Patent ID: 12196104

The following table lists the reference numerals used in the drawings with the features to which they refer:

Ref no.FeatureFIG.ACore airflow1BBypass airflow1CCamber Line6DPoint4, 5, 7EPoint4, 5, 7FPoint4, 5, 7GPoint4, 5, 7xDistance6A-ALine4, 5TLEThickness of the leading edge6T1avgFirst average value7T2avgSecond average value7T1cFirst constant value7T2CSecond constant value7T1maxFirst maximum value7T2maxSecond maximum value79Principal and rotational axis (of engine)1, 210Gas turbine engine111Core112Air intake114Low pressure compressor115High pressure compressor116Combustion equipment117High pressure turbine118Bypass exhaust nozzle119Low pressure turbine120Core exhaust nozzle121Nacelle122Bypass duct123Propulsive fan1, 224Stationary supporting structure226Shaft1, 227Interconnecting shaft128Sun gear2, 330Epicyclic gearbox1, 2, 332Planet gear2, 334Planet carrier2, 336Linkage238Ring gear2, 340Linkage2100Fan blade4, 5110Aerofoil portion4, 5120Leading edge4, 5, 6124Pressure surface6126Suction surface6130Trailing edge4, 5, 6140Root4, 5144Radial extent/region4, 5, 7145Blade span4, 5146Radial extent/region4, 5, 7147Radial extent/region4, 5, 7148Radial extent/region5, 7149Trailing Edge Span4, 5150Tip4, 5160Platform4, 5, 6170Root Portion5180Tip Portion5190Fixture5200Graph7300Method8310Step8320Step8

DETAILED DESCRIPTION

Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

As used herein, a thickness of an aerofoil section may be defined at a given location on a camber line as a length of a line that is perpendicular to a local direction of the camber line at that location and extends from a pressure surface to a suction surface of the aerofoil section.

Reference to a cross-section through an aerofoil portion at a given percentage along a blade span may mean a section through the aerofoil portion in a plane defined by: a line that passes through a point on a leading edge that is at that percentage along the leading edge from the leading edge root and points in the direction of the tangent to a circumferential direction at that point on the leading edge; and a point on a trailing edge that is at that same percentage along the trailing edge from a trailing edge root.

As referred to herein, a percentage along the leading edge or trailing edge from the root may be, for example, a radial percentage or a spanwise percentage.

Alternatively, reference to a cross-section through an aerofoil portion at a given radial percentage along the blade span may mean a section through the aerofoil that is perpendicular to the radial direction at that radial percentage along the leading edge.

Where reference is made to the axial, radial, and circumferential directions, the skilled person will readily understand this to mean conventional directions when a fan blade is assembled as part of a fan stage or is provided in a gas turbine engine. Viewing the fan blade along a circumferential direction may mean viewing the fan blade in side profile and/or in the meridional plane and/or projected onto a plane defined by the axial and radial directions.

Any fan blade and/or aerofoil portion described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example, at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example, a metal matrix composite and/or an organic matrix composite, such as carbon fibre, and/or from a metal, such as a titanium based metal or an aluminium based material (such as an Aluminium-Lithium alloy) or a steel based material.

As used herein, “at least one of A and B” should be understood to mean “only A, only B, or both A and B.”

FIG.1illustrates a gas turbine engine10having a principal rotational axis9. The engine10comprises an air intake12and a propulsive fan23that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine10comprises a core11that receives the core airflow A. The engine core11comprises, in axial flow series, a low pressure compressor14, a high pressure compressor15, combustion equipment16, a high pressure turbine17, a low pressure turbine19, and a core exhaust nozzle20. A nacelle21surrounds the gas turbine engine10and defines a bypass duct22and a bypass exhaust nozzle18. The bypass airflow B flows through the bypass duct22. The fan23is attached to and driven by the low pressure turbine19via a shaft26and an epicyclic gearbox30.

