Top PDF Computational Prediction of Tip-Vortex of a Swept Wing

Computational Prediction of Tip-Vortex of a Swept Wing

Computational Prediction of Tip-Vortex of a Swept Wing

Copyright to IJIRSET www.ijirset.com 5979 In a finite wing, there is an opportunity for the pressures acting on the upper and lower surfaces to interact near the wing tip[1]. The shorter the distance between the wing tip, the larger the downwash velocity and the induced drag[2]. The trailing vortex system also generates an upwash in the regions beyond the wing span and a downwash inside the wing span. This downwash produced by the trailing vortex system adds to the downwash produced by the bound vortex system[3].
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Computational Study of Aspect-Ratio Effects on Swept-Wing Stall.

Computational Study of Aspect-Ratio Effects on Swept-Wing Stall.

and induced downwash from tip vortices, leads to different sections along the semi-span operating at angles of attack which maybe different from the wing angle of attack. The above phenomenon is affected further for swept wings due to the spanwise propagation of separated boundary layer. Once the separation point along the section was evaluated, the local operating angle of attack was evaluated using Beddoes and Leishman model [36], developed for calculating lift over an airfoil having separation. This method, based on the Kirchhoff-Helmholtz flat-plate model, gives relations for the normal and chordwise forces experienced by the airfoil as a function of the angle of attack, α , and the non-dimensionalized chordwise location of the separation point, f . The value of f varies from 0 (leading edge, signifying fully separated flow) to 1(trailing edge, fully attached flow). The normal and chordwise force coefficients on the section are given by:
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On the turbulent mean-flow at the leading edge of a swept wing

On the turbulent mean-flow at the leading edge of a swept wing

Despite the significant change in the leading edge modelling, the profile drag prediction was not significantly affected as the predicted momentum thickness in the far wake was almost similar from both the old and the new version of Callisto. The revised computational results in Fig. 11 show a rapid growth in momentum thickness downstream of the AL followed by an equally rapid decay, presumably due to the interplay between streamline curvature and favourable pressure gradient at the LE. From the mean flow measurements it is difficult to understand and describe the physical mechanism responsible for the non-monotonic growth in θ in the vicinity of the AL, but as similar trend was predicted by Callisto a simple diagnosis was conducted by analysing the individual terms of the streamwise momentum integral equation. This stationary points in the trend of θ appears when the magnitude of the favourable pressure gradient overtakes the skin friction immediately downstream of the AL, hence slowing the growth in θ, which results in a maximum point. The minimum point is associated with the point where skin friction exceeds the magnitude of the favourable pressure gradient, thus θ grows again. Similar behaviour was also observed for calculation of flow over transonic wing at cruise condition. By x/c = 0.25, θ estimated by the modified Callisto has almost merged with that obtained from the previous version, and any residual differences are smaller than the effect of correcting for effective sweep. Based on these observations, the overall profile drag predictions obtained from earlier versions of Callisto can still be considered robust.
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An empirical approach for propeller tip vortex cavitation noise

An empirical approach for propeller tip vortex cavitation noise

vortex cavitation noise. The propeller cavitation noise has been investigated by Lafeber et al. [2] via both computational and experimental methods. ETV (Empirical cavitating Tip Vortex) noise model based on TVI technique has been used for prediction of cavitating vortex noise for three different marine propellers. Szantyr et al. [3] have focused on tip vortex cavitation in their study. The main aim of the study was to develop a reliable method for numerical prediction of tip vortex cavitation. The inception of tip vortex cavitation has been studied by Lee et al. [4]. A detailed noise spectrum analysis has been applied to determine the tip vortex cavitation noise. Wijngaarden et al. [5] have investigated the broadband inboard noise and vibration on passenger vessels for a frequency range. Hydro-acoustic calculations involving tip vortex cavitation have been examined both by sea trials and model experiments. Raestad [6] has studied on the tip vortex cavitation noise for twin screw passenger vessels. Full scale experiments have been conducted on different ship types. The tip vortex cavitation noise formulation has been developed with the help of a lifting surface method (LSM). The reference distance for the calculations has been considered as three decks above the propeller. The final report of specialist committee on hydrodynamic noise in 27th ITTC conference [7] emphasizes that TVI technique related to the tip vortex cavitation has given a good agreement with measured inboard noise data. Wave effects on cavitation and pressure pulses of a tanker with twin podded propulsor have been investigated by Taskar et al. [8]. In their study, TVI technique has been used in order to calculate the propeller noise induced by tip vortex cavitation. In the calculations, the tip circulation value has been taken from the blade section of r/R=0.997
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Energization of Boundary Layer Over Wing Surface By Vortex Generators

