GAS TURBINE ENGINE FAN BLADE WITH AXIAL LEAN

Abstract

A fan blade for a gas turbine engine is arranged such that for any two points on its leading edge that are in the radially outer 40% of the blade span and are radially separated by at least 5% of the blade span, the radially outer of the two points is axially forward of the radially inner point. The radius of the leading edge of a given fan blade at the hub divided by the radius of the leading edge of the fan blade at the tip is less than or equal to 0.3. Such an arrangement may result in an improved operability range.

Claims

1. A fan stage for a gas turbine engine, the fan stage defining axial, radial and circumferential directions, the fan stage comprising a plurality of fan blades extending from a hub, wherein: each fan blade comprises an aerofoil portion having a leading edge extending from a root to a tip the radial distance between the leading edge at the root (A) and the leading edge at the tip (C) defining a blade span; for any two points (P1, P2) on the leading edge of a fan blade that are in the radially outer 40% of the blade span and are radially separated by at least 5% of the blade span, the radially outer of the two points is axially forward of the radially inner point; and the radius of the leading edge of a given fan blade at the hub divided by the radius of the leading edge of the fan blade at the tip is less than or equal to 0.3.

2. A fan stage for a gas turbine engine according to claim 1, wherein: for any two points (P1, P2) on the leading edge of the fan blade that are radially closer to the tip than to the root and are radially separated by at least 5% of the blade span, the radially outer of the two points is axially forward of the radially inner point.

3. A fan stage according to claim 1, wherein for any two points (P1, P2) on the leading edge of the fan blade that are in the radially outer 40% of the blade span and have a difference in radius of at least 2% of the blade span, the radially outer of the two points is axially forward of the radially inner point.

4. A fan stage according to claim 1, wherein for any two points (P1, P2) on the leading edge of the fan blade 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 radially outer of the two points is axially forward of the radially inner point.

5. A fan stage according to claim 1, wherein for any two points (P1, P2) on the leading edge of a fan blade that are radially separated by at least 5% of the blade span, the radially outer of the two points is axially forward of the radially inner point.

6. A fan stage according to claim 1, wherein, when viewed along a circumferential direction, the angle ( (P1, P2)) formed between the radial direction and a line drawn between the two points (P1, P2) on the leading edge 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.

7. A fan stage according to claim 1, 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 of any given fan blade is in the range of from 6 and 0.2, 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.

8. A fan stage according to claim 1, wherein: the maximum perpendicular distance (e) between any point on the leading edge of any given fan blade and a straight line drawn between the leading edge at the root and at the tip is 2% of the blade span.

9. A fan stage according to claim 1, wherein each fan blade comprises: a platform; and a root portion, wherein the root portion extends between the platform and the root of the aerofoil portion.

10. A fan stage according to claim 9, wherein the radial extent of the root portion of each fan blade is no more than 7% of the span of the aerofoil portion.

11. A fan stage according to claim 1, wherein each fan blade comprises a tip portion that extends at least radially away from the tip of the aerofoil portion.

12. A fan stage according to claim 11, wherein the radial extent of the tip portion of each fan blade is no more than 7% of the span of the aerofoil portion.

13. A fan stage according to claim 1, wherein the aerofoil portion of each fan blade has a trailing edge extending from the root to 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.

14. A gas turbine engine comprising a fan stage according to claim 1.

15. A gas turbine engine according to claim 14, further comprising: a turbine; and a gearbox, wherein: the fan stage 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.

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

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

18. A method of manufacturing a fan stage, the fan stage defining axial, radial and circumferential directions, the fan stage comprising a plurality of fan blades extending from a hub, wherein each fan blade comprises an aerofoil portion having a leading edge extending from a root to a tip the radial distance between the leading edge at the root (A) and the leading edge at the tip (C) defining a blade span; for any two points (P1, P2) on the leading edge of a fan blade that are in the radially outer 40% of the blade span and are radially separated by at least 5% of the blade span, the radially outer of the two points is axially forward of the radially inner point; and the radius of the leading edge of a given fan blade at the hub divided by the radius of the leading edge of the fan blade at the tip is less than or equal to 0.3, the method comprising: providing the fan hub; and attaching the plurality of fan blades to the fan hub using linear friction welding.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0045] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0046] FIG. 1 is a sectional side view of a gas turbine engine on accordance with the present disclosure;

[0047] FIG. 2 is a side view of a fan blade according to an example of the present disclosure;

[0048] FIG. 3 is a close-up view of a leading edge portion of a fan blade according to an example of the present disclosure; and

[0049] FIG. 4 is a side view of a fan blade according to an example of the present disclosure.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0050] 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.

[0051] 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.

[0052] 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.

[0053] 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.

[0054] 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.

[0055] 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.

[0056] 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.

[0057] 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.

[0058] In general, the fan stage 13 (which may be referred to simply as the fan 13) has a hub to tip ratio, which may be defined as the radius of the leading edge of the fan blades 100 at the point where they extend away from the hub 200 (labelled r.sub.hub in FIG. 1) divided by the radius of the leading edge of the fan blades 100 at their tip 150 (labelled r.sub.tip in FIG. 1). The hub to tip ratio (r.sub.hub/r.sub.tip) may be in the ranges described and/or claimed elsewhere herein.

