QIN, Ning (University of Sheffield, Department of Mechnaical EngineeringMappin Street,Sheffield, South Yorkshire S1 3JD, GB)
| CLAIMS 1. A three-dimensional bump (100, 500) for location on a surface of an aerodynamic body, comprising: a first curved surface (120, 140, 160, 180, 200, 512) arranged to re-distribute pressure at a foot of a shock wave formed on the surface of the aerodynamic body; and at least one vortex generating edge (300, 550) for generating at least one vortex downstream thereof. 2. A device as claimed in claim 1 , wherein the vortex generating edge (300, 550) is curved in vertical and/or horizontal planes. 3. A device as claimed in claim 1 or 2, wherein the vortex generating edge (300) extends generally horizontally outward. 4. A device as claimed in claim 1 or 2, wherein the vortex generating edge (550) extends generally vertically upward. 5. A device as claimed in any preceding claim, wherein: the vortex generating edge comprises at least one projection (300, 550) extending from the first surface and arranged to generate the at least one vortex downstream thereof. 6. A device as claimed in claim 5, wherein the or each projection (300) comprises upper (320) and lower (340) surfaces joined at the vortex generating edge (300). 7. A device as claimed in claim 6, wherein the lower surface (340) of the or each projection (300) is curved concavely defining, at least in part, a hollow region (400) between the projection (300) and the surface of the aerodynamic body. 8. A device as claimed in claim 5, wherein the or each projection (55) comprises first and second side surfaces (554) joined at the vortex generating edge (550). 9. A device as claimed in claim 8, wherein the first and second side surfaces (554) upwardly incline toward the vortex generating edge (550). 10. A device as claimed in claim 8 or 9, wherein the vortex generating edge (550) comprises a forward facing portion (551 ) generally following a cubic curve and a rearward portion (552) generally following a cubic curve. 1 1 . A device as claimed in any preceding claim, wherein the first surface (120, 140, 160, 180, 200, 512) comprises curved forward (140) and rearward (160) facing surfaces, and curved left (180) and right (200) side surfaces. 12. A device as claimed in claim 1 1 , wherein the forward (140) and rearward (160) facing surfaces comprise, respectively, forward (140) and rearward (160) edges arranged to intersect with the surface of the aerodynamic body substantially tangentially. 13. A device as claimed in claim 1 1 or 12, wherein the forward (140) and rearward (160) facing surfaces define, in longitudinal cross section, first and second curves joined at a crest (220) extending transversely across the bump. 14. A device as claimed in claim 13, wherein the first and second curves are symmetrical or asymmetrical about the transverse crest (220). 15. A device as claimed in any of claims 1 1 to 14, the or each vortex generating edge (300) comprising an outermost portion of a respective one of the left and right side surfaces. 16. A device as claimed any preceding claim, wherein the first and/or second curves are cubic curves. 17. A device as claimed in claim 13 to 16, wherein the transverse crest (220) is positioned, in a longitudinal direction: closer to the forward (140) edge than to the rearward (160) edge; or closer to the (160) edge than to the forward (140) edge. 18. A device as claimed in any of claims 11 to 17, wherein the left and right side surfaces join at a crest (220) extending longitudinally along the bump between the forward (140) and rearward (160 edges. 19. A device as claimed in claim 18, wherein the longitudinal crest defines a centreline of the bump about which the bump is substantially symmetrical in cross section. 19. A device as claimed in any preceding claim, wherein the first surface is arranged to effect substantially isentropic compression on the shock wave to transform the shock wave into one selected from amongst a series of compression waves and a series of weaker shock waves. 20. A device as claimed in any preceding claim, wherein the at least one vortex generating edge (300) is arranged to generate a vortex in close proximity to the first surface so as to reduce and/or delay shock induced separation from a downstream portion of the first surface. 21 . A device as claimed in claim 13 or any claim dependent on claim 13, wherein the transverse crest (220) is positioned rearward of the forward (140) edge by a distance of between 40% and 70% of the overall length L of the device, preferably between 50% and 60%. 22. An aerodynamic body having at least one three-dimensional bump (100) as claimed in any preceding claim. 23. An aerodynamic body as claimed in claim 22, wherein the maximum height of the or each bump is between 0.5% and 2% of the chord length of the aerodynamic body. 24. An aerodynamic body as claimed in claim 23 or claim 24, wherein the length of the or each bump is between 20% and 35% of the chord length of the aerodynamic body. 25. An aerodynamic body as claimed in any of claims 22 to 24, wherein the or each bump is located at a position downstream of the shock wave location by a distance of between 2% and 9% of the chord length of the aerodynamic body. 26. An aerodynamic body as claimed in any of claims 2 to 25, wherein the or each bump is located at a position of between 60% and 65% of the chord length of the aerodynamic body from the leading edge thereof. 27. An aerodynamic body as claimed in any of claims 1 1 to 26, wherein the width of the or each bump is between 10% and 15% of the chord length of the aerodynamic body. 28. An aerodynamic body as claimed in any of claims 22 to 27, comprising an array of bumps extending in a span-wise direction of the aerodynamic body and spaced at a distance of between 20% and 40% of the chord length of the aerodynamic body between adjacent bumps. 29. An aerodynamic body as claimed in any one of claims 22 to 28 wherein the aerodynamic body is a wing. 30. An aerodynamic body as claimed in claim 29 wherein the wing is an aircraft wing. 31 . An aircraft having at least one bump as claimed in any of claims 1 to 21 or an aerodynamic body as claimed in any of claims 22 to 28. 32. A method of reducing drag on a transonic aerodynamic body, comprising: performing substantially isentropic compression on a shock wave formed on the aerodynamic body by means of a three-dimensional bump (100, 550) having a first curved surface (120, 140, 160, 180, 200, 512) arranged to re-distribute pressure at a foot of the shock wave; and generating, by at least one vortex generating edge (300, 550) of the three-dimensional bump (100, 550), at least one vortex so as to reduce and/or delay shock induced separation from a rear surface thereof. |
The present invention is concerned with aerodynamic drag reduction. In particular, but not exclusively, the invention relates to the reduction of drag experienced by aerodynamic surfaces, for example surfaces of aerodynamic bodies such as transonic wings. Such wings may be employed for example in the construction of aircraft. Aspects of the invention relate to a device, to an aerodynamic surface, to an aerodynamic body, to a wing, to an aircraft and to a method.
