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Influence of different flap settings on the wake-vortex structure of a rectangular wing with flaps and means of alleviation with wing fins
Aerosp. Sci. Technol. 4 (2000) 79–90 2000 Éditions scientifiques et médicales Elsevier SAS. All rights reserved S1270-9638(00)00120-6/FLA
Influence of different flap settings on the wake-vortex structure of a rectangular wing with flaps and means of alleviation with wing fins Ingmar Schell a, *, Erol Özger a , Dieter Jacob a a Institut für Luft- und Raumfahrt (ILR), RWTH Aachen, Wüllnerstraße 7, 52062 Aachen, Germany
Received 11 October 1999; revised and accepted 17 January 2000
Abstract
This paper presents the results of an experimental investigation of the wake-vortex structure behind a flapped rectangular wing. The wind-tunnel test is part of a project funded by the German Research Association (DFG) to investigate the flow structure in the wake of wings with different flap extensions. The purpose of this investigation is to study the main features of lift generated vortices in order to find ways to alleviate hazardous wake vortex encounters for following airplanes during start and approach such that an increase in airport capacity can be achieved. First, the wake structure at different flap settings is investigated by measuring the velocity field at different positions in the near field behind the wing. The measured data are evaluated in terms of vortex parameters such as core radius, maximum tangential velocities, circulation distributions, turbulence levels and maximum induced rolling moments on a following aircraft. Then additionally, for one flap setting, means of alleviation are examined with a wing fin mounted at different positions on the wing. Alleviation is judged by the decrease of the maximum induced rolling moment. A significant reduction is achieved for a wing fin placed near the outboard edge of the flap. 2000 Éditions scientifiques et médicales Elsevier SAS flapped wing / wake structure / wake vortex alleviation / induced rolling moment
Zusammenfassung
Einfluß von verschiedenen Klappenstellungen auf die Nachlaufstruktur eines Rechteckflügels mit Klappen und Möglichkeiten der Wirbelabschwächung mit Hilfe von Finnen. In diesem Bericht werden die Ergebnisse der experimentellen Untersuchung des Wirbelnachlaufs eines Rechteckflügels mit Klappen diskutiert. Die im Windkanal durchgeführten Versuche sind Teil eines von der Deutschen Forschungsgemeinschaft (DFG) finanzierten Projektes, in dem unter anderem die Strömung im Nachlauf von Flügeln mit und ohne Klappen untersucht wird. Das Ziel dieser Untersuchung ist die Bestimmung von charakteristischen Merkmalen des Wirbelnachlaufs, um anschließend gezielte Maßnahmen zur Verringerung der Gefahren für nachfolgende Flugzeuge bei Start und Landung zu ergreifen und dadurch die Kapazitäten von Flughäfen zu erhöhen. Im Rahmen dieser Arbeit wird zuerst die Nachlaufstruktur für verschiedene Klappenstellungen durch Messungen der Geschwindigkeitsverteilungen an verschiedenen Positionen im Nahfeld hinter dem Flügel untersucht. Die gemessenen Daten werden bezüglich verschiedener Wirbelparameter ausgewertet, wie z.B. Kernradius, maximale Umfangsgeschwindigkeit, Zirkulations- und Turbulenzverteilung sowie maximales induziertes Rollmoment auf ein nachfolgendes Flugzeug. Anschließend wird als Maßnahme für die Abschwächung des Wirbelnachlaufs zusätzlich eine Finne an verschiedenen Orten auf dem Flügel für eine feste Klappenstellung angebracht. Die Abschwächung wird durch die erzielte Abnahme des induzierten Rollmomentes beurteilt. Eine deutliche Verringerung wird durch eine Finne erreicht, die in der Nähe des äußeren Klappenendes angebracht wird. 2000 Éditions scientifiques et médicales Elsevier SAS
* Correspondance and reprints; E-mail: [email protected]
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Flügel mit Klappen / Nachlaufstruktur / Wirbelabschwächung / induziertes Rollmoment
Nomenclature b bf c cF c1 c1,max cL cLα D dP f hF L lf lP Rc Re t u, v, w u0 U∞ VP Vθ,max x y z α δ εF Γ Λ ω
wing span wing span of following aircraft wing chord wing fin chord induced rolling moment coefficient max. induced rolling moment coefficient lift coefficient lift gradient diameter of wind-tunnel nozzle diameter of sensors recording frequency of the hotwire probe wing fin height length of test section length of flap section active length of sensors core radius Reynolds number measurement period velocity components in x, y, z-direction turbulence in freestream direction freestream velocity measurement volume of the hotwire probe maximum tangential velocity distance in freestream direction distance in spanwise direction distance in vertical direction angle of attack flap angle angle of incidence of wing fin circulation aspect ratio vorticity
1. Introduction The shedding of strong vortices in the wake of starting and landing aircraft poses a hazard for aircraft following. This leads to a limitation of the capacity of airports due to detrimental starting and landing frequencies. Although this problem has been in existence for forty years it is pertinent at the moment because of projects like future high capacity transport aircraft (e.g. A3XX) due to the large strength of their shedded vortices. Many investigations on the wake structure of unflapped rectangular wings have been performed experimentally [5,7] as well as numerically, starting from simple Betz methods [1] to elaborated Navier–Stokes computations [6,9]. Investigations of flapped wings [2,3,10] are fewer in number and not detailed enough to cover, for example, the influence
of different flap settings on the wake structure. Therefore, the first part of this paper investigates experimentally the wake structure in the near field of a rectangular wing for different flap settings. Vortex core parameters such as core radius and maximum tangential velocities are presented, as well as overall vorticity and circulation distributions. Another aspect of this work is the alleviation of the vortex wake. In the past and nowadays many investigations have been conducted to explore methods of decreasing the rolling moment induced on an aircraft following. The value of the induced rolling moment depends on the circulation and core radius of the vortices, their relative positions and on the span of the following aircraft. All the attenuating methods have in common the dispersion of circulation over a large area in the wake. This can be achieved by either convectivity or viscosity dominated transport mechanisms. Using spoilers or mass injection in the region of the vortices [4,11] may be attributed to the viscosity dominated transport mechanisms. On the one hand in these investigations the vortices widened and the maximum induced rolling moment decreased at various rates. On the other hand Dunham [8] reports that the decrease in maximum induced rolling moment of the vortices due to mass injection was negligible. Thus, various views on the effectiveness of these methods exist. The convective transport can be influenced by changing the wing loading by means of different wing geometries and flap deflection. Rossow [12] gives a theoretical introduction to various concepts based on convective transport. Dunham [8] finds that a modification of the wing tip shape (Ogee tip) does not decrease the maximum induced rolling moment significantly. Rossow [13–16] is the first person to present a device to alleviate significantly the hazard posed by wake vortices, with small losses in lift. He investigates the effect of wing fins on the maximum induced rolling moment and finds a decrease of approximately 60%. Rossow attributes this change to the convective transport triggered by the additional wing fin vortex that disperses the vortex wake. However, this explanation has not been validated so far, so that the mechanism still remains unclear. In [15] he additionally measures the downwash velocities and states that they are diminished in the region of maximum induced rolling moment. Therefore, the second part of this paper deals with the effect of wing fins on the wake structure and the achievable alleviation. The objective is to identify whether the attenuation is based on convective or viscous mechanisms or on both.
