GAS TURBINE ENGINE FAN

Abstract

A fan stage of a ducted fan gas turbine engine has a rotor hub having a principal axis of rotation and a plurality of fan blades having a hub end attached to the hub and extending radially towards a tip end so as to define a blade span dimension. Each blade has a leading and a trailing edge, a chord for a section of the blade being a straight line joining the leading and trailing edges within the section. A difference between a stagger angle in a mid-span region and in the vicinity of the tip end of each blade is greater than or equal to 20°. The fan blades are twisted to a greater extent than conventional between the mid-span and tip end. A camber angle difference between the mid-span region and the tip end may be greater than 30 degrees.

Claims

1. A fan stage (23) of a ducted fan gas turbine engine (10), comprising: a rotor hub (42) having a principal axis of rotation (9); and a plurality of fan blades (25), each blade having a hub end (44) attached to the hub and extending radially outwardly towards a tip end (46) so as to define a blade span dimension (48), each blade further having a leading edge (52) and a trailing edge (54), a chord (64) for a section of the blade being a straight line joining the leading (52) and trailing (54) edges within the section and a stagger angle being the angle formed between the chord (64) and the principal axis of rotation (9); wherein a difference between the stagger angle in a mid-span region (50) of each blade and the stagger angle in a vicinity of the tip end (46) of each blade is greater than or equal to 20°, and less than or equal to 80°.

2. The fan stage according to claim 1, wherein the difference between the stagger angle in the mid-span region of each blade and the stagger angle in the vicinity of the tip end is greater than or equal to 35° or 38°, and less than or equal to 80°.

3. The fan stage according to claim 1, wherein the mid-span region comprises a region between 40% and 60% of the blade span between the hub end and tip end.

4. The fan stage according claim 1, wherein the stagger angle in the mid-span region comprises the stagger angle for a blade section at a midpoint of the blade span.

5. A fan stage according claim 1, wherein the vicinity of the tip end comprises a region within 5% of the blade span or less from the tip end.

6. The fan stage according to claim 1, comprising: a blade inlet angle at the leading edge of the blade, the blade inlet angle defined as an angle between a local axis of the blade at the leading edge and the principal rotation axis; a blade outlet angle at the trailing edge of the blade, the blade outlet angle defined as an angle between a local axis of the blade at the trailing edge and the principal rotation axis; a camber angle defined as a difference in the blade inlet angle and the blade outlet angle; the mid-span region comprising a camber angle; the vicinity of the tip end comprising a different camber angle; and where a camber angle difference between the mid-span camber angle and the camber angle in the vicinity of the tip end is greater than or equal to 30° and less than or equal to 70°.

7. The fan stage according to claim 6, where the mid-span region camber angle is greater than 45° and less than or equal to 70°.

8. The fan stage according to claim 6, where the camber angle in the vicinity of the tip end is less than or equal to 15°.

9. A fan blade for a gas turbine engine, comprising: a hub end for attachment to a rotor hub in use, the blade extending from the hub end towards a tip end so as to define a blade span dimension, the blade further having a leading edge and a trailing edge, a chord for a sectional profile of the blade being a straight line joining the leading and trailing edges within the sectional profile; wherein a relative difference in angle between a chord in a mid-span region of the blade and a chord in a vicinity of the tip end of the blade is greater than or equal to 20° and less than or equal to 80°.

10. The fan blade according to claim 9, wherein the relative difference in angle between the chord in a mid-span region of the blade and the chord in the vicinity of the tip end of the blade is greater than or equal to 30° and less than or equal to 80°.

11. The fan blade according to claim 9, wherein the mid-span region comprises a region between 30% and 60% of the blade span between the hub end and tip end.

12. The fan blade according to claim 9, wherein the chord in the mid-span region comprises a chord at a midpoint of the blade span.

13. The fan blade according to claim 9, wherein the vicinity of the tip end comprises a region within 5% of the blade span or less from the tip end.

14. The fan blade according to claim 9, comprising: a blade inlet angle at the leading edge of the blade, the blade inlet angle defined as an angle between a local axis of the blade at the leading edge and a common axis; a blade outlet angle at the trailing edge of the blade, the blade outlet angle defined as the angle between the local axis of the blade at the trailing edge and the common axis; a camber angle defined as a difference in the blade inlet angle and the blade outlet angle; a mid-span region comprising a camber angle; a tip end region comprising a different camber angle; and where a camber angle difference between the mid-span region and the vicinity of the tip end is greater than 30° and less than or equal to 70°.

