan, akan terjadi peningkatan gaya angkat akibat bertambah panjangnya chord aerofoil dan berubahnya kelengkungan aerofoil. Pemakaian zap flap ini juga akan membuat aerofoil masih mempunyai sifat sebagai aerofoil satu elemen meskipun flap telah digunakan. Dengan sifat ini diharapkan penambahan gaya angkat yang diperoleh tidak akan berpengaruh banyak terhadap koefisien-koefisien aerodinamika aerofoil yang lain. Aerofoil yang digunakan dalam penelitian ini adalah aerofoil Kennedy-Marsden yang telah dimodifikasi. Dengan memvariasikan panjang zap flap yang dijulurkan, dilakukan analisis perubahan prestasi aerodinamika aerofoil. Pada penelitian ini, analisis karakteristik aerodinamika aerofoil dihasilkan dari perhitungan numerik menggunakan program MSES. Persamaan atur yang dipecahkan dalam solver MSES ini adalah kombinasi persamaan Euler dan persamaan lapisan batas
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Ramesh, Kiran, Ashok Gopalarathnam, Kenneth Granlund, Michael V. Ol, and Jack R. Edwards. "Discrete-vortex method with novel shedding criterion for unsteady aerofoil flows with intermittent leading-edge vortex shedding." Journal of Fluid Mechanics 751 (June 23, 2014): 500–538. http://dx.doi.org/10.1017/jfm.2014.297.
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AbstractUnsteady aerofoil flows are often characterized by leading-edge vortex (LEV) shedding. While experiments and high-order computations have contributed to our understanding of these flows, fast low-order methods are needed for engineering tasks. Classical unsteady aerofoil theories are limited to small amplitudes and attached leading-edge flows. Discrete-vortex methods that model vortex shedding from leading edges assume continuous shedding, valid only for sharp leading edges, or shedding governed by ad-hoc criteria such as a critical angle of attack, valid only for a restricted set of kinematics. We present a criterion for intermittent vortex shedding from rounded leading edges that is governed by a maximum allowable leading-edge suction. We show that, when using unsteady thin aerofoil theory, this leading-edge suction parameter (LESP) is related to the $\def \xmlpi #1{}\def \mathsfbi #1{\boldsymbol {\mathsf {#1}}}\let \le =\leqslant \let \leq =\leqslant \let \ge =\geqslant \let \geq =\geqslant \def \Pr {\mathit {Pr}}\def \Fr {\mathit {Fr}}\def \Rey {\mathit {Re}}A_0$ term in the Fourier series representing the chordwise variation of bound vorticity. Furthermore, for any aerofoil and Reynolds number, there is a critical value of the LESP, which is independent of the motion kinematics. When the instantaneous LESP value exceeds the critical value, vortex shedding occurs at the leading edge. We have augmented a discrete-time, arbitrary-motion, unsteady thin aerofoil theory with discrete-vortex shedding from the leading edge governed by the instantaneous LESP. Thus, the use of a single empirical parameter, the critical-LESP value, allows us to determine the onset, growth, and termination of LEVs. We show, by comparison with experimental and computational results for several aerofoils, motions and Reynolds numbers, that this computationally inexpensive method is successful in predicting the complex flows and forces resulting from intermittent LEV shedding, thus validating the LESP concept.
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Mohammad, Hussein K., S. M. Jali, Arz Y. Qwam Alden Qwam Alden, Viktor Kilchyk, and Bade Shrestha. "Experimental and Numerical Investigation of Channeling Effects on Noise and Aerodynamic Performance of NACA 0012 Aerofoil in Wind Turbine Applications." Renewable Energy and Power Quality Journal 21, no. 1 (July 2023): 373–80. http://dx.doi.org/10.24084/repqj21.328.
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Wind power is considered one of the main sources of renewable energy in the market today. In this study, different sizes and directions of channels were created inside the NACA 0012 aerofoil, and the effect of these channels were investigated on aerodynamic noise and aerodynamic performance, experimentally and numerically. The results have shown several factors that could affect the aerodynamic noise such as flow velocity, angle of attack, and trailing edge blowing injection. The study also concluded an increase in drag coefficients and a decrease in lift coefficients for all channeled samples compared to the regular aerofoil. In contrast to the studies that showed improvements in the aerodynamic performance of supersonic channeled aerofoils, this study done under subsonic flow showed an increase in drag and decrease in lift.
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Tam, Christopher K. W., and Hongbin Ju. "Aerofoil tones at moderate Reynolds number." Journal of Fluid Mechanics 690 (December 1, 2011): 536–70. http://dx.doi.org/10.1017/jfm.2011.465.