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2. Experimental setup The experiments are conducted in the Göttingen type low speed wind-tunnel of the Institut für Luft- und Raumfahrt (ILR). The open test section has a length of L = 3 m and a nozzle diameter of D = 1.5 m. The maximum freestream velocity is U∞ = 75 m/s. In these experiments a half-model mounted on a test rack is used (see figure 1). Suction is applied at the nose of the rack to reduce the boundary layer growth on the half model symmetry plane. In order to adjust the angle of attack α of the wing models an electrically powered turntable is integrated within the rack. Aerodynamic forces and moments are measured by a six-component strain gauge balance. Velocities (u, v, w) in the wake of the wing are measured by means of an automatically traversed probe. The probe can be traversed in all three axis directions (x, y, z) so that velocities can be recorded in arbitrary planes perpendicular to the freestream velocity. Figure 2 shows the wing model. It is a flapped rectangular wing with a span of b = 1.11 m and a chord length of c = 0.12 m (aspect ratio Λ = 9.25). The wing is segmented in spanwise direction, each segment having a deployable flap. In the scope of the experiments only the flaps of the four inner segments are deployed (hatched area in figure 2) similar to the flap setting of a commercial aircraft in starting and landing configuration.
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The wing profile is an Eppler-1232 profile and the flaps can be deployed at δ = 0◦ , 10◦ , 20◦ , 30◦ . The open jet boundary can be considered to be sufficiently small for the investigated configurations (b/2D = 0.37). The measurements are performed by means of hotwire anemometry. The 3D-hotwire probes are designed and manufactured in the Aerodynamic Institute Aachen. The diameter of the sensors is dP = 5 µm, the active length is lP = 1 mm so that the measurement volume of the hotwire probe is about VP = 1 mm3 . Each of the hotwire sensors is operated separately using a Dantec Frame 90N10 with three anemometer units CTA 90C10. For calibration of the probes the Dantec calibration system 90H10 is used. The velocity data are recorded at a frequency of f = 1 kHz for a measurement period of t = 1 sec. The velocities downstream of the wing are measured in planes perpendicular to the freestream velocity at four distances (x/b = 0.0, 0.2, 0.4 and 1.0). First, planes with coarse step size (1y = 1z = 7.5 mm) are measured to evaluate the positions of the vortices and to get an overall impression of the wake structure. Then, grids with a fine step size (1y = 1z = 2 mm) are used to resolve finer structures of the vortices. In figure 3 the coordinate system and the investigated areas are shown. Moreover the vortices shed at the wing tip and the flap edges are sketched with their sense of rotation.
Figure 2. Sketch of flapped wing model.
Table I. Measured cases Angle of attack α Force
Figure 1. Experimental rack in open test section.
Flap angle δ
Wake Measurements
−10◦ –10◦
4◦
0◦
−10◦ –10◦
4◦
10◦
−10◦ –10◦
4◦
20◦
−10◦ –10◦
4◦
30◦
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Figure 3. Coordinate system and arrangement of measured planes.
Figure 5. Circulation distributions Γ /bU∞ for various flap settings. Figure 4. Lift curves cL for various flap settings.
3. Wake-vortex structure at different flap settings Table I shows the measured cases which will be discussed in the following. The experiments were performed at a freestream velocity of U∞ = 40 m/s without fixed transition and a Reynolds number based on the wing chord of about Re = 3 · 105 . Force measurements were conducted at an angle of attack range between α = −10◦ to α = 10◦ . For the wake vortex investigation a moderate angle of attack of α = 4◦ was chosen to ensure that no flow separation occurs on the wing for all flap configurations. The lift curves are shown in figure 4. At higher angles of attack (α > 4◦ ) a flap setting of δ = 30◦ has a lower lift than at δ = 20◦ because at δ = 30◦ and high angles of attack, local separation occurs.
The velocity data are evaluated to compute the circulation Γ in the wake of the wing according to I Γ =
vE · d sE
(1)
along the outer contour of the coarse grid. Figure 5 shows the circulation distributions at the trailing edge of the various configurations for α = 4◦ . For increasing flap angles the circulation in the flap area is also increasing. The circulation distributions are consistent with the measured lift data. The decrease in circulation for flap angles higher than δ = 20◦ is caused by separation. The spanwise oscillations in the circulation distributions can be explained by local separation due to the interrupted flap segments.
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Figure 6. Vorticity distributions ωc/U∞ in the wake of the wing.