In use, the core airflow A is accelerated and compressed by the low pressure compressor14and directed into the high pressure compressor15where further compression takes place. The compressed air exhausted from the high pressure compressor15is directed into the combustion equipment16where 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 turbines17,19before being exhausted through the core exhaust nozzle20to provide some propulsive thrust. The high pressure turbine17drives the high pressure compressor by a suitable interconnecting shaft27. The fan23generally provides the majority of the propulsive thrust. The epicyclic gearbox30is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine10is shown inFIG.2. The low pressure turbine19(seeFIG.1) drives the shaft26, which is coupled to a sun wheel, or sun gear,28of the epicyclic gear arrangement30. Radially outwardly of the sun gear28and intermeshing therewith is a plurality of planet gears32that are coupled together by a planet carrier34. The planet carrier34constrains the planet gears32to precess around the sun gear28in synchronicity whilst enabling each planet gear32to rotate about its own axis. The planet carrier34is coupled via linkages36to the fan23in order to drive its rotation about the engine axis9. Radially outwardly of the planet gears32and intermeshing therewith is an annulus or ring gear38that is coupled, via linkages40, to a stationary supporting structure24.

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 fan23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft26with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan23). 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 fan23may be referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox30is shown by way of example in greater detail inFIG.3. Each of the sun gear28, planet gears32and ring gear38comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated inFIG.3. There are four planet gears32illustrated, although it will be apparent to the skilled reader that more or fewer planet gears32may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox30generally comprise at least three planet gears32.

The epicyclic gearbox30illustrated by way of example inFIGS.2and3is of the planetary type, in that the planet carrier34is coupled to an output shaft via linkages36, with the ring gear38fixed. However, any other suitable type of epicyclic gearbox30may be used. By way of further example, the epicyclic gearbox30may be a star arrangement, in which the planet carrier34is held fixed, with the ring (or annulus) gear38allowed to rotate. In such an arrangement the fan23is driven by the ring gear38. By way of further alternative example, the gearbox30may be a differential gearbox in which the ring gear38and the planet carrier34are both allowed to rotate.

It will be appreciated that the arrangement shown inFIGS.2and3is 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 gearbox30in the engine10and/or for connecting the gearbox30to the engine10. By way of further example, the connections (such as the linkages36,40in theFIG.2example) between the gearbox30and other parts of the engine10(such as the input shaft26, the output shaft, and the fixed structure24) 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 ofFIG.2. For example, where the gearbox30has 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 inFIG.2.

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.

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

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 inFIG.1has a split flow nozzle18,20meaning that the flow through the bypass duct22has its own nozzle18that is separate to and radially outside the core exhaust nozzle20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct22and the flow through the core11are 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. 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. In some arrangements, the gas turbine engine10may not comprise a gearbox30.

The geometry of the gas turbine engine10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis9), a radial direction (in the bottom-to-top direction inFIG.1), and a circumferential direction (perpendicular to the page in theFIG.1view). The axial, radial and circumferential directions are mutually perpendicular.

As discussed above, in some embodiments, the gas turbine engine10may include the engine core11including the turbine19, the compressor14, and the core shaft26connecting the turbine19to the compressor14. The gas turbine engine10may further include the fan23located upstream of the engine core11. The fan23may include a plurality of fan blades. The gas turbine engine10may further include the gearbox30that receives an input from the core shaft26and outputs drive to the fan23so as to drive the fan23at a lower rotational speed than the core shaft26. In some embodiments, the turbine may be a first turbine19, the compressor may be a first compressor14, and the core shaft may be a first core shaft26. The engine core11may further include a second turbine17, a second compressor15, and a second core shaft27connecting the second turbine17to the second compressor15. The second turbine17, the second compressor15, and the second core shaft27may be arranged to rotate at a higher rotational speed than the first core shaft26.

FIG.4shows a fan blade100for the gas turbine engine10(shown inFIG.1) in accordance with an embodiment of the present disclosure.

The fan blade100includes an aerofoil portion110. The aerofoil portion110includes a leading edge120and a trailing edge130. The aerofoil portion110extends from a root140to a tip150in a substantially radial spanwise direction. The leading edge120may be defined as a line defined by axially forwardmost points of the aerofoil portion110from the root140to the tip150.