Drag is generally the most significant limiting factor in the aerodynamic performance and fuel efficiency of aircraft. Drag reduction is therefore an important aspect of aircraft design, not only for more cost effective operation through reduced fuel consumption but also to ensure adherence to the increasingly strict regulations for aircraft emissions.
The total drag experienced by aircraft is comprised of several components which vary in both magnitude and proportion throughout the operational envelope of the aircraft.
Lift induced drag, that is drag caused by the redirecting of the airflow by the wing in order to generate lift, is a major component of the total drag exerted on the aircraft. Form drag (also known as profile drag or pressure drag), that is drag caused by the viscosity of the air, is another prominent component of the total drag experienced by the aircraft.
One contributing factor to the increase in form drag on an aircraft is separation from the aircraft wing of the boundary layer, the layer of air closest to the surface of the wing, within which the stream-wise airflow relative to the aircraft is reduced gradually to zero at the surface. It is advantageous to maintain a laminar (non-turbulent) boundary layer across the surface of the wing since this results in lower skin friction, another component of the total drag experienced by the aircraft. Laminar boundary layers, however, contain less energy and are therefore more susceptible to the adverse (negative) pressure gradients that are formed on the upper surface of the wing in flight. Towards the rear part of the wing chord, increasing pressure causes the boundary layer to become detached from the wing surface in a process called boundary layer separation. l Separation of the boundary layer from the wing surface causes a significant increase in form drag experienced by the aircraft. In addition, early flow separation may reduce the effectiveness of the major flight control surfaces for the wing, which are generally located near the trailing edge, and is thus a significant limiting factor on aerodynamic performance. Early flow separation can be controlled by keeping the boundary layer attached to the wing surface for as long as possible.
At higher speeds, particularly transonic or supersonic speeds, aircraft may experience a third drag component known as wave drag (also termed shock drag). In transonic flight, that is flight at Mach numbers between approximately 0.6 and 1.0, wave drag occurs as the result of shock waves formed when localized regions of air on the upper surface of the wing reach supersonic velocities due to airflow acceleration over the curved surfaces of the wing. Shock waves are associated with significant increases in entropy and consume a considerable amount of energy which is experienced by the aircraft as drag and which increases as the strength of the shock wave increases.
Furthermore, in addition to the direct adverse effect of wave drag on the aircraft, the formation of a normal shock wave on the surface of the wing above the critical Mach number creates a strong adverse pressure gradient which tends to promote boundary layer separation downstream of the shock wave, thereby causing a further increase in form drag.
Although shock waves are typically associated with transonic/supersonic flows, they can form at much lower flight speeds at areas on the aircraft where local airflow accelerates to supersonic velocities. The lowest flight speed at which the localized airflow over the wings becomes supersonic is known as the critical Mach number of the aircraft which, in practice, occurs well below an airflow (flight) speed of Mach 1. Thus, aircraft flying at transonic speeds (for example Mach numbers of 0.7-0.9) routinely experience wave drag which increases significantly as the airflow speed increases.
Drag adversely affects aircraft operating speed and fuel economy. Various proposals have therefore been made to reduce the different forms of drag in an effort to improve aircraft fuel economy, range and emissions. For example, it is known that a turbulent boundary layer is less susceptible to flow separation than a laminar boundary layer and thus remains attached to the wing surface for longer. Although a turbulent boundary layer generates more skin friction than a laminar boundary layer, this may be insignificant compared to the increase in form drag generated by a separated laminar boundary layer.
Devices have been proposed to "energise" the boundary layer in order to delay or prevent separation to reduce form drag. One such form of device is a vortex generator, a generally triangular or rectangular vane positioned on the upper surface of the wing towards the leading edge. Usually having a height sufficient to protrude above the boundary layer, vortex generators generate tip vortices which draw in energetic, fast flowing air from outside the boundary layer into contact with the wing surface. By re-energising the boundary layer downstream of the vortex generator with this faster moving air, the airflow over the wing is less susceptible to separation and remains attached to the surface for longer, thereby reducing form drag.