The vorticity ω can be computed from the velocity data according to ω=
∂w ∂v − . ∂y ∂z
(2)
Typical vorticity distributions in several planes measured in the wake of the wing with a flap angle of δ = 20◦ are shown in figure 6. The wake of the wing is rolling up into three stable vortices, the wing tip vortex, the inboard and the outboard flap vortex. The mutual velocity inductions make the wing tip vortex move upward in positive z-direction while the flap vortex pair is moving downward in the negative z-direction. For a more detailed evaluation of the vortex parameters the region of the vortices were measured with a smaller step size of 1y = 1z = 2 mm. In figure 7 and 8 the vorticity distribution of the three vortices is shown in the planes x/b = 0 and x/b = 0.2. The wing tip vortex shows a single branched structure with one vorticity maximum defining the vortex center. The flap end vortices show a double branched structure with two vorticity maxima at x/b = 0 which merge into one maximum downstream at x/b = 0.2. The vortex structure described above consists of a dominant and a weak vortex. The first is lying in the flap gap between the trailing edges of the deployed flap and the retracted part of the wing and is mainly generated by the pressure difference at the flap gap
between the suction and the pressure side of the deployed flap. The weak vortex is lying beside the dominant vortex at the suction side of the deployed flap. In contrast to the dominant one, the weak vortex is generated by the pressure difference between the suction sides of the deployed and retracted part of the wing. Important vortex parameters such as core radius Rc and maximum tangential velocity Vθ,max are evaluated from the velocity data measured with the fine grids. The vortex center is defined as the locus of maximum vorticity; the core radius is defined by the location of the maximum induced tangential velocities. Figures 9 and 10 show the core radius Rc and the maximum tangential velocity Vθ,max for the wing tip and the outboard flap vortex at different downstream positions. The inboard vortex is omitted in the further investigation due to the lower amount of circulation and the less important role in the wake vortex problem compared to the other two vortices. Both the core radius Rc and the maximum tangential velocity Vθ,max of the wing tip vortex are almost independent of the flap angle. This fact may be explained by the distance between the flow at the flap region and the flow at the wing tip region. Thus, the level of interference between the wing tip and outboard flap vortex is rather low in the region measured until x/b = 1.0. In contrast to this the core radius Rc and the maximum tangential ve-
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Figure 7. Vorticity distribution ωc/U∞ at x/b = 0.0 (α = 4◦ , δ = 20◦ ).
Figure 8. Vorticity distribution ωc/U∞ at x/b = 0.2 (α = 4◦ , δ = 20◦ ).
locity Vθ,max of the outboard flap vortex depend strongly on the flap angle δ. Increasing the flap angle to δ = 10◦ and δ = 20◦ leads to an increase of the amount of circulation in the outboard flap vortex and a reduction of the core radius and therefore to higher maximum tangential velocities Vθ,max . At δ = 30◦ stronger separation at the flap edges leads to an increase of the values in core radius of the outboard flap vortex especially at greater distances downstream from the trailing edge. Therefore, the maximum tangential velocity Vθ,max is decreasing. To measure the hazard of the wake vortex posed to following aircraft the induced rolling moment is computed
from the experimental data in the wake according to
c1 =
2 bf2
Zbf /2 −bf /2
w cLα · arctan · ξ dξ. u
(3)
The induced rolling moment coefficient c1 is computed by means of the strip theory where the following wing is divided into a series of chordwise strips. Each strip is treated as a two dimensional airfoil and the lift induced by the wake of the generator aircraft is computed as a function of the local flow angle, which is obtained from
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Figure 9. Core radius Rc /c and maximum tangential velocity Vθ,max for the wing tip vortex.
Figure 10. Core radius Rc /c and maximum tangential velocity Vθ,max for the outboard flap vortex.
the measured velocity components u and w. Together with a lift gradient cLα = 2π and a lever arm ξ the incremental rolling moment is integrated on the whole following span bf . Here, a span ratio of bf /b = 0.2 is chosen which is a typical ratio for such an investigation [8,12,13]. Figure 11 shows a typical induced rolling moment distribution in the wake of the wing at a distance of x/b = 1.0. As expected the maximum induced rolling moments are encountered in the region of the wing tip and the outboard flap vortex. Especially, in the near field region of the wake two components inducing the rolling moment on a following aircraft have to be distinguished. One part is generated by the downwash of the bound vortices of the wing which fades away at larger distances to the wing. The other part is generated by the downwash of the trailing vortices that are rather persistent in the wake flow up to many spans behind the wing. In figure 12 the maximum induced rolling moments c1,max are presented as a function of the distance to the trailing edge of the wing. For each flap setting the maximum induced rolling moment c1,max shows a high value at the trailing edge (x/b = 0) caused by higher levels of downwash of the bound vortices in addition to the velocity induction of the trailing vortices. Then, a more or less
steep fall is followed by an monotonic increase in maximum induced rolling moment. The diminishing factor of the induced rolling moment is the decreasing downwash of the bound vortices further downstream of the wing. However, due to the induced velocities of the wing tip and outboard flap vortex which are rotating around each other, the induced rolling moment is increasing. This can be explained by the fact that the angle between the connecting line of the vortices and the horizontally flying follower wing is approaching π/2. Thus, the induced velocities of both vortices amplify each other leading to higher induced rolling moments. With increasing flap angles up to δ = 20◦ the induced rolling moments cl,max increase mainly because of the increasing circulation of the outboard flap vortex. At δ = 30◦ the induced rolling moment shows a decline which can be attributed to the larger core radius of the outboard flap vortex and the more dissipated flow structure in the wake of the flap region. An examination of the y-, z-positions of the maximum induced rolling moments shows that downstream from station x/b = 0.2 the maximum induced rolling moments are always lying near the wing tip vortex. The induced rolling moments are computed within the measured region corresponding to the near field of the wing. Although a relevant wake vortex encounter will
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Figure 11. Distribution of the rolling moment c1 induced on a following aircraft at x/b = 1.0.