15. The fan blade according to claim 14, where the mid-span region camber is greater than 45° and less than or equal to 70°, and the tip end camber is less than 15°.

16. The fan blade according to claim 14, wherein the camber in both the mid-span region and the vicinity of the tip end is an acute angle.

17. A fan stage of a ducted fan gas turbine engine, comprising: a rotor hub having a principal axis of rotation; a plurality of fan blades, each blade having a hub end attached to the hub and extending radially outwardly towards a tip end so as to define a blade span dimension, each blade further having a leading edge and a trailing edge, a chord for a section of the blade being a straight line joining the leading and trailing edges within the section and a stagger angle being the angle formed between the chord and the principal axis of rotation; a blade inlet angle at the leading edge of the blade, the blade inlet angle defined as an angle between a local axis of the blade at the leading edge and the principal rotation axis; a blade outlet angle at the trailing edge of the blade, the blade outlet angle defined as an angle between a local axis of the blade at the trailing edge and the principal rotation axis; a camber angle defined as a difference in the blade inlet angle and the blade outlet angle; a mid-span region comprising a camber angle; a vicinity of the tip end comprising a different camber angle; and where a camber angle difference between the mid-span camber angle and the camber angle in the vicinity of the tip end is greater than or equal to 30° and less than or equal to 70°, wherein the mid-span region comprises a region between 40% and 60% of the blade span between the hub end and tip end, wherein the vicinity of the tip end comprises a region within 5% of the blade span or less from the tip end, wherein the difference between the stagger angle in a mid-span region of each blade and the stagger angle in the vicinity of the tip end of each blade is greater than or equal to 20° and less than or equal to 80°, wherein the mid-span region camber angle is greater than 45 degrees and less than or equal to 80°.

18. The fan stage according to claim 17, wherein the camber angle difference between the mid-span camber angle and the camber angle in the vicinity of the tip end is greater than or equal to 30° and less than or equal to 60°.

19. The fan stage according to claim 17, wherein the difference between the stagger angle in a mid-span region of each blade and the stagger angle in the vicinity of the tip end of each blade is greater than or equal to 20° and less than or equal to 70°.

20. The fan stage according to claim 17, wherein the camber angle in the vicinity of the tip end is greater than or equal to 0 degrees and less than or equal to 15 degrees.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0041] FIG. 1 is a sectional side view of a gas turbine engine;

[0042] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0043] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0044] FIG. 4 is a schematic sectional side view of a fan region of a gas turbine engine;

[0045] FIG. 5 is a sectional view of a fan blade of a gas turbine engine;

[0046] FIG. 6 is an efficiency versus air flow curve for a fan blade in a conventional system;

[0047] FIG. 7 is a sectional view of the tip of a fan blade;

[0048] FIG. 8 is a sectional view of the mid-span of a modified fan blade;

[0049] FIG. 9 is a sectional view of a fan blade of a gas turbine engine;

[0050] FIG. 10 is an efficiency versus air flow curve for a fan blade in a conventional system and a modified fan blade; and,

[0051] FIG. 11 is an efficiency versus air flow curve for a fan blade in a conventional system; a modified fan blade; and further modified fan blade.

DETAILED DESCRIPTION

[0052] FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0053] The examples described hereinbelow concern a fan 23 that has been modified relative to a conventional fan by modifying the shape of the fan blades 25 thereof.

[0054] In use, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 where further compression takes place. 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 is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low-pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0055] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low-pressure turbine 19 (see FIG. 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to process around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

[0056] Note that the terms “low-pressure turbine” and “low-pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the “low-pressure turbine” and “low-pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

[0057] The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the disclosure. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0058] The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

[0059] It will be appreciated that the arrangement shown in FIGS. 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

[0060] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

[0061] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0062] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

[0063] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

[0064] FIG. 4 shows the geometry of a fan stage of a gas turbine engine. A fan blade 41 is attached to a hub 42 at a root end 44 and comprises a tip 46 at a distal/free end of the blade 41. The hub is configured to rotate about a principal engine axis 9 as described in relation to FIG. 1.

[0065] The tip 46 represents the radially outermost extent of the blade 41 with respect to the axis 9. A span dimension 48 extends between the root 44 and the tip 46, i.e. representing a radial height of the blade 41.