Devices for reducing wave drag have also been proposed. For example, the use of two- dimensional "bumps" located on the upper surface of the wing has been proposed. Such bumps generally comprise a raised profile positioned at or close to the foot of the shock wave and extending along substantially the full span of the wing. The bump is arranged to cause substantially isentropic compression in which the normal shock wave is replaced with a sequence of compression waves with little or no increase in entropy, thereby weakening the shock wave and reducing wave drag. A recent advancement on the two-dimensional bump is the three-dimensional bump, that is a bump whose width (span-wise dimension) is significantly less than the span-wise dimension of the wing. Studies have shown that the provision of an array of such devices on the upper surface of the wing can achieve significant reductions in wave drag over a wider operational range compared to two-dimensional bumps, and are more suitable for use with the types of swept wing used on transonic commercial aircraft. Further details of such three-dimensional shock control bumps is disclosed in "Three-dimensional contour bumps for transonic wing drag reduction", Proc. IMechE Vol. 222 Part G: J. Aerospace Engineering, pp.619-629, the contents of which are incorporated herein by reference. However, while improvements in drag reduction remain an important area for development in the field of aircraft design, heretofore the reduction of multiple forms of drag in a single device has not been considered.
It is an aim of the present invention to address this problem and to improve upon known drag reducing devices and techniques. Embodiments of the invention may provide a device adapted for use on an aerodynamic body such as a wing, for example the upper surface of a wing , or on the suction side of a rotor blade, which reduces both shock drag and form drag in the transonic regime. Other aims and advantages of the invention will become apparent from the following description, claims and drawings.
Aspects of the invention therefore provide a device, an aerodynamic body, a wing, an aircraft and a method as claimed in the appended claims.
According to another aspect of the invention for which protection is sought, there is provided a device for use on a surface of an aerodynamic body, the apparatus comprising shock control means for reducing the strength of a shock wave formed on the surface of the aerodynamic body and vortex generation means for generating at least one vortex downstream of the shock control means.
In an embodiment, the device comprises a three-dimensional bump for location on the surface of the aerodynamic body. The shock control means may comprise a first, curved surface arranged to re-distribute the pressure at the foot of the shock wave. The vortex generation means may comprise at least one wing-like projection extending from the first surface and arranged to generate a vortex downstream thereof.
The first surface of the device provides the beneficial effect of reducing wave drag by replacing the normal shock wave formed on the upper surface of the aerodynamic body with a series of compression waves or weaker shocks. This reduces both the strength of the shock wave and the adverse pressure gradient on the surface of the body and thus reduces the risk of flow separation. Furthermore, the vortex generated by the at least one wing-like projection reduces trailing edge separation by generating a span-wise rotational flow in the airstream over the surface of the aerodynamic body, thereby reducing form drag, in the case of a wing, improving control surface effectiveness.
In addition, however, the applicant has found that the apparatus of the invention provides the unexpected technical effect of significantly reducing shock-induced separation from a downstream portion of the first surface.
The combined effect of the first and second surfaces exceeds that which would be expected through the use of separate, conventional three-dimensional bumps and vortex generators. Moreover, achieving these three beneficial effects (the reduction of wave drag, the reduction of trailing edge flow separation and the reduction of shock induced flow separation) in a single device reduces both cost and weight compared to existing solutions, and also improves design freedom. In an embodiment, the first surface comprises curved forward and rearward facing surfaces and curved left and right side surfaces. The or each wing-like projection may comprise an outermost portion of a respective one of the left and right side surfaces.
In an embodiment, the forward and rearward facing surfaces comprise, respectively, forward and rearward edges arranged to intersect with the surface of the aerodynamic body substantially tangentially.
In an embodiment, the forward and rearward facing surfaces define, in longitudinal cross section, first and second cubic curves joined at a crest extending transversely across the bump.
In an embodiment, the first and second cubic curves are symmetrical about the transverse crest.
Alternatively, the first and second cubic curves may be asymmetrical about the transverse crest such that the transverse crest is positioned, in a longitudinal direction, either closer to the forward edge than to the rearward edge or closer to the rearward edge than to the forward edge. In an embodiment, the or each wing-like projection comprises upper and lower surfaces joined at a vortex generating edge.
In an embodiment, the vortex generating edge is curved in both vertical and horizontal planes. The height of the edge from the surface of the aerodynamic body may therefore vary along the length of the device. Alternatively, or in addition, the span-wise width of the or each wing-like projection may vary along the length of the device. The positions of maximum height of the edge from the surface of the aerodynamic body and maximum width of the or each wing-like projection may be mutually coincident and/or coincident with the transverse crest.
In an embodiment, the lower surface of the or each wing-like projection is curved concavely defining, at least in part, a hollow region between the wing-like projection and the surface of the aerodynamic body. In an embodiment, the left and right side surfaces join at a crest extending longitudinally along the bump between the forward and rearward edges. The longitudinal crest may define a centreline of the bump about which the bump is substantially symmetrical in cross section.
In an embodiment, the first surface is arranged to effect substantially isentropic compression on the shock wave to transform the shock wave into a series of compression waves or weaker shocks.
In an embodiment, the at least one wing-like projection is arranged to generate a vortex in close proximity to the first surface so as to reduce and/or delay shock induced separation from a downstream portion of the first surface.