[0059] FIG. 2 is a side view (that is, a view in the axial-radial plane) of a fan blade 100 in accordance with the present disclosure. 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 120 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.

[0060] Various features of an exemplary fan blade 100 will now be described with reference to FIGS. 2 and 3. It will be appreciated that these features may be applied alone or in combination, as defined in the claims. The variables shown in FIGS. 2 and 3 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 a percentage (for example 2% or 5%) of the aerofoil span apart, and between the radii of point F on the leading edge at 60% of the aerofoil span from the LE root A and LE Tip C e.g. [00001]$\begin{array}{c}\frac{.Math.r_{P.Math..Math.1}-r_{P.Math..Math.2}.Math.}{r_C-r_A}e.g..Math.0.05.Math..Math.\mathrm{or}.Math..Math.0.02;.Math.r_{P.Math..Math.2}{r}_{P.Math..Math.1};\\ r_A+0.6.Math.\left(r_C-r_A\right)r_{P.Math..Math.1},r_{P.Math..Math.2}{r}_C\end{array}$ P3, P4 LE points Points on leading edge 120 with radii that are at least a percentage (for example 2% or 5%) of the aerofoil span apart, and between the radii of the LE root A and the point F on the leading edge at 60% of the aerofoil span from the root A, e.g. [00002]$\begin{array}{c}\frac{.Math.r_{P.Math..Math.4}-r_{P.Math..Math.3}.Math.}{r_C-r_A}0.05;r_{P.Math..Math.3}{r}_{P.Math..Math.4};\\ r_A{r}_{P.Math..Math.3},r_{P.Math..Math.4}{r}_A+0.6.Math.\left(r_C-r_A\right)\end{array}$ E Point on LE Point on leading edge 120 with maximum perpendicular distance to the line AC H Point on AC e Distance EH [00003]$e=\sqrt{{\left(x_H-x_E\right)}^2+{\left(r_H-r_E\right)}^2}$ Alpha LE global slope [00004]$=\frac{180}{}.Math.\mathrm{atan}.Math..Math.\left(\frac{x_c-x_A}{r_c-r_A}\right)$ alpha (P1,P2) LE local slope radially outer half [00005]$\left(P.Math..Math.1,P.Math..Math.2\right)=\frac{180}{}.Math.\mathrm{atan}\left(\frac{x_{P.Math..Math.2}-x_{P.Math..Math.1}}{r_{P.Math..Math.2}-r_{P.Math..Math.1}}\right)$ alpha (P3,P4) LE local slope radially inner half [00006]$\left(P.Math..Math.3,P.Math..Math.4\right)=\frac{180}{}.Math.\mathrm{atan}\left(\frac{x_{P.Math..Math.3}-x_{P.Math..Math.4}}{r_{P.Math..Math.3}-r_{P.Math..Math.4}}\right)$ e% LE straightness [00007]$e.Math..Math.\%=\frac{e}{r_C-r_A}.Math.100$ 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 F Aerofoil span Radius of the leading edge 120 at radial 60% point position 60% between root 140 and tip 150 from root 140: r.sub.A + 0.6 (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.

[0061] For any two points P1, P2 on the leading edge that are in the region 300 that is the radially outer 40% of the blade span (i.e. radially outboard of the point F shown in FIG. 2), for example radially closer to the leading edge tip C than to the leading edge root A, and are separated by at least 1%, for example at least 2%, for example at least 3%, for example at least 4%, for example at least 5% of the span, the radially outer point P2 is axially forward of the radially inner point P1. Such points P1, P2 may be described as being radially outside (or as having a greater radius than) the point F on the aerofoil that is 60% of the span from the leading edge root A.

[0062] For points P3, P4 that are on the leading edge that are in the region 310 that is the radially inner 60% of the blade span (i.e. radially inside of the point F shown in FIG. 2), for example radially closer to the leading edge root A than to the leading edge tip C, and are separated by at least 1%, for example at least 2%, for example at least 3%, for example at least 4%, for example at least 5% of the span, the radially outer point P3 may either be axially forward (as in the FIG. 2 example) or axially rearward of the radially inner point P4. Such points P3, P4 may be described as being radially inside (or as having a smaller radius than) the point F on the aerofoil that is 60% of the span from the leading edge root A. The leading edge 120 of the aerofoil 100 may have any desired shape in the region 310.

[0063] Optionally, 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. 2, 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.

[0064] Optionally, the local slope (P1, P2), (P3, P4) 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. 2, 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), (P3, P4) 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 1%, for example at least 2%, for example at least 3%, for example at least 4%, for example at least 5% of the blade span.

[0065] The relationship between E and H as defined in the table above is seen most easily in FIG. 3. 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. However, it will be appreciated that some arrangements of the present disclosure may have a relationship between E and H that is outside this range.

[0066] 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.

[0067] The fan blade 100 comprises a platform 160. The aerofoil portion 110 may extend directly from the platform 160, as in the FIG. 2 example. Alternatively, as shown by way of example in FIG. 4, 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.

[0068] Also as shown by way of example in FIG. 4, 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.

[0069] 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.

[0070] 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.

[0071] 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.