A blade span145is defined as a distance in a radial direction between the leading edge120at the root140and the leading edge120at the tip150. A radius of the leading edge120at the root140may be referred to as a root radius. The radius of the leading edge120at the tip150may be referred to as a tip radius. A trailing edge span149may be defined as a distance in the radial direction between the trailing edge130at the root140and the tip150.

A radial extent144is shown schematically inFIG.4. The radial extent144represents a region at a radius greater than 80% of the blade span145from the root radius, with the point at a radius of 80% of the blade span145from the root radius being labelled as point G. The radial extent144may be interchangeably referred to as “the region144”.

A radial extent147is further shown schematically inFIG.4. The radial extent147represents a region at a radius greater than 70% of the blade span145from the root radius, with a point at a radius of 70% of the blade span145from the root radius being labelled as point F. The radial extent147may be interchangeably referred to as “the region147”.

A radial extent146is further shown schematically inFIG.4. The radial extent146represents a region between a point D at a radius of 50% of the blade span145from the root radius and the point F at the radius of 70% of the blade span145from the root radius. The radial extent146may be interchangeably referred to as “the region146”.

A radial extent148is further shown schematically inFIG.4. The radial extent148represents a region between a point E at a radius of 60% of the blade span145from the root radius and the point F at the radius of 70% of the blade span145from the root radius. The radial extent148may be interchangeably referred to as “the region148”.

A cross section taken along a line A-A through the aerofoil portion110within the radial extent147is shown inFIG.6. The cross-section A-A passes through a point that is greater than 70% of the blade span145from the leading edge120at the root140and a point that is the same percentage of the trailing edge span149from the trailing edge root.

The fan blade100may further include a platform160. The aerofoil portion110may extend directly from the platform160, as shown inFIG.4. Alternatively, as shown inFIG.5, the fan blade100may include a root portion170. The root portion170may extend between the platform160and the root140of the aerofoil portion110. A radial extent of the root portion170may be less than or equal to 7% of the blade span145. In some examples, the radial extent of the root portion170may be less than or equal to 5% of the blade span145.

As shown inFIG.5, the fan blade100may further include a tip portion180. The tip portion180may extend from the tip150of the aerofoil portion110. Specifically, the tip portion180may extend at least radially away from the tip150of the aerofoil portion110. A radial extent of the tip portion180may be less than or equal to 7% of the blade span145. In some examples, the radial extent of the tip portion180may be less than or equal to 5% of the blade span145.

As shown inFIG.5, regardless of the whether the fan blade100includes the root portion170and/or the tip portion180, the blade span145is defined between the root140and the tip150of the aerofoil portion110. Similarly, the regions144,146,147, and148described above in relation toFIG.4are also defined in relation to the blade span145defined between the root140and the tip150, regardless of whether the fan blade100includes the root portion170and/or the tip portion180. The cross-sectional location A-A in the region147is also shown inFIG.5.

As noted above,FIG.6shows the cross-section A-A defined herein. The cross-section includes a camber line C (which may alternatively be referred to as a mean line). The camber line C may be defined as a line formed by the points equidistant from a pressure surface124and a suction surface126of the fan blade100. A distance along the camber line C from the leading edge120is indicated by the letter x inFIG.6. A total length of the camber line C is a length of the dashed line between the leading edge120and the trailing edge130.

A thickness at a given position x along the camber line C may be defined as a length of a line that is perpendicular to the camber line C at the location x and extends from the pressure surface124to the suction surface126. InFIG.6, a thickness of the leading edge120is indicated as TLE. TLE is the thickness of an aerofoil section at a given radius at a position along the camber line C that is 9% of the total length of the camber line C from the leading edge120. This definition is used in order to be sufficiently far from the leading edge120itself to avoid the influence of a curvature of the leading edge120(which may be, for example, an ellipse shape) on the thickness.

FIG.7shows a graph200depicting a variation of the leading edge thickness (TLE) with respect to the blade span145from the root radius of the fan blade100in accordance with an embodiment of the present disclosure.