Figure 12. Maximum induced rolling moment c1,max on following aircraft (wings horizontal) versus distance to the trailing edge.
occur at larger distances to the wing differences in maximum induced rolling moments in the near field will manifest itself also in the far field region. Measurements conducted in [13] showed significant alleviation measured at about 14 spans behind the generator model which makes the assumption mentioned above reasonable. 4. Alleviation of wake-vortex by means of wing fins In the second part of the project the alleviation of the induced rolling moments by fins mounted on the wing was investigated. In figure 13 the outboard part of the wing model is shown with a wing fin at the middle
Figure 13. Wing model with wing fin.
position. The wing fin with height hF /b = 0.026 and chord cF /b = 0.026 was located at three positions (tip, middle and flap) and at two angles of incidence (εF = +15◦/ − 15◦ ). In this paper only the cases of the wing fin at the flap position will be discussed because this was the most effective position in terms of wake vortex alleviation. This testing program is shown in table II. Force measurements show (figure 14) that the lift coefficient cL is not significantly influenced by the wing fin. The effect of the wing fin on the wake is displayed in the following figures. The wake of the wing fin consists of a fin tip vortex, whose sense of rotation depends on the angle of incidence εF , and of a necklace vortex at the base of the fin. The
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Table II. Measured cases with wing fin Angle of attack α Force
Flap angle δ
Angle of incidence εF
Wake
Measurements −10◦ –10◦
4◦
20◦
15◦
−10◦ –10◦
4◦
20◦
−15◦
necklace or horseshoe vortex that wraps around the root of the wing fin could not be resolved sufficiently in the measurements. The additional vorticity in the outboard flap vortex region produced by the wing fin can be seen at the first station x/b = 0.0 by comparing figure 15 and figure 6. The wing fin also leads to diffusion of the vorticity ω downstream of the region where it is mounted. The increased transport of vorticity in the outboard flap region is caused by the higher level of turbulence generated by the partially separated flow behind the wing fin. The difference in the turbulence level between the case without and with wing fin can be seen in figures 16 and 17 which show the turbulence distribution in freestream direction (u0 /U∞ ). The amount of wake diffusion can be quantified by determining the influence of the wing fin on the core radius Rc and on the maximum tangential velocity Vθ,max . In figure 18 these values are presented according to Rc, with fin − Rc, without fin , Rc, without fin Vθ,min with fin − Vθ max,without fin . 1Vθ,max = Vθ,min, without fin 1Rc =
Figure 14. Influence of fin on lift.
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(4) (5)
The core radius of the wing tip vortex shows very little change whereas the maximum tangential velocity decreases up to 15%. The outboard flap vortex is much more influenced by the wing fin, which leads to a strong increase of the core radius by more than 100% and
Figure 15. Vorticity distributions ωc/U∞ in the wake of the wing with wing fin.
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Figure 16. Turbulence distribution in freestream direction (u0 /U∞ ) without wing fin.
Figure 17. Turbulence distribution in freestream direction (u0 /U∞ ) with wing fin.
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Figure 18. Relative difference in core radius Rc and maximum tangential velocity Vθ,max of wing tip and outboard flap vortex due to the wing fin.
to a decrease of the maximum tangential velocity by about 50%. This seems plausible because the wing fin is closer to the outboard flap vortex. As a consequence the outboard flap vortex gets into the highly turbulent wake of the wing fin and widens under this influence. The amount of circulation within the wing tip and outboard flap vortex is investigated in figure 19. The wing tip vortex contains less circulation in multiples of core radius for the wing fin cases. Thus, the structure of the wing tip vortex seems to be changed by the presence of a wing fin. The amount of circulation of the outboard flap vortex in multiples of core radius is less systematically influenced by the presence of the wing fin. The diffused wake structure with wing fin has a strong influence on the maximum induced rolling moments c1,max for a following aircraft (bf /b = 0.2). Figure 20 shows that a wing fin reduces the maximum induced rolling moment c1,max at x/b = 1.0 by almost 40%. The small difference in the maximum induced rolling moment for the wing fin with positive and negative angle of incidence suggests that the turbulence in the wake of the fin has a stronger influence on the induced rolling moment than the circulation generated by the fin. However, the y-, z-position of the maximum induced
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Figure 19. Circulation distribution Γ /bU∞ in wing tip vortex and outboard flap vortex.
Figure 20. Maximum induced rolling moment c1,max without and with wing fin.
rolling moments is little influenced by the wing fin. It remains near the wing tip vortex. The influence of wing fins on the induced rolling moment further downstream has to be investigated.
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5. Conclusion The experiments described in this paper deal with the influence of different flap settings on the wake structure of a flapped wing and the alleviating effect of wing fins on the induced rolling moments. Deploying the flap changes the wake structure from a one vortex system to a multiple vortices system. Three vortices are shed, one at the wing tip and one at each flap edge. For increasing flap angles the wing tip vortex remains approximately unchanged while the flap vortices are affected. Their core radius is reduced at lower flap angles followed by a widening at higher flap angles due to an increased flow separation at the flap edges. Vortex parameters such as core radius and maximum tangential velocity as well as vorticity and circulation distributions were investigated and presented. By means of a wing fin mounted near the outboard edge of the flap a significant decrease of the maximum induced rolling moment for a horizontally following wing could be achieved at one span downstream of the generator wing. This decrease can be attributed both to convectively and viscously dominated transport mechanisms. The amount of circulation within the core of the wing tip vortex is diminished. The outboard flap vortex is highly diffused and widened by the influence of the turbulent wake of the wing fin. Both effects lead to a changed velocity distribution which induces lower maximum rolling moments on the following aircraft.
References [1] Betz A., Verhalten von Wirbelsystemen, Z. Angew. Math. Mech. 12 (3) (1932) 164–174. [2] Bruin de A.C., Hegen S.H., Rohne P.B., Spalart P.R., Flowfield survey in trailing vortex system behind a civil aircraft model at high lift, in: The Characterization & Modification of Wakes from Lifting Vehicles in Fluids, AGARD FDP Symposium, Trondheim, Norway, May 1996, published in CP-584. [3] Chen A.L., Jacob J.D., Savas Ö., Dynamics of corotating vortex pairs in the wake of flapped airfoils, J. Fluid Mech. 382 (1999) 155–193.
