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Publication numberWO1995011160 A1
Publication typeApplication
Application numberPCT/GB1994/002278
Publication date27 Apr 1995
Filing date18 Oct 1994
Priority date18 Oct 1993
Also published asEP0734342A1
Publication numberPCT/1994/2278, PCT/GB/1994/002278, PCT/GB/1994/02278, PCT/GB/94/002278, PCT/GB/94/02278, PCT/GB1994/002278, PCT/GB1994/02278, PCT/GB1994002278, PCT/GB199402278, PCT/GB94/002278, PCT/GB94/02278, PCT/GB94002278, PCT/GB9402278, WO 1995/011160 A1, WO 1995011160 A1, WO 1995011160A1, WO 9511160 A1, WO 9511160A1, WO-A1-1995011160, WO-A1-9511160, WO1995/011160A1, WO1995011160 A1, WO1995011160A1, WO9511160 A1, WO9511160A1
InventorsPatrick Ralph Ashill, Geoffrey Leonard Riddle
ApplicantThe Secretary Of State For Defence
Export CitationBiBTeX, EndNote, RefMan
External Links: Patentscope, Espacenet
Sub boundary layer vortex generators
WO 1995011160 A1
An aircraft wing possesses at least one vortex generator for controlling boundary layer separation, each vortex generator, which may be in the form of thin wires, being located entirely within the boundary layer such that the wing generates at least one dominant lee-side vortex. The vortex generators are preferably located on the leading edge of the wing, for example in a position upstream of where boundary layer separation would occur. The application describes vortex generators which are at an angle in the range of 12° to 15° to the leading edge of the wing.
Claims  (OCR text may contain errors)
1. An aircraft wing possessing at least one vortex generator for controlling boundary layer separation, characterised in that each vortex generator is located entirely within the boundary layer such that the wing generates at least one dominant lee-side vortex.
2. An aircraft wing according to Claim 1 in which each vortex generator is in the form of one or more thin wires.
3. An aircraft wing according to either preceding claim in which the vortex generators are located on the leading edge of the wing.
4. An aircraft wing according to Claim 3 in which the vortex generators are further located in a position upstream of where boundary layer separation would occur.
5. An aircraft wing according to any preceding claim in which the vortex generators are at an angle in the range of 12° to 15° to the leading edge of the wing.
Description  (OCR text may contain errors)


This invention relates to the use of novel sub boundary layer vortex generators to control boundary layer separation on the leading edge of wings.

Aircraft wings have thickness and camber distributions designed to ensure satisfactory supersonic and subsonic performance. At subsonic manoeuvre conditions there are likely to be large adverse pressure gradients in the flow near the wing leading edge which may cause boundary layer separation. Leading edge separation causes a loss in suction on forward facing surfaces of the wing leading to an increase in drag and also adverse effects on longitudinal and lateral stability. Vortex generators can alleviate this problem as they can control boundary layer separation, and thereby upper surface flows of wings. This confers much improved manoeuvre performance at subsonic speeds.

The use of conventional, vane-type vortex generators to control boundary layers on wings is known in the art. These generators are designed to protrude outside the boundary layer with the aim of transferring high energy air in the inviscid flow into the lower-energy flow within the boundary layer. The trailing or streamwise vorticity that promotes this mixing is generated in the same way that a wing generates trailing vorticity, ie by generating circulation around streamwise sections. However, since a large part of this type of device lies in the high-dynamic pressure flow outside the boundary layer, a significant drag penalty is caused.

This problem is reduced by the use of sub boundary-layer vortex generators, These devices are buried within the part of the boundary layer where the velocity deficit is greatest, typically within 20$ of the boundary layer thickness. Consequently, they have a much lower parasitic drag than vane vortex generators, but are still effective in transferring energy from the outer regions of the boundary layer to the low energy regions near the wing surface. Sub boundary-layer vortex generators are placed too deep within the boundary layer to create circulation effectively; instead, they provide streamwise vorticity by conversion of boundary-layer vorticity.

