Relevant bibliographies by topics / Supercritical Airfoil

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  1. Journal articles
  2. Dissertations / Theses
  3. Books
  4. Book chapters
  5. Conference papers

Academic literature on the topic 'Supercritical Airfoil'

Author: Grafiati
Published: 4 June 2021
Last updated: 1 February 2022

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Journal articles on the topic "Supercritical Airfoil"

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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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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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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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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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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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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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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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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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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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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Dissertations / Theses on the topic "Supercritical Airfoil"

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Li, Daxin. "Multi-objective design optimization for high-lift aircraft configurations supported by surrogate modeling." Thesis, Cranfield University, 2013. http://dspace.lib.cranfield.ac.uk/handle/1826/8468.

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Nowadays, the competition among airlines seriously depend upon the saving operating costs, with the premise that not to degrade its services quality. Especially in the face of increasingly scarce oil resources, reducing fleets operational fuel consumption, is an important means to improve profits. Aircraft fuel economy is determined by operational management strategies and application technologies. The application of technologies mainly refers to airplane’s engine performance, Weight efficiency and aerodynamic characteristics. A market competitive aircraft should thoroughly consider to all of these aspects. Transport aircraft aerodynamic performance mainly is determined by wing’s properties. Wings that are optimized for efficient flight in cruise conditions need to be fitted with powerful high-lift devices to meet lift requirements for safe takeoff and landing. These high-lift devices have a significant impact on the total airplane performance. The aerodynamic characteristics of the wing airfoil will have a direct impact on the aerodynamic characteristics of the wing, and the wing’s effective cruise hand high-lift configuration design has a significant impact on the performance of transport aircraft. Therefore, optimizing the design is a necessary airfoil design process. Nowadays engineering analysis relies heavily on computer-based solution algorithms to investigate the performance of an engineering system. Computational fluid dynamics (CFD) is one of the computer-based solution methods which are more widely employed in aerospace engineering. The computational power and time required to carry out the analysis increases as the fidelity of the analysis increases. Aerodynamic shape optimization has become a vital part of aircraft design in the recent years. Since the aerodynamic shape optimization (ASO) process with CFD solution algorithms requires a huge amount of computational power, there is always some reluctance among the aircraft researchers in employing the ASO approach at the initial stages of the aircraft design. In order to alleviate this problem, statistical approximation models are constructed for actual CFD algorithms. The fidelity of these approximation models are merely based on the fidelity of data used to construct these models. Hence it becomes indispensable to spend more computational power in order to convene more data which are further used for constructing the approximation models. The goal of this thesis is to present a design approach for assumed wing airfoils; it includes the design process, multi-objective design optimization based on surrogate modelling. The optimization design stared from a transonic single-element single-objective optimization design, and then high-lift configurations were two low-speed conditions of multi-objective optimization design, on this basis, further completed a variable camber airfoil at low speed to high-lift configuration to improve aerodynamic performance. Through this study, prove a surrogate based model could be used in the wing airfoil optimization design.
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Hillenherms, Cornelia [Verfasser]. "Experimental Investigation of a Supercritical Airfoil Oscillating in Pitch at Transonic Flow / Cornelia Hillenherms." Aachen : Shaker, 2003. http://d-nb.info/1170540740/34.

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3

Hussain, Mian M. "Time-Resolved Analysis of Circulation Control over Supercritical Airfoil using Digital Particle Image Velocimetry (DPIV)." Thesis, Virginia Tech, 2004. http://hdl.handle.net/10919/40538.

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Active pneumatic flow control methods as applied to aerospace applications have shown noteworthy improvements in lift compared to traditional means. The General Aviation Circulation Control (GACC) concept currently under investigation at NASAâ s Langley Research Center (LaRC) is an attempt at addressing some of the fundamental obstacles related to the successful development and implementation of such techniques. The primary focus of research in the field of high lift pneumatic devices is to investigate ways of obtaining significant improvements in the lift coefficient without resorting to moving surfaces. Though it has been demonstrated that the lift coefficient can be amplified in a variety of ways, the chosen method for the current work is via enhanced circulation stemming from a trailing edge Coanda jet. A secondary objective is to reduce the amount energy expenditure used in these pneumatic techniques by implementing time-variant flow. This paper describes experimental observations of the flow behavior at the trailing edge of a modified water tunnel based supercritical airfoil model that exploits both steady and pulsed Coanda driven circulation control. A total of 10 sets of data, excluding a baseline case of no Coanda jet, were sampled with five cases each for steady and pulsed flow, the latter at a reduced frequency, f+, of 1. Two cases of equal momentum coefficient but with varying forced frequencies were isolated for further study in an attempt to accurately compare the resultant flow dynamics of each method. All measurements were taken at a zero-lift angle of attack by means of a non-invasive time accurate flow visualization technique (DPIV). Vorticity behavior was investigated using Tecplot® and a MATLAB® program was developed to quantify the Strouhal Number of time-averaged velocity fluctuations moving aft of the Coanda surface for each case.
Master of Science
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Song, Bo. "Experimental and Numerical Investigations of Optimized High-Turning Supercritical Compressor Blades." Diss., Virginia Tech, 2003. http://hdl.handle.net/10919/29727.

