GAS TURBINE ENGINE FAN BLADE WITH AXIAL LEAN

Abstract

A fan blade for a gas turbine engine is provided with forward axial lean. The fan blade may have a substantially straight leading edge. The geometry of the fan blade results in a lower susceptibility to flutter, thereby allowing a gas turbine engine comprising such a fan blade to operate over a wider range of operating conditions.

Claims

1. A fan blade for a gas turbine engine, the gas turbine engine defining axial, radial and circumferential directions, the fan blade comprising: an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span wherein, when viewed along a circumferential direction: the angle () formed between the radial direction and a straight line (AC) drawn between the leading edge at the root and at the tip is in the range of from 6 and 0.2; and the angle ( (P1, P2)) formed between the radial direction and a line drawn between any two points (P1, P2) on the leading edge that have a difference in radius of at least 5% of the blade span is in the range of from 6 and 0, where a negative angle indicates that the respective line has an axial component that is in the same direction as the axial component of the direction from a trailing edge to the leading edge of the blade.

2. A fan blade for a gas turbine engine according to claim 1, wherein: when viewed along a circumferential direction, the maximum perpendicular distance (e) between any point on the leading edge and a straight line drawn between the leading edge at the root and at the tip is 2% of the blade span.

3. A fan blade for a gas turbine engine according to claim 1, wherein the fan blade comprises: a platform; and a root portion, wherein the root portion extends between the platform and the root of the aerofoil portion.

4. A fan blade for a gas turbine engine according to claim 3, wherein the radial extent of the root portion is no more than 7% of the span of the aerofoil portion.

5. A fan blade for a gas turbine engine according to claim 1, wherein the fan blade comprises a tip portion that extends at least radially away from the tip of the aerofoil portion.

6. A fan blade for a gas turbine engine according to claim 5, wherein the radial extent of the tip portion is no more than 7% of the span of the aerofoil portion.

7. A fan blade for a gas turbine engine according to claim 1, wherein: for two points on the leading edge that are radially closer to the tip than to the root and have a difference in radius of at least 5% of the blade span, the axial position of the radially outer point is forward of the axial position of the radially inner point.

8. A fan blade for a gas turbine engine according to claim 7, wherein: for two points on the leading edge that are radially closer to the tip than to the root and have a difference in radius of at least 2% of the blade span, the axial position of the radially outer point is forward of the axial position of the radially inner point.

9. A fan blade for a gas turbine engine according to claim 1, wherein the aerofoil portion has a trailing edge extending from the root the tip, wherein, when viewed along a circumferential direction the angle () formed between the radial direction and a straight line (BD) drawn between the trailing edge at the root and at the tip is in the range of from 20 and 5, where a negative angle indicates that the respective line has an axial component that is in the same direction as the axial component of the direction from a trailing edge to the leading edge of the blade.

10. A fan stage for a gas turbine engine comprising: a hub; and a plurality of fan blades according to claim 1, wherein: the fan blades extend radially form the hub.

11. A fan stage for a gas turbine engine according to claim 10, wherein: the ratio of the radius of the position where the leading edge of one of the fan blades meets the hub to the outermost radial extent of the leading edge of the fan blade is less than 0.33.

12. A gas turbine engine comprising a fan blade according to claim 1.

13. A gas turbine engine comprising a fan stage according to claim 10.

14. A gas turbine engine comprising: a fan having a plurality of fan blades according to claim 1; a turbine; and a gearbox, wherein: the fan is driven from the turbine via the gearbox, in order to reduce the rotational speed of the fan stage compared with the driving turbine stage.