Sub boundary-layer devices use various ways to convert boundary-layer vorticity into streamwise vorticity. Sub boundary-layer devices of the Kuethe type comprise wave elements which divert the boundary-layer vorticity by means of wave elements. In devices of the Wheeler type, separation of the flow over an aft face alters the direction of the vorticity. Devices of these types are arranged so that they generate a vortex pair. It is advantageous instead to create a dominant lee side vortex which reduces the flow of the tired boundary layer air flowing towards the tip of the wing which prevents air flow parallel to the leading edge, which until now has not been achieved by the above prior art techniques.

By the method described above, it is an object of the invention to reduce the pressure drag of aircraft wings, without significantly increasing the resultant parasitic drag.

According to a first aspect of the invention there is provided a wing possessing one or more vortex generators used to control separation on wings wherein the or each vortex generator is located entirely within the boundary layer and such they produce one or more a dominant lee-side vortices.

The application of vortex generators in prior art applications has been to control of separation on wings of low sweep. A further advantage of the invention is that it is effective in controlling separations on highly-swept wings, to give significant reductions in drag of at subsonic manoeuvre conditions.

Further advantages of incorporating sub boundary-layer vortex generators according to the invention are described below.

The invention will now be described with reference to the following figures of which:

Figure 1 shows an embodiment of a sub boundary layer vortex generator in the form of a thin wire suitably located on the surface of the leading edge portion of an aircraft wing.

Figure 2 shows a further embodiment of the invention allowing for the location of multiple sub boundary layer wire vortex generators. Figure 3 shows sketches of upper surface oil-flow visualisations showing the types of flow over a) a basic wing and b) a wing with a wire vortex generator.

Figure 4 shows a more detailed sketch of the flow in the vicinity of the wire vortex generator.

Figure 5 shows the effect of a single wire used as a sub boundary layer vortex generator on vortex generator on the change on normal force, axial force and the axial position of "P".

Figure 6 shows the effect of single and multiple wires on the axial force coefficient at various angles of attack.

Figure 7 shows the effects of wire vortex generators on lift-dependant drag factor.

Figure 8 shows the effect of wire vortex generators on pitching moment-lift curve.

In order to furnish information on the way leading edge separation occurs on highly swept wings, experiments were conducted on wings using oil flow visualisation in a wind tunnel. Experiments were conducted on the port half of a wing-body configuration mounted on a mechanical balance measuring four components of force or moment. The wing had a leading-edge of 60° sweep and pronounced curvature in the spanwise direction.

An illustration of the character of the flow over the wing upper surface with¬ out a wire vortex generator is shown in figure 2a. It shows that the flow may be divided into two regions, one upstream of a point P on the leading edge, where the leading-edge flow is attached, and the other downstream of it where leading-edge separation occurs, revealing a flow with a part-span separation S2 reattaching at R, and a secondary separation of the flow towards the wing tip at S3- In the upstream region, separation occurs on the curved part of the upper surface at SI: this separation has a profound influence on the flow as a whole through its effect on the boundary-layer flow towards the point P. This separation is known to cause an increase in pressure drag over the lead¬ ing edge or, in other words, results in a reduction in leading edge thrust. Investigations were conducted into ways of controlling such flows by influencing the behaviour of the boundary layer in the region of the separation line SI, by the application of sub boundary layer vortex generators placed on the wing leading edge and to examine the subsequent effects on parameters such as axial and normal force coefficients.

Effect of single devices

In order to quantify the effects a wire vortex generator on parameters such as changes in axial and normal force acting on the wing, static pressures were measured on the wing surface at 7 stations along the model axis at a total of 300 orifices of 0.5mm diameter drilled normal to the surface. The axial position of each of these stations was defined by x, the axial distance from the apex of the wing, made non dimensional by length of centre-line chord, c . In order to determine the dimensional axial position of the point P where separation occurs, xp the pressure coefficient at an orifice close to that for minimum pressure was plotted against angle of incidence at each axial measurement station.

