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Journal articles on the topic 'Supercritical Airfoil'

Author: Grafiati

Published: 4 June 2021

Last updated: 1 February 2022

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1

Bazgir, Ali S., and Sergey A. Takovitskii. "Increase of Critical Mach Number by Local Linearization Method and Airfoil Construction at Subsonic and Transonic Regimes." MATEC Web of Conferences 221 (2018): 05001. http://dx.doi.org/10.1051/matecconf/201822105001.

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A symmetrical airfoil has been constructed by local linearization method. A single-point objective function is defined to check the convergence of the method. As an example, the nose and tail zone of supercritical airfoil is fixed and a flat line is placed between them. The optimizable element of the airfoil contour was conjoined with the nose and tail elements of fixed shape at the sections with coordinates xs1= 0.11 and xs2= 0.66, respectively. The optimizable part of airfoil (the fixed chord line) is divided into N=55 segments. The convergence of this method has been shown with the airfoil constructed with higher critical Mach number rather than the initial airfoil. Finally, this airfoil has been compared with the supercritical airfoil NASA SC (2)-0012 at M∞=0.76. At the second part, several airfoils have been constructed and simulated over different Subsonic and Transonic Mach numbers. Finally, the drag coefficient on constructed airfoils have been compared with supercritical airfoil.

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2

Zhang, Yufei, Chongyang Yan, and Haixin Chen. "An Inverse Design Method for Airfoils Based on Pressure Gradient Distribution." Energies 13, no. 13 (July 2, 2020): 3400. http://dx.doi.org/10.3390/en13133400.

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An airfoil inverse design method is proposed by using the pressure gradient distribution as the design target. The adjoint method is used to compute the derivatives of the design target. A combination of the weighted drag coefficient and the target dimensionless pressure gradient is applied as the optimization objective, while the lift coefficient is considered as a constraint. The advantage of this method is that the designer can sketch a rough expectation of the pressure distribution pattern rather than a precise pressure coefficient under a certain lift coefficient and Mach number, which can greatly reduce the design iteration in the initial stage of the design process. Multiple solutions can be obtained under different objective weights. The feasibility of the method is validated by a supercritical airfoil and a supercritical natural laminar flow airfoil, which are designed based on the target pressure gradients on the airfoils. Eight supercritical airfoils are designed under different upper surface pressure gradients. The drag creep and drag divergence characteristics of the airfoils are numerically tested. The shockfree airfoil demonstrates poor performance because of a high suction peak and the double-shock phenomenon. The adverse pressure gradient on the upper surface before the shockwave needs to be less than 0.2 to maintain both good drag creep and drag divergence characteristics.

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3

Liao, Yan Ping, Li Liu, and Teng Long. "Investigation of Various Parametric Geometry Representation Methods for Airfoils." Applied Mechanics and Materials 110-116 (October 2011): 3040–46. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.3040.

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Abstract—This paper presents the investigation of typical parametric geometry representation methods for airfoils, namely, PARSEC method, orthogonal basis function method and CST method. The investigation assesses the fitting accuracy of these parametric methods for various airfoils including the symmetric airfoil, cambered airfoil and supercritical airfoil. The design variables of these parametric methods are solved by the methods of least squares fit. The fitting results show that the fitting accuracy of CST method is better than other parametric methods for airfoil. The aerodynamics analysis models of these typical parametric geometry representation methods for airfoil are constructed. The pressure distributions calculated for different parametric methods are compared with the corresponding experimental pressure distributions for the actual airfoil geometry.Keywords-orthogonal basis function; PARSEC; CST; fitting accuracy; pressure distributions

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4

Al-Jaburi, Khider, and Daniel Feszty. "Fixed and rotary wing transonic aerodynamic improvement via surface-based trapped vortex generators." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 15 (June 12, 2019): 5522–42. http://dx.doi.org/10.1177/0954410019853902.