15. A gas turbine engine according to claim 12 with a specific thrust of less than 100 N/Kg/s.

16. A gas turbine engine according to claim 13 with a specific thrust of less than 100 N/Kg/s.

17. A gas turbine engine according to claim 14 with a specific thrust of less than 100 N/Kg/s.

18. A method of manufacturing a fan stage for a gas turbine engine comprising: providing a fan hub; and attaching a plurality of fan blades the fan hub using linear friction welding, the fan blades comprising an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span wherein, when viewed along a circumferential direction,the angle () formed between the radial direction and a straight line (AC) drawn between the leading edge at the root and at the tip is in the range of from 6 and 0.2; and the angle ( (P1, P2)) formed between the radial direction and a line drawn between any two points (P1, P2) on the leading edge that have a difference in radius of at least 5% of the blade span is in the range of from 6 and 0, where a negative angle indicates that the respective line has an axial component that is in the same direction as the axial component of the direction from a trailing edge to the leading edge of the blade.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0046] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0047] FIG. 1 is a sectional side view of a gas turbine engine on accordance with the present disclosure;

[0048] FIG. 2 is a radial view of a fan blade according to an example of the present disclosure;

[0049] FIG. 3 is a side view of a fan blade according to an example of the present disclosure;

[0050] FIG. 4 is another side view of a fan blade according to an example of the present disclosure;

[0051] FIG. 5 is a close-up view of a leading edge portion of a fan blade according to an example of the present disclosure; and

[0052] FIG. 6 is a side view of a fan blade according to an example of the present disclosure.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0053] With reference to FIG. 1, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

[0054] The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place.

[0055] The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high 17, intermediate 18 and low 19 pressure turbines drive respectively the high pressure compressor 15, intermediate pressure compressor 14 and fan 13, each by suitable interconnecting shaft.

[0056] The gas turbine engine 10 and/or the fan stage 13 and/or the fan blades 100 of the fan stage 13 shown in FIG. 1 may be in accordance with examples of the present disclosure, aspects of which are described by way of example only in relation to FIGS. 2 to 6.

[0057] Any gas turbine engine in accordance with the present disclosure (such as the gas turbine engine 10 of FIG. 1) may, for example, have a specific thrust in the ranges described herein (for example less than 10) and/or a fan blade hub to tip ratio in the ranges described herein and/or a fan tip loading in the ranges described herein.

[0058] The present disclosure may relate to any suitable gas turbine engine. For example, other gas turbine engines to which the present disclosure may be applied may have related or alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan. The gas turbine engine shown in FIG. 1 has a mixed flow nozzle 20, meaning that the flow through the bypass duct 22 and the flow through the core 15, 16, 17, 18, 19 are mixed, or combined, before (or upstream of) the nozzle 20). However, this is not limiting, and any aspect of the present disclosure may also, for example, relate to engines 10 having a split flow nozzle, which may mean that the flow through the bypass duct 22 has its own nozzle that is separate to and may be radially outside a core engine nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.

[0059] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction 30 (which is aligned with the rotational axis 11), a radial direction 40, and a circumferential direction 50 (shown perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions 30, 40, 50 are mutually perpendicular.

[0060] The fan stage 13 comprises a plurality of fan blades 100 extending from a hub 200. The fan blades 100 may be defined with respect to the axial direction 30, radial direction 40, and circumferential direction 50 shown in FIG. 1 in relation to the gas turbine engine 10.

[0061] FIG. 2 is a view in a radially inward direction of a fan blade 100. FIG. 3 is a side view (that is, a view in the axial-radial plane) of the fan blade 100. The fan blade 100 has an aerofoil portion 110. The aerofoil portion 110 has a leading edge 120 and a trailing edge 130. The aerofoil portion 110 extends from a root 140 to a tip 150 in a substantially radial spanwise direction. The leading edge 110 may be defined as the line defined by the axially forwardmost points of the aerofoil portion 110 from its root 140 to its tip 150.