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AbstractIt is known experimentally that an aerofoil immersed in a uniform stream at a moderate Reynolds number emits tones. However, there have been major differences in the experimental observations in the past. Some experiments reported the observation of multiple tones, with strong evidence that these tones are most probably generated by a feedback loop. There is also an experiment reporting the observation of a single tone with no tonal jump or other features associated with feedback. In spite of the obvious differences in the experimental observations published in the literature, it is noted that all the dominant tone frequencies measured in all the investigations are in agreement with an empirically derived Paterson formula. The objective of the present study is to perform a direct numerical simulation (DNS) of the flow and acoustic phenomenon to investigate the tone generation mechanism. When comparing with experimental studies, numerical simulations appear to have two important advantages. The first is that there is no background wind tunnel noise in numerical simulation. This avoids the signal-to-noise ratio problem inherent in wind tunnel experiments. In other words, it is possible to study tones emitted by a truly isolated aerofoil computationally. The second advantage is that DNS produces a full set of space–time data, which can be very useful in determining the tone generation processes. The present effort concentrates on the tones emitted by three NACA0012 aerofoils with a slightly rounded trailing edge but with different trailing edge thickness at zero degree angle of attack. At zero degree angle of attack, in the Reynolds number range of$2\ensuremath{\times} 1{0}^{5} $to$5\ensuremath{\times} 1{0}^{5} $, the boundary layer flow is attached nearly all the way to the trailing edge of the aerofoil. Unlike an aerofoil at an angle of attack, there is no separation bubble, no open flow separation. All the flow separation features tend to increase the complexity of the tone generation processes. The present goal is limited to finding the basic tone generation mechanism in the simplest flow configuration. Our DNS results show that, for the flow configuration under study, the aerofoil emits only a single tone. This is true for all three aerofoils over the entire Reynolds number range of the present study. In the literature, it is known that Kelvin–Helmholtz instabilities of free shear layers generally have a much higher spatial growth rate than that of the Tollmien–Schlichting boundary layer instabilities. A near-wake non-parallel flow instability analysis is performed. It is found that the tone frequencies are the same as the most amplified Kelvin–Helmholtz instability at the location where the wake has a minimum half-width. This suggests that near-wake instability is the energy source of aerofoil tones. However, flow instabilities at low subsonic Mach numbers generally do not cause strong tones. An investigation of how near-wake instability generates tones is carried out using the space–time data provided by numerical simulations. Our observations indicate that the dominant tone generation process is the interaction of the oscillatory motion of the near wake, driven by flow instability, with the trailing edge of the aerofoil. Secondary mechanisms involving unsteady near-wake motion and the formation of discrete vortices in regions further downstream are also observed.
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Gracey, M. W., A. J. Niven, F. N. Coton, R. A. McD. Galbraith, and D. Jiang. "A correlation indicating incipient dynamic stall." Aeronautical Journal 100, no. 997 (September 1996): 305–11. http://dx.doi.org/10.1017/s0001924000028955.
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AbstractThis paper proposes a correlation which attempts to relate an aerofoil's low speed dynamic stall onset incidence to particular parameters describing its stall behaviour in steady conditions. These parameters, which can be derived from experiment or predictive algorithm, are the incidence of steady stall and a term related to trailing-edge separation characteristics. The correlation is based on a large amount of data obtained from a number of aerofoils broadly classified into two families: the NACA four digit series of symmetrical sections and a family of four profiles with the NACA 23012 as the generic shape.The correlation has been extended to low pitch rates outside the “deep” dynamic stall range and, whilst it pertains to stalling under ramp motions, a link has been established with oscillatory motions. The potential of the correlation is demonstrated by its application to the design of a symmetric aerofoil for use on wind turbines.
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Ayton, Lorna J., and Paruchuri Chaitanya. "Analytical and experimental investigation into the effects of leading-edge radius on gust–aerofoil interaction noise." Journal of Fluid Mechanics 829 (September 26, 2017): 780–808. http://dx.doi.org/10.1017/jfm.2017.594.
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This paper investigates the effects of local leading-edge geometry on unsteady aerofoil interaction noise. Analytical results are obtained by extending previous work for parabolic leading edges to leading edges of the form $x^{m}$ for $0<m<1$. Rapid distortion theory governs the interaction of an unsteady vortical perturbation with a rigid aerofoil in compressible steady mean flow that is uniform far upstream. For high-frequency gusts interacting with aerofoils of small total thickness this allows a matched asymptotic solution to be obtained. This paper mainly focusses on obtaining the analytic solution in the leading-edge inner region, which is the dominant term in determining the total far-field acoustic directivity, and contains the effects of the local leading-edge geometry. Experimental measurements for the noise generated by aerofoils with different leading-edge nose radii in uniform flow with approximate homogeneous, isotropic turbulence are also presented. Both experimental and analytic results predict that a larger nose radius generates less overall noise in low-Mach-number flow. By considering individual terms in the analytic solution, this paper is able to propose reasons behind this result.
