GB2072600A

GB2072600A - Supercritical aerofoil section

Info

Publication number: GB2072600A
Application number: GB8008632A
Authority: GB
Original assignee: UK Secretary of State for Defence
Current assignee: UK Secretary of State for Defence
Priority date: 1980-03-13
Filing date: 1980-03-13
Publication date: 1981-10-07
Also published as: WO1981002557A1; GB2072600B; EP0047319A1

Abstract

An aerofoil for use on an aircraft whereover the section design mach number is 0.7-0.77 and having a thickness chord ratio of about 0.14, with t/c max forward of mid-chord, upper surface minimum curvature at or forward of max t/c, and leading edge radius and camber such that local peak M is obtained forward of 10% chord and supersonic flow maintained to aft of 40% chord, when it may terminate in a weak shock. The lower surface has rear camber and is arranged to maximise lift but with a distribution such as to alternate the pitching moment coefficient of the section as a whole.

Description

SPECIFICATION Improvements in aerofoils The present invention relates to supercritical aero foils, that is aerofoils which at their design speed maintain a supersonic flow over a region thereof, the supersonic flow being shock free or terminating in a weak shock and whereby an appreciable lift contri bution is obtained from that region without delete rious drag penalty.

It provides an aerofoil for particular application to transport aircraft, required to cruise at high subsonic Mach numbers.

According to the present invention an aerofoil suitable for use in the mainplane of a transport aircraft with a section design Mach number between 0.7 and 0.77 has a supercritical section with rear loading, the section having a thickness to chord ratio greater than 0.12, a leading edge region radius and camber such that there is generated atthe design Mach No and lift coefficient a local peak Mach No of 1.1 to 1.3 forward of 10% chord, and an upper surface curvature decreasing to a minimum of value and location in combination with the leading edge region radius and camber such that aft of the peak local Mach No supersonic flow is maintained with a net recompression which if terminating in a shock does so with a Mach No. thereat at least 0.05 less than the forward local peak Mach No., subsonic flow not being achieved till aft of 40% chord, and the lower surface contour being such that lift generated by the lower surface is maximised without excessive increase in viscous drag and is consistent with attenuating the pitching moment coefficient of the section as a whole. Preferably the commencement of the adverse pressure gradient associated with re compression to subsonic flow is as far aft as possible before the onset of excessive viscous drag.

According to feature of the invention the upper surface minimum curvature point may occur at 30 to 40% chord, and the maximum thickness at 35 to 45% chord. Preferably also, upper surface curvature rises from the minimum to a maximum at about 75% chord and then diminishes steadily toward the trailing edge. The minimum curvature value may be 2-2.5 t/c2, where t = thickness and c = chord, so that for a preferred section with a thickness to chord ratio of 0.14 the minimum curvature is 0.28 -0.35 for unit chord. The maximum value at 75% chord may be 4tIc2 or, with a 0.14 t/c section, 0.56 per unit chord length.

Sections with a thickness/chord ratio as high as those in accordance with the present invention, might be expected to incur somewhat of a drag penalty. However part of the discovery enshrined in the present invention is that by maximising the rear loading within the limitation imposed by the need to avoid unacceptable trailing edge separation, thus reducing that proportion of the total lift coefficient generated in the supercritical region, the wave drag of the section can be minimised.

A suitable leading edge radius is 1.4% chord and the slope of the camber line at the leading edge may be 6" to the chord line, nose up. This ensures weak shock waves or low wave drag over the aerofoil and no significant incidence variation over-sensitivity.

Nevertheless, in order to render sections according to the present invention compatible with requirements for satisfactory performance at other parts of the flight envelope, the section is preferably provided with variable geometry at the leading edge, obtained for example with a flexible nose flap of the kind described in UK Patent Specification 1296994, a leading edge slat or a droop flap.

The trailing edge may comprise flaps or control surfaces as usually found on aircraft aerofoils.

A family of aerofoils A to G in accordance with the invention will now be described by way of example, together with details of the behaviour of some of them, with reference to the accompanying drawings, of which Figure 1 is an outline of sections A and B Figure 2 is a table of the ordinates of Section A, Figure 3 is a table of the ordinates of Section B, Figure 4 is a table of the ordinates of the rear portions of Sections C to G Figure 5 is a graph of upper surface curvature distribution of Sections A and B Figure 6 is a graph of pressure distributions over Section B Figure 7is a graph comparing drag properties of Sections B and C As can be seen in Figure 1 Sections A and B are supercritical sections with camber extending over approximately the latter 40% chord.By virtue of the fact that Zu (the height of the upper surface above the chord) is lower up to 10% chord than Z1 (the height of the chord above the lower surface) the two sections have a nose up appearance. In fact, the slope of the leading edge is 6" to the chord line. The leading edge curvature is 1.4% chord. Compared with Section A, Section B mainly has a somewhat extended roof top and increased upper surface curvature at the trailing edge.