In an embodiment, the transverse crest is positioned rearward of the forward edge by a distance of between 40% and 70% of the overall length L of the device, preferably between 50% and 60. According to another aspect of the invention for which protection is sought, there is provided an aerodynamic body having at least one device as set out in any of the preceding paragraphs.
The aerodynamic body may be a wing such as a wing of an aircraft. In an embodiment, the maximum height of the or each device is between 0.5% and 2% of the chord length of the aerodynamic body. In an embodiment, the length of the or each device is between 20% and 35% of the chord length of the aerodynamic body.
The or each device may be located at a position downstream of the shock wave location by a distance of between 2% and 9% of the chord length. In an embodiment, the or each device is located at a position of between 60% and 65% of the chord length from the leading edge thereof. In an embodiment, the width of the or each device is between 10% and 15% of the chord length.
The aerodynamic body may comprise an array of devices extending in a span-wise direction of the aerodynamic body with a spacing between adjacent devices of between 20% and 40% of the chord length.
According to a further aspect of the invention for which protection is sought, there is provided a wing having a device as set out previously.
According to a still further aspect of the invention for which protection is sought, there is provided a method of reducing drag on a transonic aerodynamic body, comprising performing substantially isentropic compression on a shock wave formed on the aerodynamic body by means of a three-dimensional bump and generating at least one span-wise vortex adjacent or in close proximity to a rear surface of the bump such so as to energise the boundary layer on the rear surface of the bump and reduce and/or delay shock induced separation therefrom.
Within the scope of this application it is envisaged that the various aspects, embodiments, examples, features and alternatives set out in the preceding paragraphs, in the claims and/or in the following description and drawings may be taken independently or in any combination thereof.
The present invention will now be described, by way of example only, with reference to the accompanying drawings, in which: Figure 1 is a perspective view of one half of a known three-dimensional shock control bump; Figure 2 is a full cross-section through the shock control bump of Figure 1 in the y-z plane; Figure 3 is a perspective view of one half of an apparatus embodying a form of the invention; Figure 4 is a plan view of the apparatus of Figure 3;
Figure 5 is a cross-section through the apparatus of Figure 3 in the y-z plane;
Figure 6 is a side view of the apparatus of Figure 3;
Figure 7 is a cross-section through the apparatus of Figure 3 in the y-z plane illustrating a comparison with the shock control bump of Figure 1 ;
Figure 8 is a full plan view of the apparatus of Figure 3;
Figure 9 is a full cross-section through the apparatus of Figure 8 in the y-z plane along the line M-M';
Figure 10 is a plan view of an apparatus according to a further embodiment of the invention;
Figure 1 1 is a schematic illustration of the apparatus of Figure 10 from three views; and Figure 12 is a perspective view of the apparatus shown in Figures 10 and 11.
It is to be noted that the accompanying Figures are for illustrative purposes only and are not drawn to scale. Reference will be made herein to an aerodynamic body, a wing and an aerofoil. For the avoidance of doubt, reference to aerodynamic body includes reference to a wing as an example of a 3D aerodynamic surface. The term aerodynamic body also includes reference to rotor blades, for example rotor blades of a helicopter, to propeller blades, gas turbine rotor blades, engine nacelles and any other suitable aerodynamic body.
Furthermore, in the following description, all directional, dimensional and geometrical terms are used in the context of a general aerodynamic body, such as an aircraft wing. Such bodies generally comprise upper and lower surfaces, a span (i.e. a dimension from root to tip) and a chord (i.e. a dimension between the leading and trailing edges thereof). The terms "upper", "lower", "span-wise" and "chord-wise" are to be interpreted accordingly. For convenience, reference will also be made to Cartesian axes, the orientation of which is shown in the Figures. In addition, as far as possible, like reference numerals indicate like parts.
Referring to Figure 1 , the right hand half of a known form of shock control device is shown, in perspective view, generally at 10. The device 10 comprises a raised profile, three-dimensional "bump" which is arranged to be positioned on the upper surface S of an aerodynamic body, such as the wing of an aircraft. Whilst the following discussion will be made with reference to positioning of the device 10 on the upper surface S of a wing, it is to be understood that the discussion is more widely applicable to any suitable aerodynamic body and not just to a wing.
The bump 10 comprises a smoothly curved upper surface 12 consisting of forward facing and rearward facing surfaces 14, 16, and left and right side surfaces 18, 20 (shown in Figure 2). The forward and rearward facing surfaces 14, 16 intersect with the upper surface S of the wing to define, respectively, forward and rearward edges 14a, 16a of the bump 10 and are joined at a transverse crest 22 extending in the y-z plane. Likewise the left and right side surfaces 18, 20 intersect with the upper surface S of the wing to define, respectively, left and right edges 18a, 20a of the bump 10, and are joined at a longitudinal crest 24 extending in the x-y plane.
As best shown in Figure 2, which illustrates a cross section through the bump 10 in the y-z plane along the transverse crest 22, the left and right side surfaces 18, 20 are substantially symmetrical about the longitudinal crest 24 which thus defines the transverse centreline of the bump 10. In other words, the left and right halves of the bump 10 are substantially symmetrically identical.