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Domel, August G., Mehdi Saadat, James C. Weaver, Hossein Haj-Hariri, Katia Bertoldi, and George V. Lauder. "Shark skin-inspired designs that improve aerodynamic performance." Journal of The Royal Society Interface 15, no. 139 (February 2018): 20170828. http://dx.doi.org/10.1098/rsif.2017.0828.
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There have been significant efforts recently aimed at improving the aerodynamic performance of aerofoils through the modification of their surfaces. Inspired by the drag-reducing properties of the tooth-like denticles that cover the skin of sharks, we describe here experimental and simulation-based investigations into the aerodynamic effects of novel denticle-inspired designs placed along the suction side of an aerofoil. Through parametric modelling to query a wide range of different designs, we discovered a set of denticle-inspired surface structures that achieve simultaneous drag reduction and lift generation on an aerofoil, resulting in lift-to-drag ratio improvements comparable to the best-reported for traditional low-profile vortex generators and even outperforming these existing designs at low angles of attack with improvements of up to 323%. Such behaviour is enabled by two concurrent mechanisms: (i) a separation bubble in the denticle's wake altering the flow pressure distribution of the aerofoil to enhance suction and (ii) streamwise vortices that replenish momentum loss in the boundary layer due to skin friction. Our findings not only open new avenues for improved aerodynamic design, but also provide new perspective on the role of the complex and potentially multifunctional morphology of shark denticles for increased swimming efficiency.
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Benner, M. W., S. A. Sjolander, and S. H. Moustapha. "The Influence of Leading-Edge Geometry on Secondary Losses in a Turbine Cascade at the Design Incidence." Journal of Turbomachinery 126, no. 2 (April 1, 2004): 277–87. http://dx.doi.org/10.1115/1.1645533.
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The paper presents detailed experimental results of the secondary flows from two large-scale, low-speed linear turbine cascades. The aerofoils for the two cascades were designed for the same inlet and outlet conditions and differ mainly in their leading-edge geometries. Detailed flow field measurements were made upstream and downstream of the cascades using three and seven-hole pressure probes and static pressure distributions were measured on the aerofoil surfaces. All measurements were made exclusively at the design incidence. The results from this experiment suggest that the strength of the passage vortex plays an important role in the downstream flow field and loss behavior. It was concluded that the aerofoil loading distribution has a significant influence on the strength of this vortex. In contrast, the leading-edge geometry appears to have only a minor influence on the secondary flow field, at least for the design incidence.
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Muller, Jan. "Improving Initial Aerofoil Geometry Using Aerofoil Particle Swarm Optimisation." MENDEL 28, no. 1 (June 30, 2022): 63–67. http://dx.doi.org/10.13164/mendel.2022.1.063.
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Advanced optimisation of the aerofoil wing of a general aircraft is the main subject of this paper. Meta-heuristic optimisation techniques, especially swarm algorithms, were used. Subsequently, a new variant denoted as aerofoil particle swarm optimisation (aPSO) was developed from the original particle swarm optimisation (PSO). A parametric model based on B-spline was used to optimise the initial aerofoil. The simulation software Xfoil was calculating basic aerodynamic features (lift, drag, moment).
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Narkiewicz, J. P., A. Ling, and G. T. S. Done. "Unsteady aerodynamic loads on an aerofoil with a deflecting tab." Aeronautical Journal 99, no. 987 (September 1995): 282–92. http://dx.doi.org/10.1017/s0001924000028463.
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AbstractA thin, low cambered aerofoil with a tab at the trailing edge in inviscid, incompressible, two-dimensional, unstalled flow with varying freestream velocity and arbitrary, different motions for both aerofoil and tab is considered. The general expressions for unsteady lift and aerodynamic moment on aerofoil and tab are derived within the assumptions of potential theory.To verify the approach proposed in this paper, the classical Theodorsen case of an oscillating aerofoil is adopted. For this case, aerofoil loads are calculated in the time domain by applying an inverse Laplace transformation to the approximation of the lift deficiency function in the frequency domain.The comparison of the results calculated by this method with those obtained by other methods and experiments shows good agreement, which validates the general formulation. The loads on an aerofoil having different frequencies for main aerofoil pitch and tab deflection are calculated.
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NASH, EMMA C., MARTIN V. LOWSON, and ALAN McALPINE. "Boundary-layer instability noise on aerofoils." Journal of Fluid Mechanics 382 (March 10, 1999): 27–61. http://dx.doi.org/10.1017/s002211209800367x.