[4] Corsiglia V.R., Jacobsen R.A., Chigier N., An experimental investigation of trailing vortices behind a wing with a vortex dissipator, in: Proceedings of a Symposium on Aircraft Wake Turbulence, Seattle, Washington, September 1970, pp. 229–242. [5] Corsiglia V.R., Rossow V.J., Ciffone D.L., Experimental study of the effect of span loading on aircraft wakes, J. Aircraft. 13 (12) (1976) 968–973. [6] Dacles-Mariani J., Zilliac G.G., Chow J.S., Bradshaw P., Numerical/experimental study of a wingtip vortex in the near field, AIAA J. 33 (9) (1995) 1561−1568. [7] Devenport W.J., The structure and development of a wingtip vortex, J. Fluid Mech. 312 (1996) 67–106. [8] Dunham R.E., Unsuccessful Concepts for Aircraft Wake Vortex Minimization, Symposium held at Washington DC, February 1976, NASA SP-409, 1977, pp. 221–249. [9] Eaton J.A., O’ Flaherty M.P., Flowfield prediction of three-dimensional wing trailing vortices using advanced turbulence models, in: The Characterization & Modification of Wakes from Lifting Vehicles in Fluids, AGARD FDP Symposium, Trondheim, Norway, May 1996, published in CP-584. [10] Hueneke K., Structure of a transport aircraft-type near field wake, in: The Characterization & Modification of Wakes from Lifting Vehicles in Fluids, AGARD FDP Symposium, Trondheim, Norway, May 1996, published in CP-584. [11] Philips W.R.C., Graham J.A., Reynolds-stress measurements in a turbulent trailing vortex, J. Fluid Mech. 147 (1984) 353–371. [12] Rossow V.J., Inviscid Modelling of Aircraft Trailing Vortices, Symposium held at Washington DC, February 1976, NASA SP-409, 1977, pp. 9–59. [13] Rossow V.J., Effects of wing fins on lift-generated wakes, J. Aircraft 15 (3) (1978) 160–167. [14] Rossow V.J., Experimental Investigation of Wing Fin Configurations for Alleviation of Vortex Wakes of Aircraft, NASA TM 78250, November 1978. [15] Rossow V.J., Sacco J.N., Askins P.A., Bisbee L.S., Smith S.M., Wind-tunnel measurements of hazards posed by lift-generated wakes, J. Aircraft 32 (2) (1995) 278–284. [16] Rossow V.J., Fong R.K., Wright N.S., Bisbee L.S., Vortex wakes of two transports in 80 by 120 foot wind tunnel, J. Aircraft 33 (2) (1996) 399–406.
Influence of different flap settings on the wake-vortex structure of a rectangular wing with flaps and means of alleviation with wing fins Ingmar Schell a, *, Erol Özger a , Dieter Jacob a a Institut für Luft- und Raumfahrt (ILR), RWTH Aachen, Wüllnerstraße 7, 52062 Aachen, Germany
Received 11 October 1999; revised and accepted 17 January 2000
Abstract
This paper presents the results of an experimental investigation of the wake-vortex structure behind a flapped rectangular wing. The wind-tunnel test is part of a project funded by the German Research Association (DFG) to investigate the flow structure in the wake of wings with different flap extensions. The purpose of this investigation is to study the main features of lift generated vortices in order to find ways to alleviate hazardous wake vortex encounters for following airplanes during start and approach such that an increase in airport capacity can be achieved. First, the wake structure at different flap settings is investigated by measuring the velocity field at different positions in the near field behind the wing. The measured data are evaluated in terms of vortex parameters such as core radius, maximum tangential velocities, circulation distributions, turbulence levels and maximum induced rolling moments on a following aircraft. Then additionally, for one flap setting, means of alleviation are examined with a wing fin mounted at different positions on the wing. Alleviation is judged by the decrease of the maximum induced rolling moment. A significant reduction is achieved for a wing fin placed near the outboard edge of the flap. 2000 Éditions scientifiques et médicales Elsevier SAS flapped wing / wake structure / wake vortex alleviation / induced rolling moment
Zusammenfassung
Einfluß von verschiedenen Klappenstellungen auf die Nachlaufstruktur eines Rechteckflügels mit Klappen und Möglichkeiten der Wirbelabschwächung mit Hilfe von Finnen. In diesem Bericht werden die Ergebnisse der experimentellen Untersuchung des Wirbelnachlaufs eines Rechteckflügels mit Klappen diskutiert. Die im Windkanal durchgeführten Versuche sind Teil eines von der Deutschen Forschungsgemeinschaft (DFG) finanzierten Projektes, in dem unter anderem die Strömung im Nachlauf von Flügeln mit und ohne Klappen untersucht wird. Das Ziel dieser Untersuchung ist die Bestimmung von charakteristischen Merkmalen des Wirbelnachlaufs, um anschließend gezielte Maßnahmen zur Verringerung der Gefahren für nachfolgende Flugzeuge bei Start und Landung zu ergreifen und dadurch die Kapazitäten von Flughäfen zu erhöhen. Im Rahmen dieser Arbeit wird zuerst die Nachlaufstruktur für verschiedene Klappenstellungen durch Messungen der Geschwindigkeitsverteilungen an verschiedenen Positionen im Nahfeld hinter dem Flügel untersucht. Die gemessenen Daten werden bezüglich verschiedener Wirbelparameter ausgewertet, wie z.B. Kernradius, maximale Umfangsgeschwindigkeit, Zirkulations- und Turbulenzverteilung sowie maximales induziertes Rollmoment auf ein nachfolgendes Flugzeug. Anschließend wird als Maßnahme für die Abschwächung des Wirbelnachlaufs zusätzlich eine Finne an verschiedenen Orten auf dem Flügel für eine feste Klappenstellung angebracht. Die Abschwächung wird durch die erzielte Abnahme des induzierten Rollmomentes beurteilt. Eine deutliche Verringerung wird durch eine Finne erreicht, die in der Nähe des äußeren Klappenendes angebracht wird. 2000 Éditions scientifiques et médicales Elsevier SAS
* Correspondance and reprints; E-mail: [email protected]
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Flügel mit Klappen / Nachlaufstruktur / Wirbelabschwächung / induziertes Rollmoment
Nomenclature b bf c cF c1 c1,max cL cLα D dP f hF L lf lP Rc Re t u, v, w u0 U∞ VP Vθ,max x y z α δ εF Γ Λ ω
wing span wing span of following aircraft wing chord wing fin chord induced rolling moment coefficient max. induced rolling moment coefficient lift coefficient lift gradient diameter of wind-tunnel nozzle diameter of sensors recording frequency of the hotwire probe wing fin height length of test section length of flap section active length of sensors core radius Reynolds number measurement period velocity components in x, y, z-direction turbulence in freestream direction freestream velocity measurement volume of the hotwire probe maximum tangential velocity distance in freestream direction distance in spanwise direction distance in vertical direction angle of attack flap angle angle of incidence of wing fin circulation aspect ratio vorticity
1. Introduction The shedding of strong vortices in the wake of starting and landing aircraft poses a hazard for aircraft following. This leads to a limitation of the capacity of airports due to detrimental starting and landing frequencies. Although this problem has been in existence for forty years it is pertinent at the moment because of projects like future high capacity transport aircraft (e.g. A3XX) due to the large strength of their shedded vortices. Many investigations on the wake structure of unflapped rectangular wings have been performed experimentally [5,7] as well as numerically, starting from simple Betz methods [1] to elaborated Navier–Stokes computations [6,9]. Investigations of flapped wings [2,3,10] are fewer in number and not detailed enough to cover, for example, the influence