On the other hand, as best illustrated in Figure 1 , the forward and rearward facing surfaces 14, 16 are asymmetrical about the transverse crest 22 which lies, in the longitudinal direction (x direction), closer to one of the forward and rearward edges 14a, 16a than to the other. More specifically, in the illustrated embodiment, the transverse crest 22 is positioned closer to the forward edge 14a than to the rearward edge 16a. The average gradient of the forward facing surface 14, that is to say the angle of inclination of the forward facing surface 14 relative to the upper surface S of the wing, is therefore greater than the average gradient of the rearward facing surface 16.
In other forms of shock control bump, however, the transverse crest is positioned closer to the rearward edge 16a than to the forward edge 14a, such that the average gradient of the rearward facing surface 16 is greater than that of the forward facing surface 14.
As shown in Figure 2, the side surfaces 18, 20 blend into the upper surface S of the wing at the left and right edges 18a, 20a substantially tangentially. Likewise, the forward and rearward facing surfaces 14, 16 blend into the upper surface S of the wing at the forward and rearward edges 14a, 16a of the bump 10 substantially tangentially. It is widely recognised in the art that this tangential geometry is important to maintain a laminar flow at the intersections between the bump 10 and the upper surface S of the wing in order to limit increases in profile drag caused by the presence of the bump 10. It will be appreciated from the foregoing that the bump 10 has the effect of modifying the surface contour or geometrical profile of the wing, the effect of which is described below.
The parameterisation of the bump geometry is of great importance for optimising the aerodynamic performance of the bump 10. In particular, there are six parameters which are considered to contribute significantly to the drag-reducing effect of the bump 10 on the wing. These are: i) the maximum height H of the bump 10, i.e. the amount by which the bump projects above the upper surface S of the wing - it will be understood that the maximum height H of the bump 10 occurs at the intersection between the transverse crest 22 and the longitudinal crest 24, hereafter termed the peak point P. ii) the absolute position B A of the peak point P along the chord of the aerofoil; iii) the position B R of the peak point P relative to the forward and rearward edges 14a, 16a of the bump 10; iv) the overall length L of the bump 10, i.e. the distance between the forward and rearward edges 14a, 16a; v) the width W of the bump 10, i.e. the span-wise distance between the left and right edges 18a, 20a; and vi) the span-wise spacing Y between adjacent bumps 10 on the upper surface S of the wing.
Referring next to Figures 3 to 9, the right hand half of an apparatus embodying one form of the invention is shown generally at 100. The apparatus 100 takes the form of a three-dimensional shock control bump, similar in geometry to the bump 100 described above, in which one or both of the side surfaces 180, 200 is modified to provide the additional function of a vortex generator, as described below.
The apparatus 100, which is referred to hereafter by the term Shock Control Vortex Generating Bump (SCVGB), thus comprises shock control means in the form of a shock reducing portion A for reducing wave drag and vortex generation means in the form of two vortex generating portions B for reducing profile drag. In the illustrated embodiment, though not essentially, the shock reducing portion A and the vortex generating portions B are integrally formed but, for convenience and clarity, are described separately below.
The shock reducing portion A of the bump 100 comprises a smoothly curved upper surface 120 having forward facing and rearward facing surfaces 140, 160 and left and right side surfaces 180, 200. The forward and rearward facing surfaces 140, 160 intersect with the upper surface S of the aerofoil to define, respectively, forward and rearward edges 140a, 160a of the SCVGB 100 and are joined at a transverse crest 220 extending in the y-z plane.
The left and right side surfaces 180, 200 are curved substantially symmetrical about the x-y plane and are joined at a longitudinal crest 240 which thus defines the transverse centreline of the bump 100. In other words, the left and right halves of the bump 100 are substantially symmetrically identical.
As best shown in Figure 6, in this non-limiting embodiment, the geometry of the SCVGB 100 in the x-y plane is formed by the blending of two non-symmetrical cubic curves which join at the transverse crest 220. Each cubic curve is formed such that the gradient of the surface defined thereby (the angle of inclination relative to the upper surface S of the aerofoil) varies in the stream-wise direction from a minimum (substantially zero) close to surface S of the aerofoil, through a maximum, to a second minimum (substantially zero) at the transverse crest 220.
Figure 6 illustrates how the forward and rearward facing surfaces 140, 160 blend into the upper surface S of the aerofoil at the forward and rearward edges 140a, 160a of the SCVGB 100 substantially tangentially. This geometry is important to maintain a laminar flow at the intersections between the bump 100 and the upper surface S of the wing in order to limit increases in profile drag caused by the presence of the SCVGB 100 on the wing surface.
In the illustrated embodiment, the transverse crest 220 is positioned closer to the forward edge 14a than to the rearward edge 160a. The cubic curve forming the forward facing surface 140 of the shock reducing portion A therefore has an average gradient greater than that of the cubic curve forming the rearward facing surface 16 of the shock reducing portion A.