As shewn in Figure 5 the upper surface curvature of the two sections decreases to a minimum at 35% chord for Section A and 39% chord for Section B. The curvature values being 0.290 and 0.345 for unit chord length respectively. The curvature then increases to a maximum at 76% for Section A and 77% chord for Section B, the values being 0.546 for unit length chord in both cases, after which the curvature then falls again towards the trailing edge.

From the ordinates of the two sections listed in Figures 2 and 3 it can be seen that the maximum thickness of Sections A and B is about 14.07% and 14% chord respectively, both occurring at 38% chord.

On the lower surface the point of inflection occurs at 67% chord for Section A and 66% chord Section B.

The trailing edge of Section B is at 0.75% chord above the chord line.

Sections C to G all have the same ordinates as Section B forward of 65% chord. Consequently the ordinates thereof are only iisted rearward of this point, see Figure 4.

The main feature of Section C is a base thickness of 1% chord and an associated increase in rear loading as compared with Section B. Section D has an increased camber over the rear 35% chord with respect to Section B, with negligible base thickness, while Section E has even more rear camber. Sections F and G are developments of Section D, but having 0.5% chord and 1% chord base thickness respectively.

Figure 6 compares theoretical and experimental pressure distributions over Section B at design conditions (M = 0.734, GL = 0.61), the test being carried out transition fixed and at a Reynolds number of 20 x 106. It indicates that the design objectives of that section were largely achieved.

Experimentally a peak local Mach No of about 1.25 is achieved at about 5% chord on the upper surface.

Then follows a net recompression, which is advantageously substantially isentropic, terminating in a weak shock of 1.12 Mach No at 55 % chord, and an acceptable adverse pressure gradient in terms-of slope and extent. The effect of the thick trailing edge on Section C is believed to be that it has greater rear loading than Section B. For a given lift coefficient this implies that less lift is required from the supercritical flow region (from 0.01 to 0.56 xlc), which in turn implies that for a given type of pressure distribution, the shock strength and hence the wave drag are lower. Alternatively it implies improved buffet performance or relieves viscous drag problems associated with an otherwise too adverse pressure gradient.

Evidence in favour of this proposition can be derived from Figure 7 in which drag coefficient versus lift coefficient plots are compared for Section B and C. At low coefficients the latter aerofoil has higher drag than the former (this may be attributed to the non-zero base pressure drag of Section C - that is the drag resulting from suction forces acting on the thick trailing edge). Above a certain lift coefficient (0.6 and 0.65 for low and high Reynolds number tests respectively) the drag of Section C is less than that of Section B. In other words wave drag can be traded forform drag or, by variable trailing edge geometry, an excessive rise in viscous drag postponed to a higher CL.

It is customary in aerodynamic circles to measure drag increase by units called "counts", which are Cc increments of 0.0002. A drag rise of 20 counts, ie 0.002 with Mach No. is in this context regarded as excessive.

Claims

1. An aerofoil suitable for use in the mainplane of a transport aircraft with a section design Mach number between 0.7 and 0.77 having a supercritical section with rear loading, the section having a thickness to chord ratio greater than 0.12, a leading edge region radius and camber such that there is generated at the design Mach No and lift coefficient a local peak Mach No of 1.1 to 1.3 forward of 10% chord, and an upper surface curvature decreasing to a minimum of value and location in combination with the leading edge region radius and camber such that aft of the peak local Mach No supersonic flow is maintained with a net recompression which if terminating in a shock does so with a Mach No thereat at least 0.05 less than the forward local peak Mach No, subsonic flow not being achieved till aft of 40% chord, and the lower surface contour being such that lift generated by the lower surface is maximised without excessive increase in viscous drag and is consistent with attenuating the pitchirfg moment coefficient of the section as a whole.

2. An aerofoil as claimed in claim 1 and whereover in use at design conditions the commencement of the adverse pressure gradient associated with recompression to subsonic flow is as far aft as possible before the onset of excessive viscous drag.

3. An aerofoil as claimed in claim 1 or claim 2 and wherein the upper surface minimum curvature point occurs at 30-40% chord.

4. An aerofoil as claimed in any one of claims 1 to 3 and wherein the maximum thickness is at 35-45% chord.

5. An aerofoil as claimed in any one of claims 1-4 and wherein upper surface curvature rises from the minimum to a maximum at about 75% chord and then diminishes steadily toward the trailing edge.

6. An aerofoil as claimed in claim 5 and wherein the maximum curvature value at 75% chord is 4t1c2.

7. An aerofoil as claimed in any one of the preceding claims and wherein the minirnum curva- ture value is 2-2.5 t/c2.

8. An aerofoil as claimed in any one of the preceding claims and having a leading edge radius of 1.4% chord.

9. An aerofoil as claimed in any one of the preceding claims and wherein the slope of the camber line at the leading edge is 6" to the chord line, nose up.

10. An aerofoil as claimed in any one of the preceding claims and having a flexible nose flap of the kind described in UK Patent Specification 1296994.

11. An aerofoil as claimed in any one of claims 1-9 and having a leading edge slat.

12. An aerofoil as claimed in any one of claims 1-9 land having a leading edge droop flap.

13. An aerofoil as claimed substantially as hereinbefore described with reference to the accompanying drawings.