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Cascade testing and flow analysis of three high-turning supercritical compressor blades were conducted. The blades were designed at an inlet Mach number (M1) of 0.87 and inlet flow angle of 48.4 deg, with high camber angles of about 55 deg. The baseline blade was a conventional Controlled Diffusion Airfoil (CDA) design and the other two were optimized blades. The blades were tested for an inlet Mach number range from 0.61 to 0.95 and an inlet flow angle range from 44.4 deg to 50.4 deg, at high Reynolds numbers (1.2-1.9x10^6 based on the blade chord). The test results have shown lower losses and better incidence robustness for the optimized blades at higher supercritical flow conditions (M1>0.83). At the design condition, 30% loss reduction was achieved. The blade-to-blade flow was computed by solving the two-dimensional steady Navier-Stokes equations. Experimental results, in conjunction with the CFD flowfield characterization, revealed the loss reduction mechanism: severe boundary layer separation occurred on the suction surface of the baseline blade while no separation occurred for the optimized blades. Furthermore, whether the boundary layer was separated or not was found due to different shock patterns, different shock-boundary layer interactions and different pressure distributions on the blades. For the baseline blade, the strong passage shock coincided with the adverse pressure gradient due to the high blade front camber at 20% chord, leading to the flow separation. For the optimized blades, the high blade camber shifted to more downstream (30-40% chord), resulting in stronger flow leading edge acceleration, less strength of the passage shock near the blade surface, favorable pressure gradient right after the passage shock, thus no flow separation occurred. The flow understanding obtained by the current research can be used to guide the design of high-turning compressor blades at higher supercritical flow conditions.
Ph. D.
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Merchant, Ali A. (Ali Abbas). "Design and analysis of supercritical airfoils with boundary layer suction." Thesis, Massachusetts Institute of Technology, 1996. http://hdl.handle.net/1721.1/10987.

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Loo, Felipe Manuel. "Numerical Study of Limit Cycle Oscillation Using Conventional and Supercritical Airfoils." Scholarly Repository, 2008. http://scholarlyrepository.miami.edu/oa_theses/176.

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Limit Cycle Oscillation is a type of aircraft wing structural vibration caused by the non-linearity of the system. The objective of this thesis is to provide a numerical study of this aeroelastic behavior. A CFD solver is used to simulate airfoils displaying such an aeroelastic behavior under certain airflow conditions. Two types of airfoils are used for this numerical study, including the NACA64a010 airfoil, and the supercritical NLR 7301 airfoil. The CFD simulation of limit cycle oscillation (LCO) can be obtained by using published flow and structural parameters. Final results from the CFD solver capture LCO, as well as flutter, behaviors for both wings. These CFD results can be obtained by using two different solution schemes, including the Roe and Zha scheme. The pressure coefficient and skin friction coefficient distributions are computed using the CFD results for LCO and flutter simulations of these two airfoils, and they provide a physical understanding of these aeroelastic behaviors.
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Gulla, Duncan. "Ausgewählte statistische Betrachtungen im Flugzeugentwurf: Superkritische Profile und Fahrwerk." Aircraft Design and Systems Group (AERO), Department of Automotive and Aeronautical Engineering, Hamburg University of Applied Sciences, 2019. http://d-nb.info/1180601696.