Referring toFIGS.4to7, for cross-sections through the aerofoil portion110at radii between 50% (indicated by the letter D inFIGS.4,5and7) and 70% (indicated by the letter F inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE includes a first maximum value T1max. In other words, the leading edge thickness TLE has the first maximum value T1max in the region146. The first maximum value T1max may be defined as a maximum value of the leading edge thickness TLE in the region146.

Furthermore, for cross-sections through the aerofoil portion110at radii greater than 70% (indicated by the letter F inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE includes a second maximum value T2max. In other words, the leading edge thickness TLE has the second maximum value T2max in the region147. The second maximum value T2max may be defined as a maximum value of the leading edge thickness TLE in the region147.

The second maximum value T2max is between 105% and 125% of the first maximum value T1max. In other words, the second maximum value T2max is between 1.05 times to 1.25 times of the first maximum value T1max.

The second maximum value T2max being between 105% and 125% of the first maximum value T1max may significantly improve bird strike capability of the fan blade100, with a negligible decrease in an aerodynamic efficiency of the fan blade100. Therefore, the fan blade100having the second maximum value T2max between 105% and 125% of the first maximum value T1max maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

In some embodiments, for cross-sections through the aerofoil portion at radii between 60% (indicated by the letter E inFIGS.4,5, and7) and 70% (indicated by the letter F inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE may include a first average value T1avg. The first average value T1avg may be defined as an average value of the leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii between 60% and 70% of the blade span145from the root radius. In other words, the first average value T1 avg may be the average value of the leading edge thickness TLE in the region148.

Further, for cross-sections through the aerofoil portion at radii greater than 80% (indicated by the letter G inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE includes a second average value T2avg. The second average value T2avg may be defined as an average value of the leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii greater than 80% of the blade span145from the root radius. In other words, the second average value T2avg may be the average value of the leading edge thickness TLE in the region144.

The second average value T2avg may be between 105% and 125% of the first average value T1avg. The second average value T2avg being between 105% and 125% of the first average value T1 avg may significantly improve bird strike capability of the fan blade100, with a negligible decrease in an aerodynamic efficiency of the fan blade100. Therefore, the fan blade100having the second average value T2avg being between 105% and 125% of the first average value T1avg may maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

In some embodiments, for all cross-sections through the aerofoil portion110at radii between 60% (indicated by the letter E inFIGS.4,5, and7) and 70% (indicated by the letter F inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE may include a first constant value T1c. In other words, the leading edge thickness TLE may have the first constant value T1c in the region148. In some embodiments, as shown in the graph200ofFIG.7, the first constant value T1 c may be equal to the first maximum value T1max. However, in some other embodiments, the first constant value T1c may be different from the first maximum value T1max. Further, in some embodiments, as shown in the graph200ofFIG.7, the first constant value T1 c may be equal to the first average value T1 avg.

Furthermore, for all cross-sections through the aerofoil portion110at radii greater than 80% (indicated by the letter G inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE may include a second constant value T2c. In other words, the leading edge thickness TLE may have the second constant value T2c in the region144. In some embodiments, as shown in the graph200ofFIG.7, the second constant value T2c may be equal to the second maximum value T1max. However, in some other embodiments, the second constant value T2c may be different from the second maximum value T2max. Further, in some embodiments, as shown in the graph200ofFIG.7, the second constant value T2c may be equal to the second average value T2avg.

The second constant value T2c may be between 105% and 125% of the first constant value T1c. The second constant value T2c being between 105% and 125% of the first constant value T1 c may significantly improve bird strike capability of the fan blade100, with a negligible decrease in an aerodynamic efficiency of the fan blade100. Therefore, the fan blade100having the second constant value T2c being between 105% and 125% of the first constant value T1 c may maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

In some embodiments, as shown inFIG.7, for all cross-sections through the aerofoil portion110at radii greater than 80% (indicated by the letter G inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE may be greater than 105% and less than 125% of the leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii between 60% (indicated by the letter E inFIGS.4,5, and7) and 70% (indicated by the letter F inFIGS.4,5, and7) of the blade span145from the root radius. In other words, the leading edge thickness TLE at each cross-section through the aerofoil portion110in the region144may be greater than 105% and less than 125% of the leading edge thickness TLE at each cross-section through the aerofoil portion110in the region148.

The leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii greater than 80% of the blade span145from the root radius being greater than 105% and less than 125% of the leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii between 60% and 70% of the blade span145from the root radius may significantly improve bird strike capability of the fan blade100, with a negligible decrease in an aerodynamic efficiency of the fan blade100. Therefore, the fan blade100having such a configuration may maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

In some embodiments, for cross-sections through the aerofoil portion110at radii between 70% (indicated by the letter F inFIGS.4,5, and7) and 80% (indicated by the letter G inFIGS.4,5, and7) of the blade span145from the root radius, the leading edge thickness TLE may increase linearly with respect to the blade span145. For example, as shown inFIG.7, the leading edge thickness TLE may increase linearly from 70% to 80% of the blade span145from the root radius.

The leading edge thickness TLE increasing linearly with respect to the blade span145for cross-sections through the aerofoil portion110at radii between 70% and 80% from the root radius may improve the aerodynamic performance of the fan blade100as compared to an abrupt change in the leading edge thickness TLE. Furthermore, the linear increase in the leading edge thickness TLE for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span from the root radius may minimally impact the aerodynamic performance of the fan blade100.

In some embodiments, for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span145from the root radius, the leading edge thickness TLE may increase linearly from the first maximum value T1max to the second maximum value T2max. In some embodiments, for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span145from the root radius, the leading edge thickness TLE may increase linearly from the first average value T1 avg to the second average value T2avg. In some embodiments, for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span145from the root radius, the leading edge thickness TLE may increase linearly from the first constant value T1 c to the second constant value T2c.

In some embodiments, the first maximum value T1max, the first average value T1avg, and the first constant value T1c are equal to each other. That is, in some embodiments, T1max=T1 avg=T1c.

In some embodiments, the second maximum value T2max, the second average value T2avg, and the second constant value T2c are equal to each other. That is, in some embodiments, T2max=T2avg=T2c.

It will be appreciated that the geometry represented by the graph200inFIG.7is exemplary only, and a great many other geometries are possible in accordance with the present disclosure.

As discussed above with reference toFIG.1, the gas turbine engine10includes the fan23including the plurality of fan blades. The plurality of fan blades may include the fan blade100.

The fan blade100may be attached to a hub in any desired manner. For example, the fan blade100may include a fixture190, such as that shown inFIG.5, that 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 blade100to the hub/disc.

Alternatively, the fan blade100and the hub may be formed as a unitary part, with no mechanical and/or releasable connections, so as to form a unitary fan stage. Such a unitary fan stage may be referred to as a “blisk”. Such a unitary fan stage may be manufactured in any suitable manner, for example by machining and/or by linear friction welding the fan blades100to the hub, or at least linear friction welding the aerofoil portions110to a hub that includes radially inner stub portions of the fan blades100.

FIG.8shows a flowchart depicting various steps of a method300of minimising an impact of a bird strike on a fan blade (e.g., the fan blade100ofFIGS.4and5) of a gas turbine engine (e.g., the gas turbine engine10ofFIG.1). The method300may also be a method of designing the fan blade100. The method300will be described with additional reference toFIGS.4to7.

The fan blade has an aerofoil portion having a leading edge extending from a root to a tip. A distance between the leading edge at the root and the leading edge at the tip defines a blade span. A leading edge thickness is defined as a thickness of a cross-section at a given radius at a location along a camber line that is 9% of the total length of the camber line from the leading edge.

At step310, the method300includes providing a first maximum value of the leading edge thickness for cross-sections through the aerofoil portion at radii between 50% and 70% of the blade span from a root radius. For example, the method300may include providing the first maximum value T1max of the leading edge thickness TLE for cross-sections through the aerofoil portion110at radii between 50% and 70% of the blade span145from the root radius. In other words, the method300may include providing the first maximum value T1max of the leading edge thickness TLE for cross-sections through the aerofoil portion110in the region146.