It will be appreciated, however, that this particular geometrical configuration is not essential and that variations, including embodiments in which the transverse crest 220 is positioned closer to the rearward edge 160a of the SCVGB 100 than to the forward edge 140a, such that the average gradient of the rearward facing surface 160 is greater than that of the forward facing surface 140, may equally be employed. The skilled person will readily understand the aerodynamic implications of varying the gradients of the forward and rearward facing surfaces 140, 160 of the SCVGB 100 and further discussion is therefore considered unnecessary. However, the position Br of the transverse crest 220 relative to the overall length of the SCVGB 100, and therefore relative to the forward and rearward edges 140a, 160b, termed the "relative transverse crest position", is an important parameter in determining the aerodynamic performance of the apparatus. This parameter can be optimized according to the desired application, for example the type and operating conditions of the wing with which the apparatus is to be used and the skilled person will be well able to perform such parameterisation in order to achieve optimum performance without further instruction. As mentioned above, the SCVGB 100 also comprises a vortex generating portion B. In the illustrated embodiment, the vortex generating portion B comprises two substantially symmetrically identical vortex generators positioned either side, in a span-wise direction, of the shock reducing portion A. In particular, as best shown in Figures 5 and 7, each vortex generator comprises a vortex- generating edge in the form of a respective wing-like projection 300 extending outwardly from the shock reducing portion A in a span-wise direction. Each wing-like projection 300 has an upper surface 320 and a lower surface 340. The upper surface 320 of the wing-like projection 300 is substantially contiguous with, or comprises a portion of, the respective side surface 180, 200 and is thus curved both in the y-z plane (i.e. about the longitudinal crest 240) and the x-y plane (i.e. about the transverse crest 220), as best shown in Figures 5 and 6. It will be appreciated, however, that the upper surface 340 may instead be substantially planar in the y-z plane.
The lower surface 340 of each wing-like projection 300, on the other hand, is curved concavely between the upper surface 320 and a respective side edge 180a, 200a, at which the lower surface 340 intersects with the upper surface S of the wing, defining a hollow region 400 therebetween.
The upper and lower surfaces 320, 340 of the wing-like projection 300 meet at a relatively sharp edge or chine 360 which extends substantially along the full length of the SCVGB 100, intersecting with the upper surface S of the wing at or adjacent the forward and rearward edges 140a, 160a, respectively.
As shown in Figure 6, the edge 360 is curved in the x-y plane such that its height V above the upper surface S of the wing varies along the length L of the SCVGB 100 from a minimum at the forward and rearward edges 140a, 160a, to a maximum at a point 420 therebetween. When viewed in plan form, as shown in Figure 4, it can be seen that the edge 360 is also curved in the x-z plane such that the width (span-wise dimension) of the wing-like projection 300 also varies along the length L of the SCVGB 100 from a minimum at the forward and rearward edges 140a, 160a, to a maximum at a point 440 therebetween.
In the illustrated embodiment, the positions 420, 440 of maximum height of the edge 360 and maximum width of the wing-like projection 300 are substantially coincident in a chord-wise (longitudinal) direction. In addition, the positions 420, 440 may be substantially coincident with the position, in a chord-wise (longitudinal) direction, of the transverse crest 220. That is to say, the positions 420, 440 of maximum height of the edge 360 and maximum width of the SCVGB 100 may be located closer to the forward edge 140a than to the rearward edge 160a in the chord-wise direction. It will be appreciated, however, that the above-described geometrical configuration of the winglike portion 300 is not essential and that the variances in the width of the wing-like projection 300 and the height of the edge 360 can be such that the points 420, 440 of maximum height and maximum width are located closer to the rearward edge 160a. Alternatively, the height V of the edge 360 and the width of the wing-like projection 300 may be substantially constant along the length L of the SCVGB 100.
The overall width W of the SCVGB 100, i.e. the combined width of the shock reducing portion A and the vortex generating portion(s) B, is another important parameter affecting the performance of the device and, again, this can be optimized for best performance for particular operating conditions of a given wing.
Advantageously, the shock reducing portion A and the vortex generating portion(s) B may be integrally and/or unitarily formed. Figure 7 illustrates an advantageous, but non-limiting example in which each wing-like projection 300 is formed as an integral part of the SCVGB 100 by excising an outer portion or volume (shaded region O) of the respective left and right side surfaces 180, 200 to form the hollow region 400. This technique for forming the vortex generating portions reduces both weight and part count, thereby also reducing cost and manufacturing complexity. Alternatively, it is envisaged that the shock reducing portion A and the wing-like portions 300 may be formed as separate components for assembly either on or prior to installation on the wing.
Figures 8 and 9 illustrate, respectively, a complete SCVGB 100 in plan view and in cross section in the y-z plane along the line M-M'. It can be seen that the SCVGB 100 is somewhat mushroom-shaped in cross section, being substantially symmetrical about the longitudinal crest (transverse centreline) 240, but is asymmetrical about the transverse crest 220. Nevertheless, as described above, it is equally possible for the SCVGB 100 to be substantially symmetrical about the transverse crest 220, for example by positioning the transverse crest 220 at the midpoint between the forward and rearward edges 140a, 160a, and asymmetrical about the transverse centreline 240, for example by providing a single vortex generator (wing-like projection 300) on only one side of the SCVGB 100.
Use and operation of the apparatus 100 will now be described.