[0062] As mentioned above, the susceptibility of a fan blade 100 to flutter is at least part dependent on the torsional content of the lowest natural frequency mode shape (also referred to as the 1F mode shape)

[0063] This torsional content can be defined at the tip 150 of the blade 100 by a parameter l/Ch, where l is the distance (for example the shortest distance) from the leading edge 120 at the tip 150 to the centre of twist 310 in the mode shape, and Ch is the chord length at the tip 150 (see FIG. 2). Therefore a relatively higher value of l/Ch represents a relatively lower torsional content in the mode shape. The parameter l/Ch may be thought of as describing the relative motion (for example in a substantially circumferential direction) between leading edge 120 and trailing edge 130 at the tip 150 (the motion of the leading edge 120 at the tip 150 usually being greater than the motion of the trailing edge 130 at the tip 150). In FIG. 2, the reference label 120 indicates the leading edge 120 in the deformed (or displaced) 1F mode shape, and the reference label 150 indicates the tip 150 in that deformed (or displaced) 1F mode shape.

[0064] FIG. 3 shows the nodal point 320 on the leading edge 120 and the nodal point 330 on the trailing edge 130 in the 1F mode shape, the nodal points 320, 330 being points on the blade 100 (for example towards the root 140) which are stationary in the 1F mode shape. To a first approximation the displacement of the tip 150 at the leading edge 120 depends on (or at least is affected by) the distance a between the leading edge 120 at the tip 150 and the 1F nodal point 320 on the leading edge 120. Similarly, to a first approximation the displacement of the tip 150 at the trailing edge 130 depends on (or at least is affected by) the distance b between the trailing edge 130 at the tip 150 and the 1F nodal point 330 on the trailing edge 130.

[0065] Typically, the distance a is greater than the distance b. This may be at least in part because typically the 1F nodal point 320 on the leading edge 120 is radially inside the 1F nodal point 330 on the trailing edge 130.

[0066] The blade 100 shown in the Figures by way of example of the present disclosure may be said to be leant axially forwards, for example by at least having the line a pointing axially forwards, i.e. to the left in FIG. 3. As shown in the FIG. 3 example, the line b may also point axially forwards, for example at a greater angle than the line a.

[0067] As explained elsewhere herein, the blade geometry described and/or claimed herein may reduce the susceptibility of the blades to flutter. For example, and without being limited or bound to a particular theory, the ratio between lengths a and b may be decreased compared with conventional blades, which may result in a decrease in the ratio between the tip displacement at the leading edge 120 and the tip displacement at the trailing edge 130 in the 1F mode. The fan blades 100 (and/or aerofoil portions 110) may have increased l/Ch compared with conventional fan blades, which may help to reduce the torsional content of the 1F mode shape, and thus reduce the susceptibility to flutter.

[0068] Various features of an exemplary fan blade 100 will now be described with reference to FIGS. 4 and 5. It will be appreciated that these features may be applied alone or in combination, as defined in the claims. The variables shown in FIGS. 4 and 5 are explained in the table below, where the term LE refers to the leading edge 120, and the term TE refers to the trailing edge 130:

TABLE-US-00001 Point Name Definition A LE root Leading edge 120 at the root 140 of the aerofoil 110 B TE root Trailing edge 130 at the root 140 of the aerofoil 110 C LE tip Leading edge 120 at the tip 150 of the aerofoil 110 D TE tip Trailing edge 130 at the tip 150 of the aerofoil 110 P1, P2 LE points Points on leading edge 120 with radii that are at least 5% of the aerofoil span apart [00001]$\frac{.Math.r_{P.Math..Math.1}-r_{P.Math..Math.2}.Math.}{r_C-r_A}0.05;r_{P.Math..Math.2}{r}_{P.Math..Math.1}$ E Point on LE Point on leading edge 120 with maximum perpendicular distance to the line AC H Point on AC e Distance EH [00002]$e=\sqrt{{\left(x_H-x_E\right)}^2+{\left(r_H-r_E\right)}^2}$ Alpha LE global slope [00003]$=\frac{180}{}.Math.\mathrm{atan}.Math..Math.\left(\frac{x_c-x_A}{r_c-r_A}\right)$ alpha(P1, P2) LE local slope [00004]$\left(P.Math..Math.1,P.Math..Math.2\right)=\frac{180}{}.Math.\mathrm{atan}.Math..Math.\left(\frac{x_{P.Math..Math.2}-x_{P.Math..Math.1}}{r_{P.Math..Math.2}-r_{P.Math..Math.1}}\right)$ e % LE straightness [00005]$e.Math..Math.\%=\frac{e}{r_C-r_A}.Math.100$ Span Aerofoil Span Difference in the radius of the leading edge 120 at the root 140 and at the tip 150: r.sub.C r.sub.A Note in the above table, that x refers to a position in the axial direction 30 and r refers to a position in the radial direction 40.