At step320, the method300further includes providing a second maximum value of the leading edge thickness for cross-sections through the aerofoil portion at radii greater than 80% of the blade span from the root radius. The second maximum value is between 105% and 125% of the first maximum value. For example, the method300may include providing the second maximum value T2max of the leading edge thickness TLE for cross-sections through the aerofoil portion110at radii greater than 80% of the blade span145from the root radius. In other words, the method300may include providing the second maximum value T2max of the leading edge thickness TLE for cross-sections through the aerofoil portion110in the region144.

In some embodiments, the method300may further include providing a first average value of the leading edge thickness for cross-sections through the aerofoil portion at radii between 60% and 70% of the blade span from the root radius. For example, the method300may include providing the first average value T1avg of the leading edge thickness TLE for cross-sections through the aerofoil portion110at radii between 60% and 70% of the blade span145from the root radius. In other words, the method300may include providing the first average value T1 avg of the leading edge thickness TLE for cross-sections through the aerofoil portion110in the region148.

In some embodiments, the method300may further include providing a second average value of the leading edge thickness for cross-sections through the aerofoil portion at radii greater than 80% of the blade span from the root radius. The second average value is between 105% and 125% of the first average value. For example, the method300may further include providing the second average value T2avg of the leading edge thickness TLE for cross-sections through the aerofoil portion110at radii greater than 80% of the blade span145from the root radius. In other words, the method300may further include providing the second average value T2avg of the leading edge thickness TLE for cross-sections through the aerofoil portion110in the region144.

In some embodiments, the method300may further include providing a first constant value of the leading edge thickness for all cross-sections through the aerofoil portion at radii between 60% and 70% of the blade span from the root radius. For example, the method300may include providing the first constant value T1 c of the leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii between 60% and 70% of the blade span145from the root radius. In other words, the method300may include providing the first constant value T1c of the leading edge thickness TLE for all cross-sections through the aerofoil portion110in the region148.

In some embodiments, the method300may further include providing a second constant value for all cross-sections through the aerofoil portion at radii greater than 80% of the blade span from the root radius. The second constant value is between 105% and 125% of the first constant value. For example, the method300may include providing the second constant value T2c for all cross-sections through the aerofoil portion110at radii greater than 80% of the blade span145from the root radius. In other words, the method300may include providing the second constant value T2c for all cross-sections through the aerofoil portion110in the region144.

In some embodiments, the method300may further include providing the leading edge thickness for all cross-sections through the aerofoil portion at radii greater than 80% of the blade span from the root radius greater than 105% and less than 125% of the leading edge thickness for all cross-sections through the aerofoil portion at radii between 60% and 70% of the blade span from the root radius. For example, the method300may include providing the leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii greater than 80% of the blade span145from the root radius greater than 105% and less than 125% of the leading edge thickness TLE for all cross-sections through the aerofoil portion110at radii between 60% and 70% of the blade span145from the root radius. As a result, the leading edge thickness TLE for all cross-sections through the aerofoil portion110in the region144may be greater than 105% and less than 125% of the leading edge thickness TLE for all cross-sections through the aerofoil portion110in the region148.

In some embodiments, the method300may further include increasing the leading edge thickness linearly with respect to the blade span for cross-sections through the aerofoil portion at radii between 70% and 80% of the blade span from the root radius. For example, the method300may include increasing the leading edge thickness TLE linearly with respect to the blade span145for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span145from the root radius.

As an example, in some embodiments, the method300may include increasing the leading edge thickness TLE linearly from the first maximum value T1max to the second maximum value T2max for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span145from the root radius. In some embodiments, the method300may include increasing the leading edge thickness TLE linearly from the first average value T1 avg to the second average value T2avg for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span145from the root radius. In some embodiments, the method300may include increasing the leading edge thickness TLE linearly from the first constant value T1c to the second constant value T2c for cross-sections through the aerofoil portion110at radii between 70% and 80% of the blade span145from the root radius.

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.