scholarly journals Flow Past a Flat Plate With a Vortex/Sink Combination

1996 ◽  
Vol 63 (2) ◽  
pp. 543-550 ◽  
Author(s):  
N. J. Mourtos ◽  
M. Brooks
Keyword(s):  
Flat Plate ◽  
Current Model ◽  
Angle Of Attack ◽  
Separated Flow ◽  
Leading Edge ◽  
Classical Solutions ◽  
Two Dimensional ◽  
Trailing Edge ◽  
Kutta Condition ◽  
Attached Flow

This paper presents a potential flow model for the leading edge vortex over a two-dimensional flat plate at an angle of attack. The paper is an extension of a model by Saffman and Sheffield (1977). A sink has been added in this model in an effort to satisfy the Kutta condition at both the leading edge and the trailing edge of the plate. The introduction of the sink was inspired by the fact that most steady vortices in nature appear in combination with a flow feature which can be interpreted as a sink at their cores when the flow is analyzed in a two-dimensional observation plane. As in the Saffman and Sheffield model, the presence of a vortex results in increased lift; however, in the current model a unique vortex/sink position is found at each angle of attack. A comparison has also been made between the lift and the drag of this model and the corresponding results for two classical solutions of flow over a flat plate: (a) the fully attached flow with the Kutta condition satisfied at the trailing edge only and (b) the Helmholtz solution of fully separated flow.

A theoretical investigation of enhanced lift in the presence of thin aerofoil stall

1999 ◽  
Vol 103 (1023) ◽  
pp. 237-244 ◽  
Author(s):  
W. W. H. Yeung ◽  
G. V. Parkinson
Keyword(s):  
Flat Plate ◽  
Flow Model ◽  
Separated Flow ◽  
Leading Edge ◽  
Geometrical Parameters ◽  
Separation Flow ◽  
Attached Flow ◽  
Positive Angle ◽  
Edge Separation ◽  
Potential Flow Model

Abstract A theoretical study is presented for the investigation of a potential-flow model for enhancing lift over a flat-plate aerofoil experiencing thin aerofoil stall. Rather than suppressing the leading-edge separation, flow is assumed to separate tangentially at the leading edge and made to reattach smoothly at the tip of a forward-facing fence joining the plate tangentially on its upper surface to avoid any unnecessary stagnated flow. The length of the fence and its location from the leading edge form two geometrical parameters. At any positive angle of attack, the resulting bounding streamline emanating from the leading edge and terminating at the tip of the fence is simulated by using suitable mathematical singularities subject to boundary conditions such as attaining a finite velocity at each critical point of the conformal mapping involved, and the condition of finite pressure gradient at reattachment, when applicable. Computational results from varying these two geometrical parameters indicate that the lift from each model is enhanced, as compared with the attached flow model around a simple flat plate and the original separated flow model by Kirchhoff.

Calculation of Attached or Partially Separated Flow Around Airfoil Sections

1977 ◽  
Vol 21 (02) ◽  
pp. 69-81
Author(s):  
Jerome H. Milgram
Keyword(s):  
Separated Flow ◽  
Leading Edge ◽  
Perturbation Series ◽  
Unified Theory ◽  
Physical Plane ◽  
Trailing Edge ◽  
Separation Flows ◽  
Attached Flow ◽  
Approach Zero ◽  
Edge Separation

Two-dimensional airfoil theory is developed for incompressible flow, for which most aspects of the solution are found in a "mapped plane." Although the airfoil contour has large slopes in the physical plane when the airfoil has a round leading edge, the slopes in the mapped plane are small and uniformly approach zero as the airfoil thickness and camber ratios are made to approach zero by an affine transformation on the airfoil shape. Because of the small slopes in the mapped plane, a perturbation series is appropriate there. A relatively unified theory is given for attached and partially separated (trailing-edge separation) flows. The results of calculations by this theory are compared with the results of other theories and with measurements. The attached-flow theory is shown to be especially appropriate for thin marine propeller sections with blunted leading edges. Some new data on the pressure distribution across the wake at the trailing edge of an airfoil in partially separated flow are presented.

Wall Effect on Separated Flow Around an Inclined Flat Plate at High Incidence

2013 ◽  
Author(s):  
Chuanjin Lan ◽  
Laibing Jia ◽  
Zhen Li ◽  
Yanbao Ma
Keyword(s):  
Flat Plate ◽  
Practical Importance ◽  
Angle Of Attack ◽  
Flow Characteristics ◽  
Separated Flow ◽  
Bluff Bodies ◽  
Trailing Edge ◽  
Plate Length ◽  
Gap Ratio ◽  
Lift And Drag

The flow around bluff bodies has received a great deal of attention due to its practical importance in engineering and scientific relevance in fluid mechanics. The largely separated turbulent flow can be triggered by an inclined flat plate with sharp leading and trailing edge and the presence of wall can alternate the flow structures greatly. This paper applied the Unsteady Reynolds-averaged Navier-Stokes (URANS) model to simulate the vortex shedding phenomenon over the inclined flat plate for a Reynolds number Re = 20 000 with different angle of attack α and gap ratio D/L, the ratio of the distance from the plate to the wall (D) to the plate length (L). Vortex generation from the leading and trailing edge was captured clearly and the transportation and development of vortex structures were shown for different parameters. The smaller the gap ratio is, the more the flow characteristics are affected by the wall presence. The integral quantities, such as lift and drag coefficients show different peak values for different angle of attack.

Three-Dimensional and Rotational Aerodynamics on the NREL Phase VI Wind Turbine Blade

2008 ◽  
Vol 130 (3) ◽  
Author(s):  
Alvaro Gonzalez ◽  
Xabier Munduate
Keyword(s):  
Wind Turbine ◽  
Turbine Blade ◽  
Three Dimensional ◽  
Separated Flow ◽  
Leading Edge ◽  
Wind Turbine Blade ◽  
Trailing Edge ◽  
Rotating Blade ◽  
Nrel Phase Vi ◽  
Edge Separation

This work undertakes an aerodynamic analysis over the parked and the rotating NREL Phase VI wind turbine blade. The experimental sequences from NASA Ames wind tunnel selected for this study respond to the parked blade and the rotating configuration, both for the upwind, two-bladed wind turbine operating at nonyawed conditions. The objective is to bring some light into the nature of the flow field and especially the type of stall behavior observed when 2D aerofoil steady measurements are compared to the parked blade and the latter to the rotating one. From averaged pressure coefficients together with their standard deviation values, trailing and leading edge separated flow regions have been found, with the limitations of the repeatability of the flow encountered on the blade. Results for the parked blade show the progressive delay from tip to root of the trailing edge separation process, with respect to the 2D profile, and also reveal a local region of leading edge separated flow or bubble at the inner, 30% and 47% of the blade. For the rotating blade, results at inboard 30% and 47% stations show a dramatic suppression of the trailing edge separation, and the development of a leading edge separation structure connected with the extra lift.

Part II: Steady aerofoil-spoiler characteristics

1987 ◽  
Vol 91 (908) ◽  
pp. 359-366
Keyword(s):  
Separated Flow ◽  
Experimental Results ◽  
Surface Singularity ◽  
Separation Region ◽  
Two Dimensional ◽  
Trailing Edge ◽  
Singularity Method ◽  
Total Head ◽  
Low Speeds ◽  
Theoretical Results

Summary A surface singularity method has been formulated to predict two-dimensional spoiler characteristics at low speeds. Vorticity singularities are placed on the aerofoil surface, on the spoiler surface, on the upper separation streamline from the spoiler tip and on the lower separation streamline from the aerofoil trailing edge. The separation region is closed downstream by two discrete vortices. The flow inside the separation region is assumed to have uniform total head. The downstream extent of the separated wake is an empirical input. The flows both external and internal to the separated regions are solved. Theoretical results have been obtained for a range of spoiler-aerofoil configurations which compare reasonably with experimental results. The model is deficient in that it predicts a higher compression ahead of the spoiler than obtained in practice. Furthermore, there is a minimum spoiler angle below which a solution is not possible; it is thought that this feature is related to the physical observation that at small spoiler angles, the separated flow from the spoiler reattaches on the aerofoil upper surface ahead of the trailing edge.

An experimental investigation of resonant interaction of a rectangular jet with a flat plate

2015 ◽  
Vol 779 ◽  
pp. 751-775 ◽  
Author(s):  
K. B. M. Q. Zaman ◽  
A. F. Fagan ◽  
J. E. Bridges ◽  
C. A. Brown
Keyword(s):  
Mach Number ◽  
Flat Plate ◽  
Cross Section ◽  
Leading Edge ◽  
Resonant Interaction ◽  
Trailing Edge ◽  
Resonant Condition ◽  
Direction Parallel ◽  
Rectangular Jet ◽  
Axis Switching

The interaction between an 8:1 aspect ratio rectangular jet and a flat plate, placed parallel to the jet, is addressed in this study. At high subsonic conditions and for certain relative locations of the plate, a resonance takes place with accompanying audible tones. Even when the tone is not audible the sound pressure level spectra are often marked by conspicuous peaks. The frequencies of these peaks, as functions of the plate’s length, its location relative to the jet as well as jet Mach number, are studied in an effort to understand the flow mechanism. It is demonstrated that the tones are not due to a simple feedback between the nozzle exit and the plate’s trailing edge; the leading edge also comes into play in determining the frequency. With parametric variation, it is found that there is an order in the most energetic spectral peaks; their frequencies cluster in distinct bands. The lowest frequency band is explained by an acoustic feedback involving diffraction at the plate’s leading edge. Under the resonant condition, a periodic flapping motion of the jet column is seen when viewed in a direction parallel to the plate. Phase-averaged Mach number data on a cross-stream plane near the plate’s trailing edge illustrate that the jet cross-section goes through large contortions within the period of the tone. Farther downstream a clear ‘axis switching’ takes place for the time-averaged cross-section of the jet that does not occur otherwise for a non-resonant condition.

Two-Dimensional Experimental Investigation on the Effects of a Fowler Flap Gap and Overlap Size on the Wake Flow Field

2016 ◽  
Vol 139 (2) ◽  
Author(s):  
David Demel ◽  
Mohsen Ferchichi ◽  
William D. E. Allan ◽  
Marouen Dghim
Keyword(s):  
Experimental Investigation ◽  
Flow Field ◽  
Angle Of Attack ◽  
Wake Flow ◽  
Two Dimensional ◽  
Edge Region ◽  
Suction Surface ◽  
Trailing Edge ◽  
Piv Measurements ◽  
Image Velocimetry

This work details an experimental investigation on the effects of the variation of flap gap and overlap sizes on the flow field in the wake of a wing-section equipped with a trailing edge Fowler flap. The airfoil was based on the NACA 0014-1.10 40/1.051 profile, and the flap was deployed with 40 deg deflection angle. Two-dimensional (2D) particle image velocimetry (PIV) measurements of the flow field in the vicinity of the main wing trailing edge and the flap region were performed for the optimal flap gap and overlap, as well as for flap gap and overlap increases of 2% and 4% chord beyond optimal, at angles of attack of 0 deg, 10 deg, and 12 deg. For all the configurations investigated, the flow over the flap was found to be fully stalled. At zero angle of attack, increasing the flap gap size was found to have minor effects on the flow field but increased flap overlap resulted in misalignment between the main wing boundary layer (BL) flow and the slot flow that forced the flow in the trailing edge region of the main wing to separate. When the angle of attack was increased to near stall conditions (at angle of attack of 12 deg), increasing the flap gap was found to energize and improve the flow in the trailing edge region of the main wing, whereas increased flap overlap further promoted flow separation on the main wing suction surface possibly steering the wing into stall.

scholarly journals Design and Validation of a New Morphing Camber System by Testing in the Price—Païdoussis Subsonic Wind Tunnel

Aerospace ◽  
2020 ◽  
Vol 7 (3) ◽  
pp. 23 ◽  
Author(s):  
David Communier ◽  
Ruxandra Mihaela Botez ◽  
Tony Wong
Keyword(s):  
Wind Tunnel ◽  
Unmanned Aerial Vehicle ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Trailing Edge ◽  
Subsonic Wind Tunnel ◽  
Aerial Vehicle ◽  
Stall Angle

This paper presents the design and wind tunnel testing of a morphing camber system and an estimation of performances on an unmanned aerial vehicle. The morphing camber system is a combination of two subsystems: the morphing trailing edge and the morphing leading edge. Results of the present study show that the aerodynamics effects of the two subsystems are combined, without interfering with each other on the wing. The morphing camber system acts only on the lift coefficient at a 0° angle of attack when morphing the trailing edge, and only on the stall angle when morphing the leading edge. The behavior of the aerodynamics performances from the MTE and the MLE should allow individual control of the morphing camber trailing and leading edges. The estimation of the performances of the morphing camber on an unmanned aerial vehicle indicates that the morphing of the camber allows a drag reduction. This result is due to the smaller angle of attack needed for an unmanned aerial vehicle equipped with the morphing camber system than an unmanned aerial vehicle equipped with classical aileron. In the case study, the morphing camber system was found to allow a reduction of the drag when the lift coefficient was higher than 0.48.

Heat Transfer in the Separated and Reattached Flow on a Blunt Flat Plate

1974 ◽  
Vol 96 (4) ◽  
pp. 459-462 ◽  
Author(s):  
Terukazu Ota ◽  
Nobuhiko Kon
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Flat Plate ◽  
Leading Edge ◽  
Two Dimensional ◽  
Separated And Reattached Flow ◽  
The Heat Transfer Coefficient ◽  
Blunt Leading Edge ◽  
Made In

Heat transfer measurements are made in the separated, reattached, and redeveloped regions of the two-dimensional air flow on a flat plate with blunt leading edge. The flow reattachment occurs at about four plate thicknesses downstream from the leading edge and the heat transfer coefficient becomes maximum at that point and this is independent of the Reynolds number which ranged from 2720 to 17900 in this investigation. The heat transfer coefficient is found to increase sharply near the leading edge. The development of flow is shown through the measurements of the velocity and temperature in the separated, reattached, and redeveloped regions.

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