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Kenntnisse über Parametereigenschaften und Charakteristiken von Flugzeugkomponenten sind eine wesentliche Grundlage für Methoden des Flugzeugentwurfs. Daher ist Ziel dieser Arbeit, statistische Merkmale und Kenngrößen einer für den Flugzeugbau und Entwurf relevanten Auswahl an Komponenten zu erschließen. Dabei wurden zunächst superkritische Tragflügelprofile hinsichtlich ihrer geometrischen Eigenschaften (relative Profildicke, Wölbung, Dickenrücklage, Wölbungsrücklage und der sogenannte "Leading Edge Sharpness Parameter") untersucht. Diese Eigenschaften wurden mit der Software XFLR5 aus einer Auswahl an superkritischen Profilgeometrien erhoben und mit grafischen und beschreibenden Statistikmethoden ausgewertet. Die Profile wiesen relative Wölbungen von 0 % bis 3,4 % auf, die Mehrzahl entfiel auf Wölbungen von 1 % bis 2 %. Die Wölbungsrücklagen zeigten die für superkritische Profile typische Lage im hinteren Profilbereich zwischen 70 % und 90 % der Profiltiefe. Die Dickenrücklagen verteilten sich um einen Mittelwert von 37 % der Profiltiefe. Eine Betrachtung von Flugzeugreifendimensionen sollte das Verhältnis von Reifenbreite zum Durchmesser w/d charakterisieren. Es wurde ein annähernd lineares Verhalten festgestellt. Die Werte des Parameters w/d umfassten einen Bereich von 0,3 bis 0,4. Durch Regressionsanalysen konnten auch die Abhängigkeiten des Parameters w/d von nur einer bekannten Reifendimension (Breite oder Durchmesser) aufgezeigt werden. Die im Rahmen dieser Arbeit dargestellten Erkenntnisse können als Grundlage weiterführender Untersuchungen genutzt werden.
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Books on the topic "Supercritical Airfoil"

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Maurice, Holt. Supercritical flow past a symmetrical bicircular arc airfoil. Berkeley, CA: Dept. of Mechanical Engineering, University of California, 1989.

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Mokry, M. Influence of the transonic doublet in the farfield of a lifting airfoil. Ottawa: National Research Council Canada, Institute for Aerospace Research, 1993.

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Jenkins, Renaldo. Aerodynamic performance and pressure distributions for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel. Hampton, Va: Langley Research Center, 1988.

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Chan, Y. Y. Comparison of boundary layer trips of disk and grit types on airfoil performance at transonic speeds. Ottawa: National Aeronautical Establishment, National Research Council Canada, 1988.

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Jenkins, Renaldo V. NASA SC(2)-0714 airfoil data corrected for sidewall boundary-layer effects in the Langley 0.3-Meter Transonic Cryogenic Tunnel. Hampton, Va: Langley Research Center, 1989.

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Khalid, M. Further studies on the 21% thick, supercritical NLR airfoil NAE 68-060-21:1. Ottawa: National Aeronautical Establishment, 1986.

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Harris, Charles D. The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil: Design overview. Hampton, Va: Langley Research Center, 1988.

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Lee, B. H. K. An experimental study of transonic buffet of a supercritical airfoil with trailing edge flap. Ottawa: National Research Council Canada, 1988.

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Brooks, Cuyler W. The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil: Drag equations. Hampton, Va: Langley Research Center, 1989.

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Chan, Y. Y. Analysis of experimental data for cast 10-2/DOA2 supercritical airfoil at high Reynolds numbers. Ottawa: National Research Council of Canada, 1988.

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Book chapters on the topic "Supercritical Airfoil"

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Alshabu, A., H. Olivier, V. Herms, and I. Klioutchnikov. "Wave processes on a supercritical airfoil." In Shock Waves, 1321–26. Berlin, Heidelberg: Springer Berlin Heidelberg, 2009. http://dx.doi.org/10.1007/978-3-540-85181-3_85.

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Klutchnikov, Igor, and Josef Ballmann. "DNS of Transonic Flow about a Supercritical Airfoil." In Direct and Large-Eddy Simulation V, 223–30. Dordrecht: Springer Netherlands, 2004. http://dx.doi.org/10.1007/978-1-4020-2313-2_24.

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Lv, Binbin, Pengxuan Lei, Yuanjing Wang, and Wenkui Shi. "Research on the Aerodynamic Characteristics of Morphing Supercritical Airfoil." In Lecture Notes in Electrical Engineering, 828–35. Singapore: Springer Singapore, 2019. http://dx.doi.org/10.1007/978-981-13-3305-7_65.

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Garnier, E., and S. Deck. "Large-Eddy Simulation of Transonic Buffet over a Supercritical Airfoil." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 135–41. Berlin, Heidelberg: Springer Berlin Heidelberg, 2010. http://dx.doi.org/10.1007/978-3-642-14139-3_16.

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Garnier, E., and S. Deck. "Large-Eddy Simulation of Transonic Buffet over a Supercritical Airfoil." In Direct and Large-Eddy Simulation VII, 549–54. Dordrecht: Springer Netherlands, 2010. http://dx.doi.org/10.1007/978-90-481-3652-0_81.