[0066] The blade 41 has a leading edge 52 and a trailing edge 54, each extending from the root end 44 to the tip 46. The leading and trailing edges 54 may have the same or different lengths, e.g. corresponding to a varying span profile

[0067] A mid-span 50 is defined by a midpoint of the leading edge 52 of the blade 41 between the root end 44 and the tip 46, i.e. the radial height of the blade from the surface of the hub to the tip. A mid-span section 56 is defined by the section of the blade 41 at the mid-span 50. The mid-span section could be defined as a cross-section that is parallel to the axis 9 or else as a section that is in a plane containing the midpoint of the leading edge 52 and the midpoint of the trailing edge 54. A tip section 58 is defined by the cross section of the blade 41 at the tip 46 of the blade 41, e.g. in a plane containing the tip of the leading 52 and trailing 54 edges.

[0068] The fan is positioned in front of a bypass stator 60 and a core stator 62.

[0069] FIG. 5 shows a datum/normal sectional profile 43 of the blade 41 at the tip comprising leading 52 and a trailing 54 edges. A chord 64 defines a straight line extending between/through the leading edge 52 and the trailing edge 54. A chord length defines a length of the chord 64 between the leading 52 and trailing 54 edges.

[0070] The axis of rotation 9 is shown relative to the sectional profile as well as the direction of movement/rotation 55 of the blade 41 in use, i.e. perpendicular to the axis of rotation 9.

[0071] A stagger angle 66 of the blade 41 is defined by the angle the chord 64 makes relative to the rotation axis 9. A low stagger angle results in the blade 41 (i.e. chord 64) being angularly closer to the principal rotation axis 9; and a high stagger angle results in the blade being angularly further from the principal rotation axis 9 (i.e. closer to the direction of rotation 55).

[0072] Whilst the above description of the blade geometry of FIGS. 4 and 5 is for a conventional or datum blade 41, it applies equally to a modified blade 25, the features of which will be described below.

[0073] FIG. 6 shows a generic/datum example plot 69 of efficiency versus air flow for a fan blade tip section for a given rotational speed. Conventional fan blades are designed such that the sections of the blade all achieve their local peak efficiency 68 at the same point. This is normally designed to occur in the cruise operating region where the biggest benefit to fuel burn is achieved, thereby maximising the benefit of the overall peak efficiency of the fan blade. Each blade section is provided with a stagger angle 66 that yields a peak efficiency 68 that corresponds to the desired flow rate. The collection of desired sectional profiles along the span direction thereby define the overall blade profile. A section or blade that operates at a lower air flow than the peak efficiency 68 lies on the stall side 70 of the peak 68 and a blade that operates at a higher air flow than the peak lies on the choke side 72 of the peak 68.

[0074] FIG. 7 shows the cross-section 46 (in dashed lines) of the tip of a blade 41 for a conventional system designed to operate at peak efficiency 68. The cross-section 46 comprises a stagger angle 78.

[0075] FIG. 7 also shows a second cross-section 76 (in solid lines) of the tip of a modified blade 25 in accordance with an example of the present disclosure. The second cross section 76 comprises a stagger angle 80. The blade section 76 has an increase 82 in stagger angle (i.e. angled further away from the principal rotation axis 9) with respect to the stagger angle 78 of blade tip section 46 designed for operation at a conventional peak efficiency.

[0076] The tip stagger angle 80 of section 76 may be at least 1°, 3°, 5°, 7° or 9° higher than stagger angle 78. The tip stagger angle 80 of section 76 may be up to 10°, 12° or 14° higher than the stagger angle 78. Any of the above min and max values of tip stagger angle increase may be combined to define a range in which the tip stagger angle may lie.

[0077] The blade section 76 operates on the choke side of the peak efficiency 68.

[0078] FIG. 8 shows a first cross-section 56 (in dashed lines) of the mid-span of a conventional blade 41 designed to operate in accordance with a normal peak efficiency. The first cross-section 56 comprises a stagger angle 88.

[0079] FIG. 8 further shows a second cross-section 86 (in solid lines) of the mid-span of a modified blade 25 in accordance with an example of the present disclosure. The second cross section 86 comprises a stagger angle 90.

[0080] The blade section 86 has a decrease 92 in stagger angle (i.e. angled more towards a parallel axis of the principal rotation axis 9) with respect to the stagger angle 88 of blade mid-span section 56.

[0081] The modified mid-span stagger angle 90 may be at least 1°, 3°, 5°, 7° or 9° lower than the stagger angle 88. The modified mid-span stagger angle 90 may be up to 10°, 12°, or 14° lower than the stagger angle 88 to produce a conventional peak efficiency. Any of the above min and max values of mid-span stagger angle decrease may be combined to define a range in which the mid-span stagger angle may lie.