Fig la illustrates the geometry of a single wire vortex generator (WVG) . In the tests the wire was of diameter 0.51mm (0.0002δco) and of length 22.9 mm. The wire was stuck to the wing surface with a contact adhesive and was pressed onto the surface so that the height of the wire crest above the wing surface was equal to wire diameter. This was slightly larger than the calculated displacement thickness of the boundary layer just upstream of the wire and thus the device was kept well within the boundary layer. The position of the wire was defined by a particular axial station, xw which is a distance of 0.428co from the wing apex and which the nose of the wire is downstream of by a fixed distance 0.0l42co. The orientation of the wire in this figure shows that the angle between the axis of the wire and the wing leading edge in the wing chord plane was set at 16.3°.

Fig 2b shows the effect of a single wire vortex generator on the flow over the wing. A further illustration is provided by Figures 3a and 3b, which show a view of a surface oil flow and a sketched interpretation in the near region of a WVG respectively. One important feature of the flow that is readily evident is the "scarf" vortex wrapped around the nose of the wire. This vortex results from the conversion of boundary-layer vorticity into vorticity which is parallel to the streamwise direction. The mechanism for effecting this change is that of a separation of the boundary layer just upstream of the nose of the wire. The wire is at incidence to the local flow so that the vortex on the lee side of the wire is much stronger than that on the windward side, and it is the lee-side vortex that has the dominant effect on the flow further downstream. This vortex alters the flow beneath it both in the direction normal to and along the separation line further inboard. The vortex increases the velocity of the flow toward and just upstream of the separation line, while reducing the velocity of the flow along the separation line. The effect of this is to displace the separation downstream at a given spanwise position by a significant amount, typically up to about 1% of centre-line chord. This shift results, in turn, in a downstream displacement of the point P, with the consequence of increasing the extent of the leading edge over which the flow is attached causing a reduction in pressure drag or, in other words, an increase in leading-edge thrust.

The effect on axial-force coefficient which is related to pressure drag, normal-force coefficient and the axial position of the point P of a single wire at xw = 0.452co is shown against angle of incidence in figure 4. This shows that the effect of the single WVG is to reduce both normal-force and axial-force coefficients significantly over the range of angles of incidence 12° to 16.5° • Fig 4 also shows a correlation between the downstream shift of point P and the reductions in normal-force coefficient and axial force coefficient (or increase in leading-edge thrust coefficient) . These reductions in force coefficients are due to the WVG increasing the region of the wing over which the leading-edge flow is attached, decreasing pressure drag and thereby resulting in increasing the leading-edge thrust.

Effect of wire angle

The effect of wire angle (the angle between the leading edge and the wire axis) on increments in axial and normal force coefficients and the axial shift in the point P due to a wire at the same axial position was investigated . The strength of the dominant vortex depended on the angle between the wire axis and the local flow direction within the boundary layer. An angle of close to 16.3° was found to be optimum in terms of the magnitude of the increase in leading edge thrust, at least for the angles of incidence of 14° and 15° • As the wire angle increases at these angles of incidence, the reductions in the force coefficients decrease until, at about 50°, they have virtually vanished.

Effect of wire diameter

The influence of wire diameter on the increments in normal and axial-force coefficients was studied and revealed that there is a rapid reduction in the effect of WVG's as the wire diameter is reduced and no effect when submerged within the viscous sub layer.

Multiple wires

According to a preferred further embodiment of the invention there is provided a plurality of WVG's so as to further enhance the performance of aircraft wings. Experiments were conducted in assessing whether as the number of WVG's on the wing is increased, there is a further downstream shift of P resulting in a further increase in leading-edge thrust. Configurations with multiple wires were tested according to the three sets of configurations as described in Table 1 below and Figure lb.

Table 1: Numerical details of geometries of families of wires

Set Number X l X of Wires ( 1c o ) ' ( c )

1 2 0. 368 0.0843

2 3.7 0.284 0.0843

3 7. 14 xwl (n/2 ) = xωl + xw(n) xw (n/2 ) = 2 xw (n) 28 , 56

The position of the wires are defined by the notation given in the table above and according to the the non-dimensional distance between where the axis of the most upstream wire intersects the leading edge, and the wing apex, xwl. The wires do not extend completely to the leading edge, there being a distance of 0.0l4Co between the lee-side end and the intersect. The non-dimensional axial distance between successive wires is denoted, x W and the number of wires n. In case of Family 3. the largest number of wires tested was 5 , the corresponding value of xwl being shown in Table 1; the other members of the family were obtained by successively removing every other wire, beginning with the most upstream wire.

Figure 5 shows plots of the reduction in overall axial-force coefficient C and the shift in the axial position of "P" respectively, for single WVG's at xw = O.368 and 0.452 and for two devices, one at each of these positions against angle of attack. It is seen that the combined effects of two devices is greater than the sum of the individual contributions both in terms of the magnitude of the axial-force increment and the downstream shift of 'P' . At an angle of attack of 16.3°. a single device at x = 0.452, case A, has no effect on the flow, and the sketch at the bottom of the figure shows that this is because the device is downstream of the separation line SI. The single device at xw = O.368 (case B) , on the other hand, is upstream of the separation line and consequently moves this line and the point P downstream. This movement is sufficiently large to ensure that a second device placed at xw = 0.452 then becomes effective (case c) . This favourable interference increases the range of angle of incidence over which the devices are effective.

The variations with angle of incidence and of wire number on axial force increments are shown in Fig 6 for members of family 3- The effect of increasing the number of wires from 7 to 28 is to increase the range of angles of incidence over which the WVG.s are effective in increasing leading edge thrust. Furthermore, the change in this range of angle of incidence with wire number correlates well with the corresponding change in the range of angle of incidence for the downstream shift of the point P. A further increase to 56 wires reduces the range for which leading-edge thrust increases, and leads to an adverse effect, with 6 wires actually increasing the axial force for angles of incidence below 13°. As the number of wires increases between 7 and 28 , the reduction in axial force coefficient increases; further increase causes the benefit to decrease, indicating that 28 wires is the optimum for Family 3- Surface oil flows show that the vortices for the case with 56 wires interfere destructively explaining why there is a deterioration in performance with this number of wires. In summary therefore, an increase in the number of wires is shown to enlarge the range of angles of incidence over which the wires are effective in increasing leading-edge thrust.

Further advantages

Another advantage of the use of sub boundary layer vortex generators is that reductions in the lift dependent drag factor K =CDA/CL2 occur and are shown as a percentage of the lift-dependent drag factor of the datum (wire-off) configuration in Fig 7 • The maximum reduction occurs at a lift coefficient of about 0.6 for all wire numbers. A reduction in lift-dependent drag factor of up to about lβ% is shown to be achievable by using multiple wire vortex generators and calculations indicate that the overall parasitic drag penalty of the wire configuration giving the largest reduction in lift-dependent drag is insignificant in high-speed flight.

A further advantage of wire vortex generators delay to a higher lift coefficient the departure of the pitching moment-lift curve from a linear trend, as can be seen from figure 8. The reduced pitching moment coefficient cm, of the datum, no-wire case is seen to be close to zero between the lift coefficients 0.4 and 0.55 hut at higher lift coefficients, to increase rapidly with lift coefficient. The departure from linear law has an adverse influence on longitudinal stability and control. The effect of the WVG's is to increase the value of lift coefficient at which this increase in pitching moment occurs, and ensures that, over the range of lift coefficients where the WVG's produce the largest benefit in terms of reduced drag, the departures of pitching moment from a linear relationship are small. WVG's therefore result in enhanced stability and control of aircraft wings.

Patent Citations
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WO1984003867A1 *3 Apr 198411 Oct 1984Anders Edvard Hugo MalmstroemMethod and means for preventing wall turbulence
US4354648 *6 Feb 198019 Oct 1982Gates Learjet CorporationAirstream modification device for airfoils
Non-Patent Citations
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
WO2011113880A3 *16 Mar 20111 Dec 2011Airbus Operations GmbhFlexible sheet material for reducing the air drag of an aircraft
International ClassificationB64C21/10, B64C23/06
Cooperative ClassificationY02T50/162, B64C23/06, B64C21/10, Y02T50/166
European ClassificationB64C23/06, B64C21/10
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