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A novel passive flow control concept for transonic flows over airfoils is proposed and examined via computational fluid dynamics. The control concept is based on the local modification of the airfoil's geometry. It aims to reduce drag or to increase lift without deteriorating the original lift and/or drag characteristics of the airfoil, respectively. Such flow control technique could be beneficial for improving the range or endurance of transonic aircraft or for mitigating the negative effects of transonic flow on the advancing blades of helicopter rotors. To explore the feasibility of the concept, two-dimensional computational fluid dynamics simulations of a NACA 0012 airfoil exposed to a freestream of Mach 0.7 and Re = 9 × 106 as well as of a NASA SC(3)−0712(B) supercritical airfoil exposed to a freestream of Mach 0.78 and Re = 30 × 106 were conducted. The baseline airfoil simulations were carefully verified and validated, showing excellent agreement with wind tunnel data. Then, 32 various local geometry modifications were proposed and systematically examined, all functioning as a trapped-vortex generator. The surface modifications were examined on both the upper and lower surfaces of the airfoils. The upper surface modifications demonstrated remarkable ability to reduce the strength of the shockwave on the upper surface of the airfoil with only a small penalty in lift. On the other hand, the lower surface modifications could significantly increase the lift-to-drag ratio for the full range of the investigated angles of attack, when compared to the baseline airfoil.

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5

Sonoda, Toyotaka, and Heinz-Adolf Schreiber. "Aerodynamic Characteristics of Supercritical Outlet Guide Vanes at Low Reynolds Number Conditions." Journal of Turbomachinery 129, no. 4 (August 19, 2006): 694–704. http://dx.doi.org/10.1115/1.2720868.

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As a part of an innovative aerodynamic design concept for a single stage low pressure turbine, a high turning outlet guide vane is required to remove the swirl from the hot gas. The airfoil of the vane is a highly loaded compressor airfoil that has to operate at very low Reynolds numbers (Re∼120,000). Recently published numerical design studies and experimental analysis on alternatively designed airfoils showed that blade profiles with an extreme front loaded pressure distribution are advantageous for low Reynolds number conditions. The advantage even holds true for an increased inlet Mach number at which the peak Mach number on the airfoils reaches and exceeds the critical conditions (Mss>1.0). This paper discusses the effect of the inlet Mach number and Reynolds number on the cascade performance for both a controlled diffusion airfoil (CDA) (called baseline) and a numerically optimized front loaded airfoil. The results show that it is advantageous to design the profile with a fairly steep pressure gradient immediately at the front part in order to promote early transition or to prevent too large laminar—even shock induced—separations with the risk of a bubble burst. Profile Mach number distributions and wake traverse data are presented for design and off-design conditions. The discussion of Mach number distributions and boundary layer behavior is supported by numerical results obtained from the blade-to-blade flow solver MISES.

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6

Xu, Xin, Da Wei Liu, De Hua Chen, and Yuan Jing Wang. "Reynolds Number Effect Investigation of Shock Wave on Supercritical Airfoil." Applied Mechanics and Materials 548-549 (April 2014): 520–24. http://dx.doi.org/10.4028/www.scientific.net/amm.548-549.520.

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The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. But at transonic speeds, the shock wave on upper surface of supercritical airfoil may induce boundary layer separation, which would change the aerodynamic characteristics. The shock characteristics such as location and intensity are sensitive to Reynolds number. In order to predict aerodynamic characteristics of supercritical airfoil exactly, the Reynolds number effects of shock wave must be investigated.The transonic flows over a typical supercritical airfoil CH were numerically simulated with two-dimensional Navier-Stokes equations, and the numerical method was validated with test results in ETW(European Transonic Windtunnel). The computation attack angles of CH airfoil varied from 0oto 8o, Mach numbers varied from 0.74 to 0.82 while Reynolds numbers varied from 3×106 to 50×106 per airfoil chord. It is obvious that shock location moves afterward and shock intensity strengthens as Reynolds number increasing. The similar curves of shock location and intensity is linear with logarithm of Reynolds number, so that the shock location and intensity at flight condition could be extrapolated from low Reynolds number.

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7

Zhang, Z. Y., X. T. Yang, and B. Laschka. "Design of a supercritical airfoil." Journal of Aircraft 25, no. 6 (June 1988): 503–6. http://dx.doi.org/10.2514/3.45613.

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8

Liu, Da Wei, Xin Xu, Zhi Wei, and De Hua Chen. "Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results." Applied Mechanics and Materials 444-445 (October 2013): 517–23. http://dx.doi.org/10.4028/www.scientific.net/amm.444-445.517.

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Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106to 30×106per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

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9

Xu, Xin, Da Wei Liu, De Hua Chen, and Yuan Jing Wang. "Numerical Investigation on Shock-Induced Separation Structure of Supercritical Airfoil." Advanced Materials Research 756-759 (September 2013): 4502–5. http://dx.doi.org/10.4028/www.scientific.net/amr.756-759.4502.

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The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. But at transonic speeds, the complicated shock-induced separation on the upper surface of supercritical airfoil will change the aerodynamic characteristics. The transonic flows over a typical supercritical airfoil CH were numerically investigated in this paper, in order to analyses different shock-induced separation structure. The two-dimensional Navier-Stokes equations were solved with structure grids by utilizing the S-A turbulence model. The computation attack angles of CH airfoil varied from 0oto 4o, Mach numbers varied from 0.74 to 0.82 while Reynolds numbers varied from 3×106to 50×106per airfoil chord. It is shown that with the attack angle increases, the separation bubble occurred on the upper surface first, then the trailing-edge separation occurred, the trailing-edge would separate totally at last. The different separation structure would result in different pressure coefficient distribution and boundary layer thickness.

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10

Nakayama, A. "Characteristics of the flow around conventional and supercritical airfoils." Journal of Fluid Mechanics 160 (November 1985): 155–79. http://dx.doi.org/10.1017/s0022112085003433.

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Measurements of the mean and fluctuating velocities have been obtained with pressure and hot-wire probes in the attached boundary layers and wakes of two airfoil models at a low Mach number. The first model is a conventional airfoil at zero incidence and the second an advanced supercritical airfoil at an angle of attack of 4°. The mean-flow and Reynolds-stress data and related quantities are presented with emphasis on the trailing-edge region. The results indicate that the flow around the conventional airfoil is a minor perturbation of a symmetric flat-plate flow with small wake curvature and weak viscous–inviscid interaction. The flow around the supercritical airfoil is in considerable contrast with strong streamwise pressure gradients, non-negligible normal pressure gradients, and large surface and streamline curvatures of the trailing-edge flow. The near wake is strongly curved and intense mixing occurs between the retarded upper-surface boundary layer and strongly accelerated lower-surface boundary layer.

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11

Geissler, W., G. Dietz, and H. Mai. "Dynamic stall on a supercritical airfoil." Aerospace Science and Technology 9, no. 5 (July 2005): 390–99. http://dx.doi.org/10.1016/j.ast.2005.01.012.

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12

Wang, Ming-Yang, Zi-Liang Li, Sheng-Feng Zhao, Yan-Feng Zhang, and Xin-Gen Lu. "Effects of Reynolds number and loading distribution on the aerodynamic performance of a high subsonic compressor airfoil." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 234, no. 8 (January 13, 2020): 1069–83. http://dx.doi.org/10.1177/0957650919899541.

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The laminar-turbulent transition process on the compressor blade surface is often induced by the laminar separation flow at low Reynolds number ( Re). In the present study, numerical simulations were conducted to investigate the structure of the laminar separation bubble and its effects on the profile loss of a high subsonic compressor airfoil under different Re conditions, and the mechanism for the performance deterioration of compressor airfoil at low Re was clarified. Besides, the airfoil was redesigned to obtain a series of airfoils with different loading distributions, and the aerodynamic performance of these airfoils was compared and analyzed in detail. According to the simulation results, the laminar separation bubble mainly determined the loss generation process of a compressor airfoil. When Re decreased from 12 × 105 to 1.5 × 105, the laminar separation bubble on the suction surface grew thicker and the length was increased by 11.2% of the axial chord. As such, the reversed flow inside the laminar separation bubble became more obvious and the turbulence level downstream of the maximum thickness of laminar separation bubble was increased. Also, the growth in the turbulent boundary layer was enhanced, causing more serious flow blockage and wake mixing. According to the Denton's profile loss model, the larger trailing edge loss caused by the stronger displacement effect of laminar separation bubble was supposed to be the main reason for the performance deterioration of compressor airfoil under low Re conditions. The ultra-front loading distribution for airfoil has the possibility to suppress or even eliminate the negative effect of laminar separation bubble, and the profile loss was decreased by 26.7% at Re = 1.5 × 105; however, the less significant performance improvement was observed at some higher Re. Moreover, the ultra-front loaded airfoil was less sensitive to the inlet turbulence level and the superiority still holds even at some supercritical conditions.

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13

Xu, Xin, Da Wei Liu, De Hua Chen, and Yuan Jing Wang. "Influencing Factors Analysis of the Shock-Induced Separation for Supercritical Airfoil." Applied Mechanics and Materials 444-445 (October 2013): 221–26. http://dx.doi.org/10.4028/www.scientific.net/amm.444-445.221.

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The shock-induced separation easily occurred on the upper surface of supercritical airfoil at transonic speeds, which would change the aerodynamic characteristics. The problem of the shock-induced separation was not solved completely for the complicated phenomena and flow mechanism. In this paper, the influencing factors of shock-induced separation for supercritical airfoil CH was analyzed at transonic speeds. The Navier-Stokes equations were solved, in order to investigate influence of different attack angles, Mach numbers and Reynolds numbers. The computation attack angles of CH airfoil varied from 0oto 7o, Reynolds numbers varied from 5×106to 50×106per airfoil chord while Mach number varied from 0.74 to 0.82. It was shown that the shock-induced separation was affected by attack angles, Mach numbers and Reynolds numbers, but the influence tendency and areas were quite different. The shock wave location and intensity were affected by the three factors, and the boundary layer thickness was mainly affected by Reynolds number, while the separation structure was mainly determined by the attack angle and Mach number.

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14

Alshabu, A., and H. Olivier. "Unsteady Wave Phenomena on a Supercritical Airfoil." AIAA Journal 46, no. 8 (August 2008): 2066–73. http://dx.doi.org/10.2514/1.35516.

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15

Butefisch, Karl A., and Egon Stanewsky. "Experimental flowfield study on a supercritical airfoil." Journal of Aircraft 24, no. 11 (November 1987): 783–88. http://dx.doi.org/10.2514/3.45521.

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16

PAMBAGJO, Tjoetjoek Eko, Kazuhiro NAKAHASHI, and Shigeru OBAYASHI. "Inverse Design of a Thick Supercritical Airfoil." TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 43, no. 140 (2000): 61–66. http://dx.doi.org/10.2322/tjsass.43.61.

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17

Liu, Da Wei, Xin Peng, Xin Xu, and De Hua Chen. "Investigation on the Multi-Objective Optimization of Supercritical Airfoil Based on Nondominated Sorting Genetic Algorithm." Applied Mechanics and Materials 444-445 (October 2013): 357–62. http://dx.doi.org/10.4028/www.scientific.net/amm.444-445.357.

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This paper aimed to investigate the multi-objective optimization method of supercritical airfoil. To achieve the optimal design of supercritical airfoil Rae2822, an improved NSGA-2 (Nondominated Sorting Genetic Algorithm) method was utilized, while the cross-operator and adaptive-variation operator were introduced to improve the convergence speed of the algorithm. During the optimization, the airfoil parametric modeling was achieved based on the Bezier-Bernstein method, and the objective function was obtained through solving the N-S equations. Considering the parallel computation characteristics of the algorithm, the computation was conducted in large-scale Linux computer system to reduce the solving time. Optimization results showed that the undominate solution with high quality obtained through the NSGA-2 method distributed evenly, which provided the designer a wider choosing space. It was also showed that the multi-objective optimization method presented in this paper was feasible and reliable.

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18

Zika, V. J. "Correlation and Prediction of Rotating Stall Inception by Divergence Method." Journal of Fluids Engineering 107, no. 2 (June 1, 1985): 191–96. http://dx.doi.org/10.1115/1.3242459.

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An empirical correlation of rotating stall inception points of elementary compressors (isolated rotors, stages without prerotation, complete single stages, and multi-stage machines with repeating stages), modeled as equivalent diffusers, is presented. From it, two inception criteria for self-induced rotating stall are derived. Compressor blade rows are classified according to a geometric form parameter, (L/A∞)cor, into two groups, subcritical and supercritical. The subcritical geometries stall at a constant kinematic area ratio AE/A∞, in what appears to be a pure rotating stall mode, which occurs before the airfoil stalls. In supercritical geometries, the rotating stall is delayed until it is triggered by the airfoil stall. Thus, for the latter geometries, the airfoil stall and rotating stall are coincident. In contrast to other diffuser-analog methods, the divergence method determines the stall angle and the stalled flow coefficient rather than the stalled pressure rise.

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19

Niederdrenk, P., H. Sobieczky, and G. S. Dulikravich. "Supercritical Cascade Flow Analysis With Shock-Boundary Layer Interaction and Shock-Free Redesign." Journal of Turbomachinery 109, no. 3 (July 1, 1987): 413–19. http://dx.doi.org/10.1115/1.3262121.

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This paper describes improvements made in a user-oriented analysis code for steady two-dimensional transonic flows in turbomachinery cascades. The full potential equation is solved by a finite area technique, using a C-type grid and an analytical wake model. Solution adaptive grid clustering refines the inviscid computationally captured shock. The boundary layer is computed by an integral method except in the shock region, where the analytical interaction model of Bohning and Zierep smoothes out the pressure distribution on the airfoil surface. The code is applied in its analysis and design modes to an experimentally tested cascade of airfoils.

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20

Wang, Yuanjing, Binbin Lv, Pengxuan Lei, Wenkui Shi, and Yu Yan. "Study on Flow Mechanism of a Morphing Supercritical Airfoil." Shock and Vibration 2021 (April 19, 2021): 1–11. http://dx.doi.org/10.1155/2021/5588056.

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In order to maintain the best performance in flight, a new concept, morphing aircraft, has been proposed, which can change the real-time aerodynamic characteristics under different flight conditions. The key problem is to figure out the response of strong flow instability caused by structure changes during the morphing. To solve this problem, computational fluid dynamics (CFD) and wind tunnel tests (WTT) were employed. The results show that the deformation of thickness and camber angle of the airfoil will significantly change the distribution of pressure and result in obvious hysteresis loops of lift and drag. With the increase of deformation frequency and amplitude, the instability increases correspondingly. Moreover, the unsteady effect caused by camber deformation is much stronger than that caused by thickness deformation. In addition, the flow structures on the airfoil, such as the shock strength and boundary separation location, have a delay in response to structure changes. Therefore, there will be a hysteresis between airfoil deformation and aerodynamic characteristics, which means strong flow instability.

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21

Akbar, Hassan, Yu Qin Jiao, and Abu Bakar. "Flow Control Using Air-Jet to Improve the Aerodynamic Performance of a Multi-Element Airfoil." Applied Mechanics and Materials 772 (July 2015): 441–45. http://dx.doi.org/10.4028/www.scientific.net/amm.772.441.

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This paper describes the application of active flow control for the NLR7301 supercritical airfoil/flap configuration at Re = 2.51x106. A parametric analysis is conducted to investigate the effects of jet parameters (jet direction, jet location and momentum coefficient) on the aerodynamic performance of a multi-element airfoil. The results indicate that flow separation is delayed and efficiency of jet can be improved with specific momentum coefficient (the best lift-drag ratio at Cμ=0.16) and jet angle (16°) when the jet is located near the separation point of the airfoil.

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22

Xu, Zhaoyi, Joseph H. Saleh, and Vigor Yang. "Optimization of Supercritical Airfoil Design with Buffet Effect." AIAA Journal 57, no. 10 (October 2019): 4343–53. http://dx.doi.org/10.2514/1.j057573.

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23

Lee, B. H. K. "Investigation of flow separation on a supercritical airfoil." Journal of Aircraft 26, no. 11 (November 1989): 1032–37. http://dx.doi.org/10.2514/3.45876.

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24

Tian, Yun, PeiQing Liu, and PeiHua Feng. "Shock control bump parametric research on supercritical airfoil." Science China Technological Sciences 54, no. 11 (September 24, 2011): 2935–44. http://dx.doi.org/10.1007/s11431-011-4582-y.

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25

Ramaswamy, M. A. "Characteristics of a typical lifting symmetric supercritical airfoil." Sadhana 10, no. 3-4 (August 1987): 445–58. http://dx.doi.org/10.1007/bf02811306.

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26

Брутян, М. А., А. В. Волков, and А. В. Потапчик. "Экспериментальное исследование пассивного способа ослабления трансзвукового баффета." Письма в журнал технической физики 45, no. 21 (2019): 19. http://dx.doi.org/10.21883/pjtf.2019.21.48467.17881.

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Experimental studies of buffeting onset phenomenon physical picture on a model of a supercritical airfoil at transonic speeds are presented. Passive control method of this phenomenon based on application of the special jet vortex generators is given. It is shown that application of this method leads to delaying of the onset and decreasing of buffeting phenomenon, and at the result – to improve aerodynamic characteristics of airfoil.

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27

Fang, Ming Xia. "Aeroelastic Analysis of Supersonic Airfoil with Hysteretic Nonlinearities." Applied Mechanics and Materials 141 (November 2011): 180–85. http://dx.doi.org/10.4028/www.scientific.net/amm.141.180.

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Considering nonlinear aerodynamic forces and airfoil structure nonlinearity, aeroelasticity of supersonic airfoil was researched. Using the bifurcation diagram, phase diagram and Poincare map，effect of airfoil aeroelasticity by nonlinear aerodynamic forces and parameters of structural hysteresis was analyzed. Research shows that only consider nonlinear aerodynamic force, systematic motion tends to periodic and quasi- periodic LCO movement, and the vibration amplitude of plunge and pitch movement leap with Ma increasing .When taking into account the nonlinear aerodynamic force and structural nonlinearity, with the change of damping factor, the system will appear subcritical and supercritical flutter, and for the influence of nonlinear structure, movement leap of vibration amplitude does not occur. This means the choice of reasonable structure parameters will help to improve the aeroelastic characteristics of airfoil.

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28

Masdari, Mehran, Maryam Ghorbani, and Arshia Tabrizian. "Experimental study of wake steadiness of an airfoil in pitch–hold–return motion." Aircraft Engineering and Aerospace Technology 92, no. 7 (June 4, 2020): 1019–25. http://dx.doi.org/10.1108/aeat-07-2019-0154.

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Purpose The purpose of this paper is to analyze experimentally subsonic wake of a supercritical airfoil undergoing a pitch–hold–return motion. The focus of the investigation has been narrowed to concentrate on the steadiness of the flow field in the wake of the airfoil and the role of reduced frequency, amplitude and the hold phase duration. Design/methodology/approach All experiments were conducted in a low sub-sonic closed-circuit wind tunnel, at a Reynolds number of approximately 600,000. The model was a supercritical airfoil having 10% thickness and wall-to-wall in ground test facilities. To calculate the velocity distribution in the wake of the airfoil, total and static pressures were recorded at a distance of one chord far from the trailing edge, using pressure devices. The reduced frequency was set at 0.012, 0.03 and the motion pivot was selected at c/4. Findings Analysis of the steadiness of the wake flow field ascertains that an increase in reduced frequency leads to further flow time lag in the hold phase whereas decreases the time that the wake remains steady after the start of the return portion. Also, the roles of amplitude and stall condition are examined. Practical implications Examination of a pitch–hold–return motion is substantial in assessment of aerodynamics of maneuvers with a rapid increase in angle of attack. Moreover, study of aerodynamic behavior of downstream flow field and its steadiness in the wake of the airfoil is vital in drag reduction and control of flapping wings, dynamic stability and control of aircrafts. Originality/value In the present study, to discuss the steadiness of the flow field behind the airfoil some statistical methods and concept of histogram using an automatic algorithm were used and a specific criterion to characterize the steadiness of flow field was achieved.

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29

Rokoni, Arif Abdullah, and A. B. M. Toufique Hasan. "PREDICTION OF TRANSONIC BUFFET ONSET FOR FLOW OVER A SUPERCRITICAL AIRFOIL- A NUMERICAL INVESTIGATION." Journal of Mechanical Engineering 43, no. 1 (July 23, 2013): 48–53. http://dx.doi.org/10.3329/jme.v43i1.15782.

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Transonic flow over a supercritical airfoil leads to the appearances of unsteady shock waves in theflow field. At certain flow conditions, the interaction of unsteady shock waves with boundary layer becomescomplex and generates self-excited shock oscillation, lift fluctuation and thus initiate the buffet. In the presentstudy, Reynolds averaged Navier-Stokes equations with k-? SST turbulence model has been applied to predictthe shock induced buffet onset for the flow over a supercritical airfoil NASA SC(2) 0714. The free streamtransonic Mach number is kept in the range of 0.71 to 0.75 while the angle of attack is varied in a wide range.The onset of buffet is confirmed by the fluctuating aerodynamic properties such as lift-coefficient, pressurecoefficient, static pressure and so on. The self-excited shock oscillation and the corresponding buffet frequencyare numerically analyzed.DOI: http://dx.doi.org/10.3329/jme.v43i1.15782

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30

Deck, Sébastien. "Numerical Simulation of Transonic Buffet over a Supercritical Airfoil." AIAA Journal 43, no. 7 (July 2005): 1556–66. http://dx.doi.org/10.2514/1.9885.

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31

Li, Haoran, Yufei Zhang, and Haixin Chen. "Optimization of Supercritical Airfoil Considering the Ice-Accretion Effects." AIAA Journal 57, no. 11 (November 2019): 4650–69. http://dx.doi.org/10.2514/1.j057958.

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32

Inger, G. R. "Application of Oswatitsch's theorem to supercritical airfoil drag calculation." Journal of Aircraft 30, no. 3 (May 1993): 415–16. http://dx.doi.org/10.2514/3.46354.

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33

Weber, Stefan, Kevin D. Jones, John A. Ekaterinaris, and Max F. Platzer. "Transonic flutter computations for the NLR 7301 supercritical airfoil." Aerospace Science and Technology 5, no. 4 (June 2001): 293–304. http://dx.doi.org/10.1016/s1270-9638(01)01099-9.

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34

Hasan, A. B. M. Toufique, and Md Mahbub Alam. "RANS Computation of Transonic Buffet over a Supercritical Airfoil." Procedia Engineering 56 (2013): 303–9. http://dx.doi.org/10.1016/j.proeng.2013.03.123.

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35

Boroumand, Behnaz Beheshti, and Mahmoud Mani. "Wake measurements of oscillating supercritical airfoil in compressible flow." Transactions of the Canadian Society for Mechanical Engineering 43, no. 1 (March 1, 2019): 112–21. http://dx.doi.org/10.1139/tcsme-2017-0022.

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Boundary layer and wake behaviors are strongly affected by airfoil motion. Moreover, parameters like body oscillation frequency, oscillation type, Mach number, and angle of attack play main roles in wake characteristics. In this research, both static and dynamic tests were carried out in a tri-sonic wind tunnel to study wake profiles experimentally by hot wire anemometry. All data were recorded at a free stream Mach number of 0.4. Quarter-length and half-length of chord were also considered as downstream distances from the trailing edge in pitching motions of mean angle of attack of −0.4°. Frequencies of 3 Hz and 6 Hz with amplitude of 3° were chosen as oscillation parameters. Voltages at hot wire outputs were measured and analyzed qualitatively and statistically with root-mean-square, correlation, mean value distribution, time history, and frequency. Flow parameters were obtained by computational studies under similar experimental test conditions. The wake characteristics obtained from numerical and experimental methods were compared.

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36

Xiao, Q., H. M. Tsai, and F. Liu. "Numerical Study of Transonic Buffet on a Supercritical Airfoil." AIAA Journal 44, no. 3 (March 2006): 620–28. http://dx.doi.org/10.2514/1.16658.

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37

Wang, Na, and Chao Gao. "Variable Reynolds Number Experimental Study on Aerodynamic Characteristic of Supercritical Airfoil RAE2822." Applied Mechanics and Materials 420 (September 2013): 42–46. http://dx.doi.org/10.4028/www.scientific.net/amm.420.42.

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An experimental study of pressure distributions over RAE2822 airfoil in the two-dimensional test section 0.8×0.4 meter of a transonic wind tunnel which is the first pressruized continuous wind tunnel in China is presented. This paper in order to further study the influence of the dynamic of continuous changes Reynolds number at Mach number is 0.66 and 0.80, and the attack angle is from-2 degree to 10 degree, and especially the Reynolds number range from3.0×106to 12×106. The study is focalized on the subsonic range of flow conditions with separation and shock wave in the boundary layer. The influence of pressure distribution and pressure coefficient and moment coefficient caused by Reynolds number increasing are analyzed and discussed. The conclusions showed that the pressure distribution of the lower surface of the airfoil get the influence of the Reynolds number is negligible. The Reynolds number impact on the pressure distribution is faintness at Ma=0.66. Reynolds number increases affect the airfoil central and trailing edge pressure. As the Reynolds number increases, the CL curve move and the gradient increasing. The moment coefficient decreased as the Reynolds number increasing. The CL curve with Cd curve moves left as Reynolds number increasing.

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38

Брутян, М. А., А. В. Волков, and А. В. Потапчик. "Экспериментальное исследование нового способа уменьшения волнового сопротивления профиля при трансзвуковых скоростях." Письма в журнал технической физики 46, no. 12 (2020): 34. http://dx.doi.org/10.21883/pjtf.2020.12.49525.18054.

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A new method is proposed for wave drag reducing of a supercritical airfoil at transonic speeds, which is associated with the organization of a microwave section in a local supersonic zone on the upper surface. Simultaneously with optical studies, aerodynamic loads acting on a model were obtained. It is experimentally established that the proposed approach leads to the formation of a system of weak compression waves, to consecutive deceleration of the supersonic flow, to decrease of the Mach number in front of the shock wave, and, as a consequence, to weakening of its intensity and to reducing of the airfoil wave drag

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39

King, L. S., and D. A. Johnson. "Comparison of supercritical airfoil flow calculations with wind tunnel results." AIAA Journal 23, no. 9 (September 1985): 1301–7. http://dx.doi.org/10.2514/3.9085.

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40

Biber, Kasim, and Carl P. Tilmann. "Supercritical Airfoil Design for Future High-Altitude Long-Endurance Concepts." Journal of Aircraft 41, no. 1 (January 2004): 156–64. http://dx.doi.org/10.2514/1.1049.

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41

Rahman, Muhammad Rizwanur, Mohammad Itmam Labib, Abul Bashar Mohammad Toufique Hasan, Mohammad Saddam Hossain Joy, Toshiaki Setoguchi, and Heuy Dong Kim. "Control of Transonic Shock Wave Oscillation over a Supercritical Airfoil." Open Journal of Fluid Dynamics 05, no. 04 (2015): 302–10. http://dx.doi.org/10.4236/ojfd.2015.54031.

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42

Lee, B. H. K., and F. C. Tang. "Transonic buffet of a supercritical airfoil with trailing-edge flap." Journal of Aircraft 26, no. 5 (May 1989): 459–64. http://dx.doi.org/10.2514/3.45785.

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43

Yu, T., J. J. Wang, and P. F. Zhang. "Numerical Simulation of Gurney Flap on RAE-2822 Supercritical Airfoil." Journal of Aircraft 48, no. 5 (September 2011): 1565–75. http://dx.doi.org/10.2514/1.c031285.

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44

Zhou, Qiang, Jian Xiong, Liusheng Chen, Husheng Ma, and Yang Tao. "Surface Pressure Measurements on Supercritical Airfoil Employing Pressure -Sensitive Paint." Procedia Engineering 31 (2012): 1160–67. http://dx.doi.org/10.1016/j.proeng.2012.01.1157.

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45

Liu, Dawei, Yuanjing Wang, Dehua Chen, Xin Peng, and Xing Xu. "Numerical Investigation on the Reynolds Number Effects of Supercritical Airfoil." Procedia Engineering 31 (2012): 103–9. http://dx.doi.org/10.1016/j.proeng.2012.01.998.

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46

Alshabu, Atef, Herbert Olivier, and Igor Klioutchnikov. "Investigation of upstream moving pressure waves on a supercritical airfoil." Aerospace Science and Technology 10, no. 6 (September 2006): 465–73. http://dx.doi.org/10.1016/j.ast.2006.04.003.

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47

Zhao, Tong, Yufei Zhang, Haixin Chen, Yingchun Chen, and Miao Zhang. "Supercritical wing design based on airfoil optimization and 2.75D transformation." Aerospace Science and Technology 56 (September 2016): 168–82. http://dx.doi.org/10.1016/j.ast.2016.07.010.

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48

Anders, J. B., W. K. Anderson, and A. V. Murthy. "Transonic Similarity Theory Applied to a Supercritical Airfoil in Heavy Gas." Journal of Aircraft 36, no. 6 (November 1999): 957–64. http://dx.doi.org/10.2514/2.2557.

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49

Xu, Xin, Da-wei Liu, De-hua Chen, Zhi Wei, and Yuan-jing Wang. "Investigation on Improved Correlation of CFD and EFD for Supercritical Airfoil." Research Journal of Applied Sciences, Engineering and Technology 7, no. 5 (February 5, 2014): 1007–11. http://dx.doi.org/10.19026/rjaset.7.350.

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50

Kuya, Yuichi, Kenshiro Boda, and Keisuke Sawada. "Numerical Study of Transonic Shock Buffet Control over a Supercritical Airfoil." Journal of Aircraft 57, no. 6 (November 2020): 1242–51. http://dx.doi.org/10.2514/1.c035902.

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