[0082] The combination of a higher stagger at the tip section 76 and a reduced stagger at the mid-span section 86 results in the modified blade 25 having a larger stagger angle change between the mid-span 50 and tip. This may be described as a blade that is twisted to a greater extent than is conventional between the mid span and tip.

[0083] The change in stagger angle between the mid-span section and the tip section is greater than or equal to 20°, 25°, 30° 35°, 40° and less than or equal to 60°, 65°, 70°, 75°, or 80°. In some examples, the change in stagger angle between the mid-span section and the tip section may be greater than or equal to 36°, 37° 38°, 39°, 40°, 41°, or 42° and less than or equal to 60°, 65°, 70°, 75°, or 80°. The precise value of twist selected may be based at least in part on the hub-to-tip ratio of the fan blades and/or the intended speed of rotation in use. That is to say, the speed of the blade tip is greater than the speed of the mid-span by an amount according to the radial length of the fan blades. For relatively slow rotation speeds associated with more efficient engines a relatively smaller change in stagger angle may be desirable.

[0084] The twist between the mid and tip section can be adapted to take into account the aerodynamic robustness of the corresponding sections. A section robustness depends on the amount of pressure ratio generated by that section (camber and rotational speed), the incoming Mach number at the section leading edge together with the chord allocated to that section to generate the flow turning. A key parameter to set the required chord is solidity, defined as the ratio of chord and blade to blade spacing at the section height. The aerodynamic robustness together with the requirements on blade weight, stress modal frequencies and mode shapes will help at defining an optimum blade twist added to the conventional, nearly-linear blade-speed-related twist between the hub and tip sections.

[0085] A high stagger angle of the type disclosed herein may be referred to as the blade being ‘over-closed’.

[0086] The modified blade 25 provides an increased rate of change of stagger angle local to the blade tip, which provides greater aerofoil lean (pressure surface inward, suction surface outward) local to the blade leading edge.

[0087] As shown in FIG. 9, the leading edge 52 of the blade 41 comprises a blade inlet angle 98. The blade inlet angle 98 is defined as the angle between the local axis 94 of the blade 41 at the leading edge 52 and the rotation axis 9.

[0088] The trailing edge 54 of the blade 41 comprises a blade outlet angle 100. The blade outlet angle 100 is defined as the angle between the local axis 96 of the blade 41 at the trailing edge 52 and the rotation axis 9. A camber angle for a blade section (i.e. blade 25 or 41) is defined as the difference in the blade inlet angle 98 and the blade outlet angle 100.

[0089] The tip section 76 of the modified blade 25 may comprise a camber angle of less than 15°. In some examples, the tip section 76 of modified blade 25 may comprise a camber angle of greater than or equal to 0° and less than or equal to 15°. This threshold may be used to protect the robustness of the tip section which operates at the highest leading edge Mach number and whose chord is limited by mechanical requirements.

[0090] The midspan section of the modified blade 25 may comprise a camber angle of greater than or equal to 40°, 45°, or 50° and less than or equal to 60°, 65°, 70°, or 75°. This threshold may be used since the mid-section provides the highest pressure rise, e.g. since it has the best robustness in terms of combination of lower leading edge Mach number whilst still provided with a significant rotational speed to generate a large enthalpy rise.

[0091] In various examples of this disclosure, the difference in camber angle between the mid-span section 86 and the tip section 76 may be greater than or equal to 30° and less than or equal to 65°, 70°, or 75°. The difference in camber angle between the mid-span section and the tip section may be greater than or equal to 31°, 32° or 33°, and less than or equal to 68°, 69°, or 70°.

[0092] This difference in camber may characterise aspects of this disclosure, e.g. instead of, or in addition to, the change in stagger angle between the mid-section and tip. The difference in camber angle may be described as a blade that undergoes a greater change in curvature between the mid-section and tip. This may be described as a blade that is more arched than a conventional blade towards the mid-section and/or less arched (i.e. flatter) than a conventional blade towards its tip.

[0093] FIG. 10 shows the engine efficiency versus flow curve for an example 102 of a modified fan blade tip section 76 (e.g. in a geared turbofan) and the conventional curve 69 of FIG. 6.

[0094] The tip sections 76 of the fan blades 25 are deliberately designed to operate away from their local peak efficiency 106, as shown by point 104 in FIG. 10, i.e. on the choked side of their loss loops.