[0069] The global slope of the aerofoil portion 110 of fan blades 100 as described and/or claimed herein, such as that shown by way of example in FIG. 4, may be in the range of from 6 and 0.2, for example within any of the ranges defined elsewhere herein. In this regard, the global slope may represent the angle formed between the radial direction and a straight line AC drawn between the leading edge point A at the root 140 and the leading edge point C at the tip 150.

[0070] The local slope (P1, P2) of the aerofoil portion 110 of fan blades 100 as described and/or claimed herein, such as that shown by way of example in FIG. 4, may be in the range of from 6 and 0, for example within any of the ranges defined elsewhere herein. In this regard, the local slope (P1, P2) may represent the angle formed between the radial direction and a line drawn between any two points on the leading edge that have a difference in radius of at least 5% of the blade span.

[0071] The relationship between E and H as defined in the table above is seen most easily in FIG. 5. The distance e between the points E and H may be said to represent the maximum perpendicular distance between any point on the leading edge and a straight line drawn between the leading edge at the root. As a percentage (e%)the distance e of the aerofoil portion 110 of fan blades 100 as described and/or claimed herein may be less than 5%, for example less than 2% (or any other range as described and/or claimed herein) of the span of the aerofoil portion 110.

[0072] The trailing edge 130 of the aerofoil portion 110 may also define a global slope . The global slope of the trailing edge 130 of fan blades 100 may be in the range of from 40 and 0, for example 30 and 1, for example 25 and 2.5, for example 20 and 5, for example 15 and 7.5, for example around 10. In this regard, the global slope of the trailing edge 130 may represent the angle between the radial direction and a straight line I drawn between a point B on the trailing edge 130 at the root 140 and a point D on the trailing edge 130 at the tip 150.

[0073] The fan blade 100 comprises a platform 160. The aerofoil portion 110 may extend directly from the platform 160, as in the FIG. 4 example. Alternatively, as shown by way of example in FIG. 6, a fan blade 100 may have a root portion 170. The root portion 170 may be said to extend between the platform 160 and the root 140 of the aerofoil portion 110. The radial extent of the root portion 170 may be no more than 7%, for example no more than 5%, of the span of the aerofoil portion 110.

[0074] Also as shown by way of example in FIG. 6, the fan blade 100 may comprise a tip portion 180. The tip portion 180 may be said to extend from the tip 150 of the aerofoil portion 110. The radial extent of the tip portion 180 may be no more than 5% of the span of the aerofoil portion 110.

[0075] The fan blade 100 may be attached to the hub 200 in any desired manner. For example, the fan blade 100 may comprise a fixture 190 such as that shown by way of example in FIG. 6 which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc.

[0076] Alternatively, the fan blade 100 and the hub 200 may be formed as a unitary part, with no mechanical and/or releasable connections, so as to form a unitary fan stage 13. Such a unitary fan stage 13 may be referred to as a blisk. Such a unitary fan stage 13 may be manufactured in any suitable manner, for example by machining and/or by linear friction welding the fan blades 100 to the hub 200, or at least linear friction welding the aerofoil portions 110 to a hub 200 that includes radially inner stub portions of the fan blades 100.

[0077] The hub to tip ratio, which may have a value as indicated elsewhere herein, may be defined as the radius of the leading edge 120 at the root 140 (which may itself be referred to as a hub) of the aerofoil 110 (point A) divided by the radius of the leading edge 120 at the tip 150 of the aerofoil 110 (point B), i.e. r.sub.A/r.sub.B.

[0078] It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.