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Tang, Lei, D. D. Liu, and P. C. Chen. "Nonlinear Aerodynamic Effects on Transonic LCO Amplitude of a Supercritical Airfoil." In IUTAM Symposium Transsonicum IV, 53–58. Dordrecht: Springer Netherlands, 2003. http://dx.doi.org/10.1007/978-94-010-0017-8_9.

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van Rooij, A. C. L. M., and W. Wegner. "Numerical Investigation of the Flutter Behaviour of a Laminar Supercritical Airfoil." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 33–41. Cham: Springer International Publishing, 2014. http://dx.doi.org/10.1007/978-3-319-03158-3_4.

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Banavara, Nagaraj K., and Diliana Dimitrov. "Prediction of Transonic Flutter Behavior of a Supercritical Airfoil Using Reduced Order Methods." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 365–73. Cham: Springer International Publishing, 2014. http://dx.doi.org/10.1007/978-3-319-03158-3_37.

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Grossi, F., M. Braza, and Y. Hoarau. "Delayed Detached-Eddy Simulation of the Transonic Flow around a Supercritical Airfoil in the Buffet Regime." In Progress in Hybrid RANS-LES Modelling, 369–78. Berlin, Heidelberg: Springer Berlin Heidelberg, 2012. http://dx.doi.org/10.1007/978-3-642-31818-4_32.

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Tô, J. B., D. M. Zilli, N. Simiriotis, I. Asproulias, D. Szubert, A. Marouf, Y. Hoarau, and M. Braza. "Numerical Simulation and Modelling of a Morphing Supercritical Airfoil in a Transonic Flow at High Reynolds Numbers." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 371–84. Cham: Springer International Publishing, 2021. http://dx.doi.org/10.1007/978-3-030-55594-8_31.

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Conference papers on the topic "Supercritical Airfoil"

1

Vaishak, V. N., M. Soumith Reddy, and S. Shali. "Aerodynamic analysis of supercritical airfoil." In SEVENTH INTERNATIONAL SYMPOSIUM ON NEGATIVE IONS, BEAMS AND SOURCES (NIBS 2020). AIP Publishing, 2021. http://dx.doi.org/10.1063/5.0057918.

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2

Choudhury, Mushrif, Jie Cui, and Vahid Motevalli. "Effects of Supercritical Airfoil Upper Camber Modification on Overall Airfoil Performance." In The 5th World Congress on Momentum, Heat and Mass Transfer. Avestia Publishing, 2020. http://dx.doi.org/10.11159/enfht20.02.

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3

Biber, Kasim, and Carl Tilmann. "Supercritical Airfoil Design for Future HALE Concepts." In 41st Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2003. http://dx.doi.org/10.2514/6.2003-1095.

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4

Haderlie, Jacob, and William Crossley. "A Parametric Approach to Supercritical Airfoil Design Optimization." In 9th AIAA Aviation Technology, Integration, and Operations Conference (ATIO). Reston, Virigina: American Institute of Aeronautics and Astronautics, 2009. http://dx.doi.org/10.2514/6.2009-6950.

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5

Tartinville, Benoit, Virginie Barbieux, and Lionel Temmerman. "High Fidelity Gust Simulations Over a Supercritical Airfoil." In 2018 Applied Aerodynamics Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2018. http://dx.doi.org/10.2514/6.2018-3634.

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6

INGER, G. "Application of Oswaititsch's theorem to supercritical airfoil drag calculation." In 9th Applied Aerodynamics Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1991. http://dx.doi.org/10.2514/6.1991-3210.

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7

Dawei, Liu, Xu Xin, Wei Zhi, and Chen Dehua. "Investigation on the Reynolds Number Simulation of Supercritical Airfoil." In 2013 Fourth International Conference on Digital Manufacturing & Automation (ICDMA). IEEE, 2013. http://dx.doi.org/10.1109/icdma.2013.176.

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8

Sonoda, Toyotaka, and Heinz-Adolf Schreiber. "Aerodynamic Characteristics of Supercritical Outlet Guide Vanes at Low Reynolds Number Conditions." In ASME Turbo Expo 2006: Power for Land, Sea, and Air. ASMEDC, 2006. http://dx.doi.org/10.1115/gt2006-90882.

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Abstract:
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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9

Aley, Kade, Tufan K. Guha, and Rajan Kumar. "Experimental Characterization of a High-Lift Supercritical Airfoil with Microjets." In 2018 Flow Control Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2018. http://dx.doi.org/10.2514/6.2018-3688.

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10

Xiao, Q., H. M. Tsai, and F. Liu. "A Numerical Study of Transonic Buffet on a Supercritical Airfoil." In 42nd AIAA Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2004. http://dx.doi.org/10.2514/6.2004-1056.

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