[0095] By designing the tip sections of the fan blades 25 on the choke side of their loss loops, away from their local peak efficiency, they have a greater flow range before stall, which gives the overall fan blade 25 a greater stability margin.

[0096] A key feature of geared fans is the reduced rotational speed of the fan blades, which reduces inlet relative flow Mach number. This lower Mach number gives shallower efficiency characteristics (or loss loops) which allows the tip sections of the fan blade to be operated away from their local peak efficiency with only a small efficiency reduction 108.

[0097] FIG. 10 shows that increasing the stagger of the fan tip sections reduces their flow capacity. If the overall capacity of the fan blade is not to be reduced the capacity of the blade inboard of the tip can be increased as described in relation to FIG. 8. This is achieved by opening the blade 25 (i.e. a reduced stagger) in the mid-span region. The mid-span region operates at a lower leading edge Mach number, due to being closer to the axis of rotation 9, and higher solidity (Blade chord/blade spacing/pitch ratio), and hence is inherently more aerodynamically robust in terms of efficiency, flow capacity and stability margin. Whilst the examples given above concern the mid-section, i.e. at half the blade span, it will be appreciated that the relevant changes may be made to a region comprising the mid-section or close to the mid-section.

[0098] FIG. 11 shows a flow versus efficiency curves 69 and 102 of FIG. 10 but including a plot/curve 110 for a further modified blade section that is thinner than the sections 46 and 76. In this regard it has been found to be a benefit of the modified fan blade 25 geometry described herein that the maximum tip thickness, i.e. the dimension 112 shown in FIG. 7, can be reduced by at least 5% relative to a conventional blade (i.e. a blade of optimal thickness for a conventional stagger angle 78 at the tip). In differing examples, the maximum tip thickness may be reduced by at least 10%, 15%, 20%, 25%, 30% or 35%. The tip thickness may be reduced by up to 30%, 35% or 40%. Any range of tip thickness reduction may be applied as defined by the optional min/max thresholds given above.

[0099] The reduction in thickness of the blade tip creates a corresponding reduction in weight towards the tip. A reduction in thickness of the blades 25 towards the tip would reduce structural strength. However, a key design consideration affecting the required strength of fan blades is the ability to withstand bird strike (or other foreign object impact). It has been found that the higher tip stagger provides a bird strike benefit due to a lowering of the incidence angle of birds/objects entering the engine intake relative to the rotating fan blade.

[0100] The typical axial velocity of a bird/object entering an engine is much lower than the air velocity into the engine. This normally results in a significant incidence angle relative to the blade leading edge and can result in significant bending or cupping of the blade leading edge geometry upon impact. The higher stagger fan tip sections disclosed herein reduce the incidence angle onto the leading edge, making the blade inherently more robust to bird, or foreign object, strikes.

[0101] This allows the tip sections to be thinned, thereby improving aerodynamic performance without compromising structural integrity. The beneficial impact of thinning the fan blades 25 towards the tip, and thereby reducing weight, can be seen by way of the efficiency improvement 114 between curves 102 and 114 in FIG. 11.

[0102] This allows the increased stagger angle tip section 76 to operate on the choke side of the peak efficiency, as shown by operating point 116 on curve 110, whilst still maintaining an operating efficiency that is the same as, or similar to, optimal point 68 for curve 69.

[0103] The modified blades 25 are mounted/attached to the hub 42 in a conventional array (i.e. with the desired circumferential/angular spacing) so as to provide a fan assembly or fan stage to be used in a gas turbine engine 10 for an aircraft.

[0104] Potential Advantages

[0105] The modified blade can increase the flow range ΔF before which the fan blade will stall. This gives the blade a greater stability margin.

[0106] The modified blade maintains overall fan blade capacity.

[0107] The modified blade reduces the operating pressure ratio of the fan tip sections and increases the operating pressure ratio of the mid-height sections of the blade. This maintains high efficiency.

[0108] The modified blade gives a bird strike benefit due to a lower bird incidence rate. This allows the tip sections to be thinned relative to a conventional blade, e.g. helping the whole engine fan blade out loads, out of balance loads and protecting the fan blade structural forcing thanks to higher blade frequencies. This may contribute to an efficiency increase Δe.sub.2 of the tip sections and/or may offset any efficiency loss due to designing the sections on the choke side of peak efficiency.

[0109] It will be understood that the disclosure 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.