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An experimental and theoretical investigation has been carried out to understand the tonal noise generation mechanism on aerofoils at moderate Reynolds number. Experiments were conducted on a NACA0012 aerofoil section in a low-turbulence closed working section wind tunnel. Narrow band acoustic tones were observed up to 40 dB above background noise. The ladder structure of these tones was eliminated by modifying the tunnel to approximate to anechoic conditions. High-resolution flow velocity measurements have been made with a three-component laser-Doppler anemometer (LDA) which have revealed the presence of strongly amplified boundary-layer instabilities in a region of separated shear flow just upstream of the pressure surface trailing edge, which match the frequency of the acoustic tones. Flow visualization experiments have shown these instabilities to roll up to form a regular Kármán-type vortex street.A new mechanism for tonal noise generation has been proposed, based on the growth of Tollmien–Schlichting (T–S) instability waves strongly amplified by inflectional profiles in the separating laminar shear layer on the pressure surface of the aerofoil. The growth of fixed frequency, spatially growing boundary-layer instability waves propagating over the aerofoil pressure surface has been calculated using experimentally obtained boundary-layer characteristics. The effect of boundary-layer separation has been incorporated into the model. Frequency selection and prediction of T–S waves are in remarkably good agreement with experimental data.
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Ayton, Lorna J., and N. Peake. "On high-frequency noise scattering by aerofoils in flow." Journal of Fluid Mechanics 734 (October 8, 2013): 144–82. http://dx.doi.org/10.1017/jfm.2013.477.
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AbstractA theoretical model is developed for the sound scattered when a sound wave is incident on a cambered aerofoil at non-zero angle of attack. The model is based on the linearization of the Euler equations about a steady subsonic flow, and is an adaptation of previous work which considered incident vortical disturbances. Only high-frequency sound waves are considered. The aerofoil thickness, camber and angle of attack are restricted such that the steady flow past the aerofoil is a small perturbation to a uniform flow. The singular perturbation analysis identifies asymptotic regions around the aerofoil; local ‘inner’ regions, which scale on the incident wavelength, at the leading and trailing edges of the aerofoil; Fresnel regions emanating from the leading and trailing edges of the aerofoil due to the coalescence of singularities and points of stationary phase; a wake transition region downstream of the aerofoil leading and trailing edge; and an outer region far from the aerofoil and wake. An acoustic boundary layer on the aerofoil surface and within the transition region accounts for the effects of curvature. The final result is a uniformly-valid solution for the far-field sound; the effects of angle of attack, camber and thickness are investigated.
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Patel, Karna S., Saumil B. Patel, Utsav B. Patel, and Prof Ankit P. Ahuja. "CFD Analysis of an Aerofoil." International Journal of Engineering Research 3, no. 3 (March 1, 2014): 154–58. http://dx.doi.org/10.17950/ijer/v3s3/305.
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Dahdi, B., M. Mamou, M. Khalid, S. Benissaad, and Z. Nemouchi. "Investigation of skin porosity damping effects on free stream disturbance induced unsteady wing loads." Aeronautical Journal 116, no. 1184 (October 2012): 1041–60. http://dx.doi.org/10.1017/s0001924000007478.
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Abstract Numerical simulations were performed to analyse the possibility of damping abrupt incoming free stream disturbances upon a porous aerofoil using an unsteady Reynolds-averaged Navier-Stokes (URANS) model. To mimic the turbulence disturbance levels that are typically encountered in the atmosphere, two flow configurations were considered. In the first configuration, the unsteadiness of the flow was created with vortices shed from a circular cylinder installed ahead of a WTEA-TE1 aerofoil. The continuous von Kármán shedding vortices contained within the cylinder wake were convected downstream and projected upon the aerofoil. In the second configuration, an instantaneous pair of discrete vortices was created by a rotational snapping of a flat plate, installed upstream of the aerofoil. Solid and porous aerofoil configurations, with porosity settings of 11 and 22%, were applied on 50% of the chord of the aerofoil starting from the leading edge. Both steady and unsteady flow simulations were performed to assess the performance of the porosity under steady and unsteady effects. The steady state flow simulations revealed a noticeable reduction in the aerofoil lift coefficient for the porous aerofoil. For unsteady solutions with a continuous or distinct series of vortices interacting with the aerofoil, the porosity showed insignificant damping of the lift coefficient amplitude. The porosity values investigated in the current exercise had indiscernible effect upon the unsteady lift-load alleviations caused by free stream disturbances.
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Aoki, Makoto. "Helicopter blade aerofoil." Journal of the Acoustical Society of America 108, no. 1 (2000): 20. http://dx.doi.org/10.1121/1.429485.
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