In use, one or more SCVGBs 100 are positioned on an aerodynamic surface, such as an aircraft wing. As described previously, the optimal number and position of SCVGBs 100 on the wing, as well as the various dimensions and geometrical parameters thereof, will vary according to application, that is to say according to the properties of the wing and the operating conditions thereof.
By way of example only, however, the following specific values were found to be advantageous for a constant section natural laminar flow (NLF) wing (RAE5243) operating in flow conditions wherein Mach no. M = 0.68, Reynolds No. Re = 1.9 x 107, and Lift Coefficient CL = 0.82:
The overall length L of the SCVGB 100 from the forward edge 140a to the rearward edge 160a thereof is between 20% and 35% of the chord length, preferably between 30% and 34%, more preferably approximately 33%. A potentially optimal value was found to be 32.9% of the wing chord.
The transverse crest 220 is positioned on the upper surface S of the wing at a point downstream of the normal shock wave by a distance of between 2% and 9% of the wing chord. For example, the position BA of the transverse crest 220 relative to the wing may be between 60% and 65% of the chord from the leading edge of the aerofoil, preferably between 62% and 64%, more preferably approximately 63%. A potentially optimal value was found to be 62.7% downstream of the leading edge of the aerofoil.
The transverse crest 220 relative to is positioned rearward of the forward edge 140a by a distance of between 40% and 70% of the overall length L of the SCVGB 100, preferably between 50% and 60%, more preferably approximately 54%. A potentially optimal value was found to be 54.1 % rearward of the forward edge 140a. That is to say, as shown in Figure 6, the position of the transverse crest 220 is slightly rearward of the midpoint between the forward and rearward edges 14a, 16a of the SCVGB 100. As discussed above, the optimal value will be typically be a function of airflow conditions and the shape of the wing with which the SCVGB 100 is provided.
The maximum height H of the SCVGB 100 is between 0.5% and 2% of the chord length, preferably between 1 % and 1.5%, more preferably between 1.3% and 1.4%. A potentially optimal value was found to be 1.38% of the wing chord. The maximum width W of the SCVGB 100 is between 10% and 15% of the chord length, preferably between 1 1 % and 13%, more preferably approximately 12%. A potentially optimal value was found to be 1 1.9% of the wing chord.
Spacing Y between adjacent SCVGBs 100 on the wing surface S is between 20% and 40% of chord length, preferably between 23% and 25%, more preferably approximately 24%. It will be appreciated that this parameter determines the number of bump devices that are required on a given wing for optimum performance.
In use, the shock reducing portion A of the SCVGB 100 has the effect of redistributing the pressure near the foot of the normal shock wave through isentropic compression, thereby reducing its strength. In particular, the smoothly curved contour of the upper surface 12 has a "smearing" effect on the shock wave, spreading it into a series of more gradual compression waves or weaker shocks having an overall smaller entropy increase. The reduction of the shock strength directly reduces the wave drag exerted on the wing.
Furthermore, the increased pressure distribution reduces the post-shock adverse pressure gradient on the wing surface S which results in the boundary layer downstream of the SCVGB 100 being less likely to separate from the wing surface S. Delaying boundary layer separation aft of the SCVGB 100 advantageously reduces form or profile drag.
The vortex generating portions B enhance the drag reducing effects of the SCVGB 100 by generating a pair of strong tip vortices downstream of the shock-boundary layer interaction region. Each vortex, generated by the edge 360 of a respective one of the wing-like projections 300, draws energetic, rapidly-moving air from outside the boundary layer into contact with the surface S of the wing, thereby re-energising the boundary layer and avoiding streamwise separation immediately downstream of the SCVGB 100. By re-energising the boundary layer downstream of the vortex generator with this faster moving air, the airflow over the wing surface S is less susceptible to separation and remains attached to the surface for longer, thereby delaying trailing edge separation. This delay in trailing edge separation not only reduces profile drag but also improves flight control surface effectiveness. In addition, however, the applicants have found a further, unexpected technical benefit provided by the present invention. In particular, the applicants have recognised that shock induced flow separation on the rearward facing surface of a conventional shock control bump, such as that described with reference to Figure 1 , can degrade the bump's performance. However, such shock induced separation is difficult to prevent.
Whilst it would be preferable to use gradually sloping forward and rearward facing surfaces 140, 160 on the bump 100 to reduce downstream flow separation from the rearward facing surface 16, the length L of the bump would have to be increased significantly in order to achieve an optimal maximum height H of the bump. In practice, this would result in the length L of the bump 10 approaching 100% of the chord length, disadvantageously increasing surface friction, cost and weight.
On the other hand, reducing the overall length L of the shock control bump 100 reduces cost and weight, but requires more steeply angled forward and rearward facing surfaces 140, 160 in order to achieve the optimal maximum height H for the bump. The increased gradient of the forward and rearward facing surfaces 140, 160 tends to increase flow separation from the rearward facing surface 160.