Energization of Boundary Layer Over Wing Surface By Vortex Generators

Abstract— The issue taken into consideration is sudden increase in drag over an aircraft wing due to three dimensional flow, tip vortices and flow separation. When flow separates its displacement thickness increases sharply this modifies the outside potential flow and pressure field. The pressure field modification results in an increase in pressure drag, and if severe enough will also result in loss of lift and stall. This study presents computational analysis results of a prototype wing with and without vortex generators of two different shapes located at leading and trailing edges of a linear wing. Here both wind tunnel testing and computational fluid dynamic analysis is carried out. The effect of the vortex generators are studied in four different cases. Nine sets of rectangular shaped vortex generators inclined at 15 degree were placed in the leading edge and trailing edge of the wing, nine sets of ogive shaped vortex generators inclined 15 degree were placed in the leading edge and trailing edge of the wing, are the cases analyzed. The studies also focus on prevention of downstream flow separation and improve overall performance by reducing drag. Both analytical and experimental results are compared where it shows that the pressure over the upper surface increases, so that the boundary layer is reenergized and attached with the body surface thus reducing the drag.
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The Aerodynamic Analysis and Aeroelastic Tailoring of a Forward-Swept Wing

The Aerodynamic Analysis and Aeroelastic Tailoring of a Forward-Swept Wing

distribution from the pressures obtained from the experimental testing. Since only two 16- port chordwise distributions were used on the pressure port wing, the precise pressure distribution was unable to be determined from the pressure plots. This occurred because the location and intensity of the chordwise pressures was unknown at locations other than 33 and 80 percent span. To enable the locations to be determined and increase the accuracy of the analysis, the number of chordwise pressure ports should be significantly increased with a large concentration of orifices near the trailing edge of the upper surface. However, the conclusions drawn from the pressure port data on separation and crossflow locations were verified through the tuft analysis. The prediction that the root separated at an alpha of 12 degrees and the flow remained attached at the wing tips above 18 degrees angle of attack was verified through the tuft analysis. Furthermore, the use of two pressure ports provided an accurate spanwise lift distribution prior to separation. The estimation of C l using the two
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Low-Order Modeling of Wing-Tip Vortices Using an Augmented Vortex Lattice Method.

Low-Order Modeling of Wing-Tip Vortices Using an Augmented Vortex Lattice Method.

Wing-tip flow effects become increasingly significant in the prediction of forces and moments on a wing as the aspect ratio decreases. Similar to the attached leading-edge vortex of delta wings, the wing-tip vortex is characterized by a separated flow structure that is not included in conventional vortex lattice method (VLM) models. These models assume fully-attached flow at the tip, and therefore fail to capture what portion of the wing-tip vortex exists as a free vortex upstream of the trailing edge. Even at small angles of attack, some degree of the shear layer separation that feeds the free tip vortex can and generally does occur. The most significant consequence of this attached vortex structure is the increased suction, and therefore, normal force developed near the wing tips. This increased normal force results in a supplementary lift known as “vortex lift” (and “vortex drag”), which acts in addition to the “linear lift” associated with fully-attached flow. For low-aspect-ratio wings this effect on lift, as well as all forces and moments, is significant.
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A vortex-based tip/smearing correction for the actuator line

A vortex-based tip/smearing correction for the actuator line

Abstract. The actuator line (AL) was intended as a lifting line (LL) technique for computational fluid dynamics (CFD) applications. In this paper we prove – theoretically and practically – that smearing the forces of the actuator line in the flow domain forms a viscous core in the bound and shed vorticity of the line. By combining a near-wake representation of the trailed vorticity with a viscous vortex core model, the missing induction from the smeared velocity is recovered. This novel dynamic smearing correction is verified for basic wing test cases and rotor simulations of a multimegawatt turbine. The latter cover the entire operational wind speed range as well as yaw, strong turbulence and pitch step cases. The correction is validated with lifting line simulations with and without viscous core, which are representative of an actuator line without and with smearing correction, respectively. The dynamic smearing correction makes the actuator line effectively act as a lifting line, as it was originally intended.
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Transonic flow past symmetrical unswept and swept wings 
		with elliptic nose

Transonic flow past symmetrical unswept and swept wings with elliptic nose

The outer boundary of the 2D lens-type computational domain is composed by two curves   and   . The size of domain along the x axis is 80 chords; the size of domain along the y axis is 200 chords. A program written in Pascal language transformed it into a three- dimensional mesh. The mesh is in the TGrid / Fluent format [15], which is suitable for the calculation in the package Ansys CFX [16]. In this paper we used a method [13] for generating a three-dimensional mesh with elements elongated along a wing. The basis of generating is the two-dimensional mesh. Elementary volumes have hexahedral shape near the surface of the wing in the boundary layer and the pentahedral shape in the rest part of the mesh. Grid is constructed sequentially by parallel transport of a two-dimensional mesh along the wing and the mesh scaling. Computational domain includes the region that surrounds the wing tip. Total size of computational domain along wingspan is 14.5 root chords for the swept wing. The sizes of domain along wingspan for unswept wings are 13.5 and 14.5 chords for aspect ratios 3 and 4 respectively.
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Mean flow and turbulence measurements in the near field of a fighter aircraft wing-tip vortex at high Reynolds number and transonic flow conditions

Mean flow and turbulence measurements in the near field of a fighter aircraft wing-tip vortex at high Reynolds number and transonic flow conditions

The investigation of the wing-tip vortex generated by a model F/A-18 fitted with AIM-9 tip missiles, via five-hole probe measurements as well as hot-wire anemometry, has provided an overview of the mean and turbulent flowfield. At α = 4.5°, the tip vortex showed certain similarities with some of those described in the literature and generated at lower speeds with wings of simpler geometry. At α = 13°, the vortex increased in strength and size, the turbulence levels were an order of magnitude higher than at α = 4.5°, and the effects of the tip missile on the flow were more prominent. The missile body produced an axial velocity deficit which became more significant at the higher angle of attack, as well as increased turbulence levels due to its own wake and that of its canards and fins. The tip missile thus significantly increases the complexity of the studied flow. More detailed measurements, ideally at several streamwise stations, could help to understand more fully the development of this complex flowfield.
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Design and Structural Analysis of the Ribs and Spars of Swept Back Wing

Design and Structural Analysis of the Ribs and Spars of Swept Back Wing

Ribs give the shape to the wing section, support the skin (prevent buckling) and act to prevent the fuel surging around as the aircraft maneuvers. They serve as attachment points for the control surfaces, flaps, under carriage and engines. The ribs need to support the wing-panels, achieve the desired aerodynamic shape and keep it, provide points for conducting large forces, add strength, prevent buckling, and separate the individual fuel tanks within the wing. Milled ribs are solid structures, manufactured by milling away excess material from the solid block of metal, and are also used where very high loads apply.
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Design and Stress Analysis of Swept Back Wing with Nastran and Patran

Design and Stress Analysis of Swept Back Wing with Nastran and Patran

Selecting the wing span is one of the most basic decisions to made in the design of a wing. The span is sometimes constrained by contest rules, hangar size, or ground facilities but when it is not we might decide to use the largest span consistent with structural dynamic constraints (flutter). This would reduce the induced drag directly. However, as the span is increased, the wing structural weight also increases and at some point the weight increase offsets the induced drag savings. This point is rarely reached, though, for several reasons.

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Effect of Wing Sweep Angle on the Vortex Interaction of a Tail-Wing Configuration

Effect of Wing Sweep Angle on the Vortex Interaction of a Tail-Wing Configuration

68 | w \#^ } w \!^ } w \!8 } w \~^when }\^ & be seen on the wing. Note that for this wing having a sweep angle of 30 the vortices unlike delta wings with leading edge sweep of 60 & type ones. A small part of the wing is affected by the &} w \!^ and the leading edge vortex covers a large portion of the wing suction surface by <G<@} w H

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Analysis of Low Speed Aerodynamics of Double Delta Wing

Analysis of Low Speed Aerodynamics of Double Delta Wing

Figure 14 to 18 shows vortex core regions over 80/56.867-deg double-delta wing at 0º, 5º, 10º, 20º, and 30º angles of attack. The method by which these vortex regions were generated was Absolute Helicity with level 0.1 for each vortex region contour. The absolute helicity vortex which is being displayed is created by a mixture of effects from the walls, the curve on the wing, and the interaction of the fluids. If I had chosen the vorticity method instead, wall effects would have dominated. At 30º angle of attack, there are strong vortices on the upper surface of the wing, which in turn indicates the region of vortex breakdown, a phenomenon which describes the deceleration of the vortex core and an increase in the vortex core diameter [16].
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Aerodynamics of tip reversal upstroke in a revolving pigeon wing

Aerodynamics of tip reversal upstroke in a revolving pigeon wing

During slow flight, bird species vary in their upstroke kinematics using either a ‘flexed wing’ or a distally supinated ‘tip-reversal’ upstroke. Two hypotheses have been presented concerning the function of the tip-reversal upstroke. The first is that this behavior is aerodynamically inactive and serves to minimize drag. The second is that the tip-reversal upstroke is capable of producing significant aerodynamic forces. Here, we explored the aerodynamic capabilities of the tip-reversal upstroke using a well- established propeller method. Rock dove ( Columba livia, N 3) wings were spread and dried in postures characteristic of either mid-upstroke or mid-downstroke and spun at in vivo Reynolds numbers to simulate forces experienced during slow flight. We compared 3D wing shape for the propeller and in vivo kinematics, and found reasonable kinematic agreement between methods (mean differences 6.4% of wing length). We found that the wing in the upstroke posture is capable of producing substantial aerodynamic forces. At in vivo angles of attack (66  deg at mid-upstroke, 46  deg at mid-downstroke), the upstroke wings averaged for three birds produced a lift-to-drag ratio of 0.91, and the downstroke wings produced a lift-to-drag ratio of 3.33. Peak lift-to-drag ratio was 2.5 for upstroke and 6.3 for downstroke. Our estimates of total force production during each half-stroke suggest that downstroke produces a force that supports 115% of bodyweight, and during upstroke a forward-directed force (thrust) is produced at 36% of body weight.
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Experimental investigation of the vorticity amplification on a swept wing with a blunt leading edge

Experimental investigation of the vorticity amplification on a swept wing with a blunt leading edge

It is known that the freestream turbulence influences transition from laminar to turbulent boundary layer [14], and in particular the cross-flow instability on swept wings [2], but the physical mechanism of interaction between freestream turbulence and boundary layer is not completely understood. The attachment line of a swept wing plays a crucial role in establishing the initial conditions for the downstream flow over an airfoil. The study of the interaction of freestream turbulence and boundary layer at leading edge may be of help in understanding this and other phenomena.
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The vortex wake of blackcaps (Sylvia atricapilla L ) measured
using high speed digital particle image velocimetry (DPIV)

The vortex wake of blackcaps (Sylvia atricapilla L ) measured using high speed digital particle image velocimetry (DPIV)

Traditionally, the study of bird aerodynamics has been restricted to kinematic analyses and modelling of force generation, simplified to varying degrees (Brown, 1953; Norberg, 1975; Norberg, 1990, Hedrick et al., 2002; Hedrick et al., 2004). But, in the past few years, direct measurements of the wake circulation of free-flying birds have become available for a number of species, allowing a direct estimate of the forces generated (Spedding et al., 2003; Warrick et al., 2005; Hedenström et al., 2006a; Hedenström et al., 2006b; Rosén at al., 2007; Henningsson et al., 2008; Tobalske et al., 2009). Circulation has been measured in the streamwise vertical plane at different positions along the span of the birds, but none of these species have been studied in the transverse plane [with the exception of hovering humming birds (Warrick et al., 2005)]. Recent findings from studies of bat wakes have pointed to the importance of transverse plane data for the reconstruction of the wake topology because important structures may be otherwise missed (Hedenström et al., 2007; Johansson et al., 2008). The suggested vortex wake model for bats (Hedenström et al., 2007) is more complex than the ones suggested for birds (Spedding et al., 2003; Henningsson et al., 2008) and includes streamwise structures (for example vortices shed at the wing root) not suggested by the bird data (Spedding et al., 2003). Whether these differences between birds and bats in the suggested wake models are due to real differences or rather to the lack of appropriate data in the bird’s case is thus an open question.
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Experimental investigation of wing tip vortices in the near-field

Experimental investigation of wing tip vortices in the near-field

The cross-sectional lines algorithm [13] has been used for obtaining coordinates of the tip vortices. In this method, a vortex origin is defined as a point, where the differences of largest and smallest tangential components of velocities along horizontal and vertical axes take largest values. To this end, a code was written to find vortex coordinates for each distance from the trailing edge to the laser sheet and for every angle of attack.

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Reduction of Induced Drag Using Wing Tip Propeller

Reduction of Induced Drag Using Wing Tip Propeller

[16] Smith M. J., Komerath N., Ames R., Wong O., and Pearson J., “Performance Analysis of A Wing with Multiple Winglets,” AIAA Jr., 2001, Pp. 2001-2407. [17] Smith S.C., Kroot I. M., “Induced Drag Computations on Wings with Accurately Modelled Wakes,” Jr. Aircraft, Vol. 34, No. 2, Pp. 253- 255. [18] Srikanth G., Bogadi S., “Experimental Investigation on the Effect of Multi-Winglets”, Int. Jr. of Mech. & Ind. Engg., Vol.1, Issue-1, 2011.

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Soaring and Gliding Flight of the Black Vulture

Soaring and Gliding Flight of the Black Vulture

The opening of the wing feathers clearly increases the span of the wing and thereby reduces the trailing vortex, or induced drag, but probably at the expense of increasing the profile dr[r]

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