of different flap settings on the wake structure. Therefore, the first part of this paper investigates experimentally the wake structure in the near field of a rectangular wing for different flap settings. Vortex core parameters such as core radius and maximum tangential velocities are presented, as well as overall vorticity and circulation distributions. Another aspect of this work is the alleviation of the vortex wake. In the past and nowadays many investigations have been conducted to explore methods of decreasing the rolling moment induced on an aircraft following. The value of the induced rolling moment depends on the circulation and core radius of the vortices, their relative positions and on the span of the following aircraft. All the attenuating methods have in common the dispersion of circulation over a large area in the wake. This can be achieved by either convectivity or viscosity dominated transport mechanisms. Using spoilers or mass injection in the region of the vortices [4,11] may be attributed to the viscosity dominated transport mechanisms. On the one hand in these investigations the vortices widened and the maximum induced rolling moment decreased at various rates. On the other hand Dunham [8] reports that the decrease in maximum induced rolling moment of the vortices due to mass injection was negligible. Thus, various views on the effectiveness of these methods exist. The convective transport can be influenced by changing the wing loading by means of different wing geometries and flap deflection. Rossow [12] gives a theoretical introduction to various concepts based on convective transport. Dunham [8] finds that a modification of the wing tip shape (Ogee tip) does not decrease the maximum induced rolling moment significantly. Rossow [13–16] is the first person to present a device to alleviate significantly the hazard posed by wake vortices, with small losses in lift. He investigates the effect of wing fins on the maximum induced rolling moment and finds a decrease of approximately 60%. Rossow attributes this change to the convective transport triggered by the additional wing fin vortex that disperses the vortex wake. However, this explanation has not been validated so far, so that the mechanism still remains unclear. In [15] he additionally measures the downwash velocities and states that they are diminished in the region of maximum induced rolling moment. Therefore, the second part of this paper deals with the effect of wing fins on the wake structure and the achievable alleviation. The objective is to identify whether the attenuation is based on convective or viscous mechanisms or on both.
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2. Experimental setup The experiments are conducted in the Göttingen type low speed wind-tunnel of the Institut für Luft- und Raumfahrt (ILR). The open test section has a length of L = 3 m and a nozzle diameter of D = 1.5 m. The maximum freestream velocity is U∞ = 75 m/s. In these experiments a half-model mounted on a test rack is used (see figure 1). Suction is applied at the nose of the rack to reduce the boundary layer growth on the half model symmetry plane. In order to adjust the angle of attack α of the wing models an electrically powered turntable is integrated within the rack. Aerodynamic forces and moments are measured by a six-component strain gauge balance. Velocities (u, v, w) in the wake of the wing are measured by means of an automatically traversed probe. The probe can be traversed in all three axis directions (x, y, z) so that velocities can be recorded in arbitrary planes perpendicular to the freestream velocity. Figure 2 shows the wing model. It is a flapped rectangular wing with a span of b = 1.11 m and a chord length of c = 0.12 m (aspect ratio Λ = 9.25). The wing is segmented in spanwise direction, each segment having a deployable flap. In the scope of the experiments only the flaps of the four inner segments are deployed (hatched area in figure 2) similar to the flap setting of a commercial aircraft in starting and landing configuration.
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The wing profile is an Eppler-1232 profile and the flaps can be deployed at δ = 0◦ , 10◦ , 20◦ , 30◦ . The open jet boundary can be considered to be sufficiently small for the investigated configurations (b/2D = 0.37). The measurements are performed by means of hotwire anemometry. The 3D-hotwire probes are designed and manufactured in the Aerodynamic Institute Aachen. The diameter of the sensors is dP = 5 µm, the active length is lP = 1 mm so that the measurement volume of the hotwire probe is about VP = 1 mm3 . Each of the hotwire sensors is operated separately using a Dantec Frame 90N10 with three anemometer units CTA 90C10. For calibration of the probes the Dantec calibration system 90H10 is used. The velocity data are recorded at a frequency of f = 1 kHz for a measurement period of t = 1 sec. The velocities downstream of the wing are measured in planes perpendicular to the freestream velocity at four distances (x/b = 0.0, 0.2, 0.4 and 1.0). First, planes with coarse step size (1y = 1z = 7.5 mm) are measured to evaluate the positions of the vortices and to get an overall impression of the wake structure. Then, grids with a fine step size (1y = 1z = 2 mm) are used to resolve finer structures of the vortices. In figure 3 the coordinate system and the investigated areas are shown. Moreover the vortices shed at the wing tip and the flap edges are sketched with their sense of rotation.
Figure 2. Sketch of flapped wing model.
Table I. Measured cases Angle of attack α Force
Figure 1. Experimental rack in open test section.
Flap angle δ
Wake Measurements
−10◦ –10◦
4◦
0◦
−10◦ –10◦
4◦
10◦
−10◦ –10◦
4◦
20◦
−10◦ –10◦
4◦
30◦
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Figure 3. Coordinate system and arrangement of measured planes.
Figure 5. Circulation distributions Γ /bU∞ for various flap settings. Figure 4. Lift curves cL for various flap settings.
3. Wake-vortex structure at different flap settings Table I shows the measured cases which will be discussed in the following. The experiments were performed at a freestream velocity of U∞ = 40 m/s without fixed transition and a Reynolds number based on the wing chord of about Re = 3 · 105 . Force measurements were conducted at an angle of attack range between α = −10◦ to α = 10◦ . For the wake vortex investigation a moderate angle of attack of α = 4◦ was chosen to ensure that no flow separation occurs on the wing for all flap configurations. The lift curves are shown in figure 4. At higher angles of attack (α > 4◦ ) a flap setting of δ = 30◦ has a lower lift than at δ = 20◦ because at δ = 30◦ and high angles of attack, local separation occurs.
The velocity data are evaluated to compute the circulation Γ in the wake of the wing according to I Γ =
vE · d sE
(1)
along the outer contour of the coarse grid. Figure 5 shows the circulation distributions at the trailing edge of the various configurations for α = 4◦ . For increasing flap angles the circulation in the flap area is also increasing. The circulation distributions are consistent with the measured lift data. The decrease in circulation for flap angles higher than δ = 20◦ is caused by separation. The spanwise oscillations in the circulation distributions can be explained by local separation due to the interrupted flap segments.
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Figure 6. Vorticity distributions ωc/U∞ in the wake of the wing.
The vorticity ω can be computed from the velocity data according to ω=
∂w ∂v − . ∂y ∂z
(2)
Typical vorticity distributions in several planes measured in the wake of the wing with a flap angle of δ = 20◦ are shown in figure 6. The wake of the wing is rolling up into three stable vortices, the wing tip vortex, the inboard and the outboard flap vortex. The mutual velocity inductions make the wing tip vortex move upward in positive z-direction while the flap vortex pair is moving downward in the negative z-direction. For a more detailed evaluation of the vortex parameters the region of the vortices were measured with a smaller step size of 1y = 1z = 2 mm. In figure 7 and 8 the vorticity distribution of the three vortices is shown in the planes x/b = 0 and x/b = 0.2. The wing tip vortex shows a single branched structure with one vorticity maximum defining the vortex center. The flap end vortices show a double branched structure with two vorticity maxima at x/b = 0 which merge into one maximum downstream at x/b = 0.2. The vortex structure described above consists of a dominant and a weak vortex. The first is lying in the flap gap between the trailing edges of the deployed flap and the retracted part of the wing and is mainly generated by the pressure difference at the flap gap
between the suction and the pressure side of the deployed flap. The weak vortex is lying beside the dominant vortex at the suction side of the deployed flap. In contrast to the dominant one, the weak vortex is generated by the pressure difference between the suction sides of the deployed and retracted part of the wing. Important vortex parameters such as core radius Rc and maximum tangential velocity Vθ,max are evaluated from the velocity data measured with the fine grids. The vortex center is defined as the locus of maximum vorticity; the core radius is defined by the location of the maximum induced tangential velocities. Figures 9 and 10 show the core radius Rc and the maximum tangential velocity Vθ,max for the wing tip and the outboard flap vortex at different downstream positions. The inboard vortex is omitted in the further investigation due to the lower amount of circulation and the less important role in the wake vortex problem compared to the other two vortices. Both the core radius Rc and the maximum tangential velocity Vθ,max of the wing tip vortex are almost independent of the flap angle. This fact may be explained by the distance between the flow at the flap region and the flow at the wing tip region. Thus, the level of interference between the wing tip and outboard flap vortex is rather low in the region measured until x/b = 1.0. In contrast to this the core radius Rc and the maximum tangential ve-
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Figure 7. Vorticity distribution ωc/U∞ at x/b = 0.0 (α = 4◦ , δ = 20◦ ).
Figure 8. Vorticity distribution ωc/U∞ at x/b = 0.2 (α = 4◦ , δ = 20◦ ).
locity Vθ,max of the outboard flap vortex depend strongly on the flap angle δ. Increasing the flap angle to δ = 10◦ and δ = 20◦ leads to an increase of the amount of circulation in the outboard flap vortex and a reduction of the core radius and therefore to higher maximum tangential velocities Vθ,max . At δ = 30◦ stronger separation at the flap edges leads to an increase of the values in core radius of the outboard flap vortex especially at greater distances downstream from the trailing edge. Therefore, the maximum tangential velocity Vθ,max is decreasing. To measure the hazard of the wake vortex posed to following aircraft the induced rolling moment is computed
from the experimental data in the wake according to
c1 =
2 bf2
Zbf /2 −bf /2
w cLα · arctan · ξ dξ. u
(3)
The induced rolling moment coefficient c1 is computed by means of the strip theory where the following wing is divided into a series of chordwise strips. Each strip is treated as a two dimensional airfoil and the lift induced by the wake of the generator aircraft is computed as a function of the local flow angle, which is obtained from
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Figure 9. Core radius Rc /c and maximum tangential velocity Vθ,max for the wing tip vortex.
Figure 10. Core radius Rc /c and maximum tangential velocity Vθ,max for the outboard flap vortex.
the measured velocity components u and w. Together with a lift gradient cLα = 2π and a lever arm ξ the incremental rolling moment is integrated on the whole following span bf . Here, a span ratio of bf /b = 0.2 is chosen which is a typical ratio for such an investigation [8,12,13]. Figure 11 shows a typical induced rolling moment distribution in the wake of the wing at a distance of x/b = 1.0. As expected the maximum induced rolling moments are encountered in the region of the wing tip and the outboard flap vortex. Especially, in the near field region of the wake two components inducing the rolling moment on a following aircraft have to be distinguished. One part is generated by the downwash of the bound vortices of the wing which fades away at larger distances to the wing. The other part is generated by the downwash of the trailing vortices that are rather persistent in the wake flow up to many spans behind the wing. In figure 12 the maximum induced rolling moments c1,max are presented as a function of the distance to the trailing edge of the wing. For each flap setting the maximum induced rolling moment c1,max shows a high value at the trailing edge (x/b = 0) caused by higher levels of downwash of the bound vortices in addition to the velocity induction of the trailing vortices. Then, a more or less
steep fall is followed by an monotonic increase in maximum induced rolling moment. The diminishing factor of the induced rolling moment is the decreasing downwash of the bound vortices further downstream of the wing. However, due to the induced velocities of the wing tip and outboard flap vortex which are rotating around each other, the induced rolling moment is increasing. This can be explained by the fact that the angle between the connecting line of the vortices and the horizontally flying follower wing is approaching π/2. Thus, the induced velocities of both vortices amplify each other leading to higher induced rolling moments. With increasing flap angles up to δ = 20◦ the induced rolling moments cl,max increase mainly because of the increasing circulation of the outboard flap vortex. At δ = 30◦ the induced rolling moment shows a decline which can be attributed to the larger core radius of the outboard flap vortex and the more dissipated flow structure in the wake of the flap region. An examination of the y-, z-positions of the maximum induced rolling moments shows that downstream from station x/b = 0.2 the maximum induced rolling moments are always lying near the wing tip vortex. The induced rolling moments are computed within the measured region corresponding to the near field of the wing. Although a relevant wake vortex encounter will
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Figure 11. Distribution of the rolling moment c1 induced on a following aircraft at x/b = 1.0.
Figure 12. Maximum induced rolling moment c1,max on following aircraft (wings horizontal) versus distance to the trailing edge.
occur at larger distances to the wing differences in maximum induced rolling moments in the near field will manifest itself also in the far field region. Measurements conducted in [13] showed significant alleviation measured at about 14 spans behind the generator model which makes the assumption mentioned above reasonable. 4. Alleviation of wake-vortex by means of wing fins In the second part of the project the alleviation of the induced rolling moments by fins mounted on the wing was investigated. In figure 13 the outboard part of the wing model is shown with a wing fin at the middle
Figure 13. Wing model with wing fin.
position. The wing fin with height hF /b = 0.026 and chord cF /b = 0.026 was located at three positions (tip, middle and flap) and at two angles of incidence (εF = +15◦/ − 15◦ ). In this paper only the cases of the wing fin at the flap position will be discussed because this was the most effective position in terms of wake vortex alleviation. This testing program is shown in table II. Force measurements show (figure 14) that the lift coefficient cL is not significantly influenced by the wing fin. The effect of the wing fin on the wake is displayed in the following figures. The wake of the wing fin consists of a fin tip vortex, whose sense of rotation depends on the angle of incidence εF , and of a necklace vortex at the base of the fin. The
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Table II. Measured cases with wing fin Angle of attack α Force
Flap angle δ
Angle of incidence εF
Wake
Measurements −10◦ –10◦
4◦
20◦
15◦
−10◦ –10◦
4◦
20◦
−15◦
necklace or horseshoe vortex that wraps around the root of the wing fin could not be resolved sufficiently in the measurements. The additional vorticity in the outboard flap vortex region produced by the wing fin can be seen at the first station x/b = 0.0 by comparing figure 15 and figure 6. The wing fin also leads to diffusion of the vorticity ω downstream of the region where it is mounted. The increased transport of vorticity in the outboard flap region is caused by the higher level of turbulence generated by the partially separated flow behind the wing fin. The difference in the turbulence level between the case without and with wing fin can be seen in figures 16 and 17 which show the turbulence distribution in freestream direction (u0 /U∞ ). The amount of wake diffusion can be quantified by determining the influence of the wing fin on the core radius Rc and on the maximum tangential velocity Vθ,max . In figure 18 these values are presented according to Rc, with fin − Rc, without fin , Rc, without fin Vθ,min with fin − Vθ max,without fin . 1Vθ,max = Vθ,min, without fin 1Rc =
Figure 14. Influence of fin on lift.
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(4) (5)
The core radius of the wing tip vortex shows very little change whereas the maximum tangential velocity decreases up to 15%. The outboard flap vortex is much more influenced by the wing fin, which leads to a strong increase of the core radius by more than 100% and
Figure 15. Vorticity distributions ωc/U∞ in the wake of the wing with wing fin.
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Figure 16. Turbulence distribution in freestream direction (u0 /U∞ ) without wing fin.
Figure 17. Turbulence distribution in freestream direction (u0 /U∞ ) with wing fin.
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Figure 18. Relative difference in core radius Rc and maximum tangential velocity Vθ,max of wing tip and outboard flap vortex due to the wing fin.
to a decrease of the maximum tangential velocity by about 50%. This seems plausible because the wing fin is closer to the outboard flap vortex. As a consequence the outboard flap vortex gets into the highly turbulent wake of the wing fin and widens under this influence. The amount of circulation within the wing tip and outboard flap vortex is investigated in figure 19. The wing tip vortex contains less circulation in multiples of core radius for the wing fin cases. Thus, the structure of the wing tip vortex seems to be changed by the presence of a wing fin. The amount of circulation of the outboard flap vortex in multiples of core radius is less systematically influenced by the presence of the wing fin. The diffused wake structure with wing fin has a strong influence on the maximum induced rolling moments c1,max for a following aircraft (bf /b = 0.2). Figure 20 shows that a wing fin reduces the maximum induced rolling moment c1,max at x/b = 1.0 by almost 40%. The small difference in the maximum induced rolling moment for the wing fin with positive and negative angle of incidence suggests that the turbulence in the wake of the fin has a stronger influence on the induced rolling moment than the circulation generated by the fin. However, the y-, z-position of the maximum induced
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Figure 19. Circulation distribution Γ /bU∞ in wing tip vortex and outboard flap vortex.
Figure 20. Maximum induced rolling moment c1,max without and with wing fin.
rolling moments is little influenced by the wing fin. It remains near the wing tip vortex. The influence of wing fins on the induced rolling moment further downstream has to be investigated.
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5. Conclusion The experiments described in this paper deal with the influence of different flap settings on the wake structure of a flapped wing and the alleviating effect of wing fins on the induced rolling moments. Deploying the flap changes the wake structure from a one vortex system to a multiple vortices system. Three vortices are shed, one at the wing tip and one at each flap edge. For increasing flap angles the wing tip vortex remains approximately unchanged while the flap vortices are affected. Their core radius is reduced at lower flap angles followed by a widening at higher flap angles due to an increased flow separation at the flap edges. Vortex parameters such as core radius and maximum tangential velocity as well as vorticity and circulation distributions were investigated and presented. By means of a wing fin mounted near the outboard edge of the flap a significant decrease of the maximum induced rolling moment for a horizontally following wing could be achieved at one span downstream of the generator wing. This decrease can be attributed both to convectively and viscously dominated transport mechanisms. The amount of circulation within the core of the wing tip vortex is diminished. The outboard flap vortex is highly diffused and widened by the influence of the turbulent wake of the wing fin. Both effects lead to a changed velocity distribution which induces lower maximum rolling moments on the following aircraft.
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