Figures 10-12 illustrate an apparatus 500 of SCVGB according to a further embodiment of the invention. The apparatus 500 includes a shock reducing portion A and a vortex generating portion B. The shock reducing portion comprises a bump 510 and the vortex generating portion B comprises one or more vortex generating edges 550, as in the previously described embodiments. Figure 10 illustrates the apparatus 500 in plan view, Figure 1 1 illustrates the apparatus 500 from three different views and Figure 12 illustrates the apparatus 500 in a perspective view from an upstream direction showing only one vortex generating edge 550. i.e. the left hand edge of the diagram corresponds to the longitudinal axis 570. The bump 510 and one or more vortex generating edges 550 are integrally formed. In the embodiment shown in Figures 10-12 the vortex generating edges 550 protrude generally upward from an upper curved surface 512 of the bump 510. The embodiment shown in Figures 10-12 comprises two vortex generating edges 550, wherein each edge is offset i.e. spaced apart in a width-wise direction from the longitudinal axis of the bump 550 by a distance V Z i . The upper curved surface 512 of the bump 510 may be as previously described with reference to the embodiment shown in Figures 3-9 and repetition of those features will be omitted here for clarity. The bump 510 has a width W and a length L as indicated in Figure 1 1. A transverse crest of the bump 510 is positioned at a distance B R from a forward edge of the bump 510, and a distance B A from a forward edge of the aerofoil, as previously described.
Each vortex generating edge 550 is formed by a first, forward facing, cubic curve 551 which protrudes or extends upward from the surface 512 of the bump 510 and varies in the stream- wise direction 560 from a minimum (substantially zero) close to the surface 512 of the bump, through a maximum, to a second minimum (substantially zero) at a transverse crest 552 of the vortex generating edge 550. The transverse crest 552 of the edge 550 is spaced a distance in from the front edge of the bump 510 by a distance V R and a distance V A from the forward edge of the wing. A second, rearward facing, cubic curve 553 blends with the first cubic curve to meet the surface of the bump 512. A length V L of the vortex generating edge 550 from where the first cubic surface gradually blends upward from the surface 512 of the bump to where the second cubic surface blends into the surface 512 of the bump 510 is shown in Figure 1 1. The point at which the second cubic surface meets the surface 512 of the bump 510 is spaced apart from the centre of the bump 510 by a distance Vz2 indicated in Figure 11.
Side surfaces 554 of the vortex generating edges 550 extend upward from the surface of the bump to meet the first and second cubic curves at a longitudinally extending crest or knife edge. The side surfaces 554 may extend generally inwardly from the surface 512 of the bump, wherein the side surfaces are spaced apart at their base by a width of the vortex generating edge V w (shown in Figure 1 1 ). A height V of the edge 550 is shown in Figure 1 1 and is less, that is the edge is lower than, a peak height of the bump 510. The peak height H at the crest of the bump is also illustrated in Figure 1 1.
Each vortex generating edge 510 has a longitudinal axis 552. The longitudinal axis 570 of the bump 510 and the longitudinal axis of each vortex generating edge 550 may intersect at an angle Θ which may be in the range of 1 to 30° , 1 to 20°, or 1 to 10°. In some embodiments, the angle Θ may be around 5° It will be realised that the angle Θ may be selected according to various design criteria. The angular position of the vortex generating edge should be chosen to create a sufficient vortex 590 downstream thereof without causing excessive drag.
The applicants have discovered that the combining of the shock reducing portion A and the vortex generating portions B in a single device permits the span-wise vortices formed by the vortex generating portions B to be generated in sufficiently close proximity to the rearward facing surface 16 to have a significant beneficial effect on shock induced separation.
In particular, the increased energy of the boundary layer caused by the vortices is more resistant to the post-shock negative pressure gradient rearward of the SCVGB 100, 500 which, although considerably reduced by the shock control portion A, remains relatively high. Boundary layer separation on the rearward facing surface 16 of the SCVGB 100, 500 is thus avoided further reducing form drag.
Analysis has shown that, depending on the original shock wave strength, a total drag reduction of between 10 to 60 drag counts can be achieved by the SCVGB 100, 500 for M max between 1.2 to 1.5. This result is considerably better than would be expected from the use of separate vortex generators which are incapable of influencing shock induced separation from the rearward facing surface 160 of a conventional shock control bump. The applicants have therefore found that the fuel economy, range and emissions of an aircraft fitted with an array of SCVGBs 100, 500 are improved by an amount which exceeds that obtained by conventional means.
It will be appreciated by those skilled in the art that all of the above-described design parameters, for example bump height H, length L and width W, bump spacing Y, and relative and absolute crest positions B A , B R can be modified and optimised for best performance for particular flight conditions of a given aircraft wing.
The SCVGB 100, 500 may be formed of any suitable material, such as metal or composite, and may be solid or substantially hollow. One advantageous form of SCVGB 100, 500 is blow- moulded or injection moulded from a material and attached to the surface of the wing by any suitable means such as rivets or adhesive. Alternatively, the SCVGB 100, 500 may be formed integrally with the wing. Nothing in the above description is intended to limit the invention in any way, the scope of which is defined by the appended claims.
Throughout the description and claims of this specification, the words "comprise" and "contain" and variations of them mean "including but not limited to", and they are not intended to (and do not) exclude other moieties, additives, components, integers or steps. Throughout the description and claims of this specification, the singular encompasses the plural unless the context other-wise requires. In particular, where the indefinite article is used, the specification is to be understood as contemplating plurality as well as singularity, unless the context requires other-wise.
All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive. The invention is not restricted to the details of any foregoing embodiments. The invention extends to any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed.