Patent Description:
A gas turbine engine includes a fan having fan blades in front of the gas turbine engine. The fan may produce a substantial amount of engine thrust, for example, about <NUM>% of an overall engine thrust. Thus, an aerodynamic efficiency of the fan blades may have a significant impact on overall engine efficiency, and therefore a significant impact on fuel burn. The aerodynamic efficiency of the fan blades may therefore be of critical importance when designing the fan blades.

During operation, the gas turbine engine may encounter a bird strike, in which a bird, an animal, or a foreign object may collide with the fan blades. Therefore, the fan blades of the gas turbine engine may need to be capable of withstanding an impact from such bird strikes. In other words, a fan blade may need to be designed to be able to retain its structural integrity upon experiencing the bird strike. However, increasing impact withstanding capability of the fan blades often comes at a cost of a severe decrease in the aerodynamic efficiency of the fan blades.

United States patent application <CIT> discloses a gas turbine engine fan blade that has an aerofoil portion for which, at radii between <NUM> percent and <NUM> percent of the blade span, the location of the position of maximum thickness along the camber line is at less than a defined percentage of the total length of the camber line. For all cross-sections through the aerofoil portion at radii greater than <NUM> percent of the blade span, the location of the position of maximum thickness along the camber line is at more than a defined percentage of the total length of the camber line. The geometry of the fan blade may result in a lower susceptibility to flutter.

There remains a need of a fan blade that has improved impact withstanding capability without a significant decrease in its aerodynamic efficiency.

According to a first aspect there is provided a fan blade for a gas turbine engine as set out in claim <NUM>.

The second maximum value of the leading edge thickness being between <NUM>% and <NUM>% of the first maximum value of the leading edge thickness may significantly improve bird strike capability of the fan blade, with a negligible decrease in an aerodynamic efficiency of the fan blade. Therefore, the fan blade having the second maximum value between <NUM>% and <NUM>% of the first maximum value may maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

Using computational fluid dynamics (CFD), it has been found that the fan blade having the second maximum value of the leading edge thickness between <NUM>% and <NUM>% of the first maximum value of the leading edge thickness may present a significant improvement in impact capability, while causing a very limited decrease in the aerodynamic efficiency.

In some embodiments, for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius, the leading edge thickness includes a first average value. For cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius, the leading edge thickness includes a second average value. The second average value is between <NUM>% and <NUM>% of the first average value.

The second average value of the leading edge thickness being between <NUM>% and <NUM>% of the first average value of the leading edge thickness may significantly improve bird strike capability of the fan blade, with a negligible decrease in an aerodynamic efficiency of the fan blade. Therefore, the fan blade having the second average value between <NUM>% and <NUM>% of the first average value may maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

In some embodiments, the first average value may be equal to the first maximum value and the second average value may be equal to the second maximum value.

In some embodiments, for all cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius, the leading edge thickness includes a first constant value. For all cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius, the leading edge thickness includes a second constant value. The second constant value is between <NUM>% and <NUM>% of the first constant value.

The second constant value of the leading edge thickness being between <NUM>% and <NUM>% of the first constant value of the leading edge thickness may significantly improve bird strike capability of the fan blade, with a negligible decrease in an aerodynamic efficiency of the fan blade. Therefore, the fan blade having the second constant value between <NUM>% and <NUM>% of the first constant value may maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

In some embodiments, the first constant value may be equal to the first maximum value and the second constant value may be equal to the second maximum value.

In some embodiments, for all cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius, the leading edge thickness is greater than <NUM>% and less than <NUM>% of the leading edge thickness for all cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius.

The leading edge thickness for all cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius being greater than <NUM>% and less than <NUM>% of the leading edge thickness for all cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius may significantly improve bird strike capability of the fan blade, with a negligible decrease in an aerodynamic efficiency of the fan blade. Therefore, the fan blade having 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 portion at radii between <NUM>% and <NUM>% of the blade span from the root radius, the leading edge thickness increases linearly with respect to the blade span.

The leading edge thickness increasing linearly with respect to the blade span for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% may improve the aerodynamic performance of the fan blade as compared to an abrupt change in the leading edge thickness. Furthermore, the linear increase in the leading edge thickness for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius may minimally impact the aerodynamic performance of the fan blade.

In some embodiments, for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius, the leading edge thickness may increase linearly from the first maximum value to the second maximum value. In some embodiments, for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius, the leading edge thickness may increase linearly from the first average value to the second average value. In some embodiments, for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius, the leading edge thickness may increase linearly from the first constant value to the second constant value.

In some embodiments, the first maximum value, the first average value, and the first constant value are equal to each other.

In some embodiments, the second maximum value, the second average value, and the second constant value are equal to each other.

In some embodiments, the fan blade further includes a platform and a root portion. The root portion extends between the platform and the root of the aerofoil portion.

In some embodiments, a radial extent of the root portion is less than or equal to <NUM>% of the blade span.

In some embodiments, the fan blade further includes a tip portion that extends at least radially away from the tip of the aerofoil portion.

In some embodiments, a radial extent of the tip portion is less than or equal to <NUM>% of the blade span.

According to a second aspect there is provided a gas turbine engine for an aircraft as set out in claim <NUM>.

In some embodiments, the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft. The engine core further includes a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

According to a third aspect there is provided a method of minimising an impact of a bird strike on a fan blade of a gas turbine engine as set out in claim <NUM>.

The method may significantly improve bird strike capability of the fan blade, with a negligible decrease in an aerodynamic efficiency of the fan blade. The method may allow the fan blade to maintain its structural integrity upon experiencing a bird strike while having excellent aerodynamic efficiency.

In some embodiments, the method further includes providing a first average value of the leading edge thickness for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius. The method further includes providing a second average value of the leading edge thickness for cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius. The second average value is between <NUM>% and <NUM>% of the first average value.

In some embodiments, the method further includes providing a first constant value of the leading edge thickness for all cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius. The method further includes providing a second constant value for all cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius. The second constant value is between <NUM>% and <NUM>% of the first constant value.

In some embodiments, the method further includes providing the leading edge thickness for all cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius greater than <NUM>% and less than <NUM>% of the leading edge thickness for all cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius.

In some embodiments, the method further includes increasing the leading edge thickness linearly with respect to the blade span for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than <NUM>, for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>, for example on the order of or at least <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from <NUM> or <NUM> to <NUM>. In some arrangements, the gear ratio may be outside these ranges.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> to <NUM>, <NUM> to <NUM>, or <NUM> to <NUM>. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s or <NUM> Nkg-<NUM>s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e., the values may form upper or lower bounds), for example in the range of from <NUM> Nkg-<NUM>s to <NUM> Nkg-<NUM>s, or <NUM> Nkg-<NUM>s to <NUM> Nkg-<NUM>s. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A fan blade and/or aerofoil portion of a fan blade 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.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> fan blades.

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

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.

The engine core <NUM> comprises, in axial flow series, a low pressure compressor <NUM>, a high pressure compressor <NUM>, combustion equipment <NUM>, a high pressure turbine <NUM>, a low pressure turbine <NUM>, and a core exhaust nozzle <NUM>.

The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines <NUM>, <NUM> before being exhausted through the core exhaust nozzle <NUM> to provide some propulsive thrust.

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 <NUM>) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft <NUM> with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan <NUM>).

By way of further example, the connections (such as the linkages <NUM>, <NUM> in the <FIG> example) between the gearbox <NUM> and other parts of the engine <NUM> (such as the input shaft <NUM>, the output shaft, and the fixed structure <NUM>) 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>.

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 in <FIG> has a split flow nozzle <NUM>, <NUM> meaning that the flow through the bypass duct <NUM> has its own nozzle <NUM> that is separate to and radially outside the core exhaust nozzle <NUM>. 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 <NUM> and the flow through the core <NUM> 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. 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 engine <NUM> may not comprise a gearbox <NUM>.

As discussed above, in some embodiments, the gas turbine engine <NUM> may include the engine core <NUM> including the turbine <NUM>, the compressor <NUM>, and the core shaft <NUM> connecting the turbine <NUM> to the compressor <NUM>. The gas turbine engine <NUM> may further include the fan <NUM> located upstream of the engine core <NUM>. The fan <NUM> may include a plurality of fan blades. The gas turbine engine <NUM> may further include the gearbox <NUM> that receives an input from the core shaft <NUM> and outputs drive to the fan <NUM> so as to drive the fan <NUM> at a lower rotational speed than the core shaft <NUM>. In some embodiments, the turbine may be a first turbine <NUM>, the compressor may be a first compressor <NUM>, and the core shaft may be a first core shaft <NUM>. The engine core <NUM> may further include a second turbine <NUM>, a second compressor <NUM>, and a second core shaft <NUM> connecting the second turbine <NUM> to the second compressor <NUM>. The second turbine <NUM>, the second compressor <NUM>, and the second core shaft <NUM> may be arranged to rotate at a higher rotational speed than the first core shaft <NUM>.

<FIG> shows a fan blade <NUM> for the gas turbine engine <NUM> (shown in <FIG>) in accordance with an embodiment of the present disclosure.

The fan blade <NUM> includes an aerofoil portion <NUM>. The aerofoil portion <NUM> includes a leading edge <NUM> and a trailing edge <NUM>. The aerofoil portion <NUM> extends from a root <NUM> to a tip <NUM> in a substantially radial spanwise direction. The leading edge <NUM> may be defined as a line defined by axially forwardmost points of the aerofoil portion <NUM> from the root <NUM> to the tip <NUM>.

A blade span <NUM> is defined as a distance in a radial direction between the leading edge <NUM> at the root <NUM> and the leading edge <NUM> at the tip <NUM>. A radius of the leading edge <NUM> at the root <NUM> may be referred to as a root radius. The radius of the leading edge <NUM> at the tip <NUM> may be referred to as a tip radius. A trailing edge span <NUM> may be defined as a distance in the radial direction between the trailing edge <NUM> at the root <NUM> and the tip <NUM>.

A radial extent <NUM> is shown schematically in <FIG>. The radial extent <NUM> represents a region at a radius greater than <NUM>% of the blade span <NUM> from the root radius, with the point at a radius of <NUM>% of the blade span <NUM> from the root radius being labelled as point G. The radial extent <NUM> may be interchangeably referred to as "the region <NUM>".

A radial extent <NUM> is further shown schematically in <FIG>. The radial extent <NUM> represents a region at a radius greater than <NUM>% of the blade span <NUM> from the root radius, with a point at a radius of <NUM>% of the blade span <NUM> from the root radius being labelled as point F. The radial extent <NUM> may be interchangeably referred to as "the region <NUM>".

A radial extent <NUM> is further shown schematically in <FIG>. The radial extent <NUM> represents a region between a point D at a radius of <NUM>% of the blade span <NUM> from the root radius and the point F at the radius of <NUM>% of the blade span <NUM> from the root radius. The radial extent <NUM> may be interchangeably referred to as "the region <NUM>".

A radial extent <NUM> is further shown schematically in <FIG>. The radial extent <NUM> represents a region between a point E at a radius of <NUM>% of the blade span <NUM> from the root radius and the point F at the radius of <NUM>% of the blade span <NUM> from the root radius. The radial extent <NUM> may be interchangeably referred to as "the region <NUM>".

A cross section taken along a line A-A through the aerofoil portion <NUM> within the radial extent <NUM> is shown in <FIG>. The cross-section A-A passes through a point that is greater than <NUM>% of the blade span <NUM> from the leading edge <NUM> at the root <NUM> and a point that is the same percentage of the trailing edge span <NUM> from the trailing edge root.

The fan blade <NUM> may further include a platform <NUM>. The aerofoil portion <NUM> may extend directly from the platform <NUM>, as shown in <FIG>. Alternatively, as shown in <FIG>, the fan blade <NUM> may include a root portion <NUM>. The root portion <NUM> may extend between the platform <NUM> and the root <NUM> of the aerofoil portion <NUM>. A radial extent of the root portion <NUM> may be less than or equal to <NUM>% of the blade span <NUM>. In some examples, the radial extent of the root portion <NUM> may be less than or equal to <NUM>% of the blade span <NUM>.

As shown in <FIG>, the fan blade <NUM> may further include a tip portion <NUM>. The tip portion <NUM> may extend from the tip <NUM> of the aerofoil portion <NUM>. Specifically, the tip portion <NUM> may extend at least radially away from the tip <NUM> of the aerofoil portion <NUM>. A radial extent of the tip portion <NUM> may be less than or equal to <NUM>% of the blade span <NUM>. In some examples, the radial extent of the tip portion <NUM> may be less than or equal to <NUM>% of the blade span <NUM>.

As shown in <FIG>, regardless of the whether the fan blade <NUM> includes the root portion <NUM> and/or the tip portion <NUM>, the blade span <NUM> is defined between the root <NUM> and the tip <NUM> of the aerofoil portion <NUM>. Similarly, the regions <NUM>, <NUM>, <NUM>, and <NUM> described above in relation to <FIG> are also defined in relation to the blade span <NUM> defined between the root <NUM> and the tip <NUM>, regardless of whether the fan blade <NUM> includes the root portion <NUM> and/or the tip portion <NUM>. The cross-sectional location A-A in the region <NUM> is also shown in <FIG>.

As noted above, <FIG> shows 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 surface <NUM> and a suction surface <NUM> of the fan blade <NUM>. A distance along the camber line C from the leading edge <NUM> is indicated by the letter x in <FIG>. A total length of the camber line C is a length of the dashed line between the leading edge <NUM> and the trailing edge <NUM>.

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 surface <NUM> to the suction surface <NUM>. In <FIG>, a thickness of the leading edge <NUM> is 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 <NUM>% of the total length of the camber line C from the leading edge <NUM>. This definition is used in order to be sufficiently far from the leading edge <NUM> itself to avoid the influence of a curvature of the leading edge <NUM> (which may be, for example, an ellipse shape) on the thickness.

<FIG> shows a graph <NUM> depicting a variation of the leading edge thickness (TLE) with respect to the blade span <NUM> from the root radius of the fan blade <NUM> in accordance with an embodiment of the present disclosure.

Referring to <FIG>, for cross-sections through the aerofoil portion <NUM> at radii between <NUM>% (indicated by the letter D in <FIG>, <FIG> and <FIG>) and <NUM>% (indicated by the letter F in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from 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 region <NUM>. The first maximum value T1max may be defined as a maximum value of the leading edge thickness TLE in the region <NUM>.

Furthermore, for cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% (indicated by the letter F in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from 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 region <NUM>. The second maximum value T2max may be defined as a maximum value of the leading edge thickness TLE in the region <NUM>.

The second maximum value T2max is between <NUM>% and <NUM>% of the first maximum value T1max. In other words, the second maximum value T2max is between <NUM> times to <NUM> times of the first maximum value T1 max.

The second maximum value T2max being between <NUM>% and <NUM>% of the first maximum value T1max may significantly improve bird strike capability of the fan blade <NUM>, with a negligible decrease in an aerodynamic efficiency of the fan blade <NUM>. Therefore, the fan blade <NUM> having the second maximum value T2max between <NUM>% and <NUM>% of the first maximum value T1 max 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 <NUM>% (indicated by the letter E in <FIG>, <FIG>, and <FIG>) and <NUM>% (indicated by the letter F in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from the root radius, the leading edge thickness TLE may include a first average value T1 avg. 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 portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius. In other words, the first average value T1avg may be the average value of the leading edge thickness TLE in the region <NUM>.

Further, for cross-sections through the aerofoil portion at radii greater than <NUM>% (indicated by the letter G in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from 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 portion <NUM> at radii greater than <NUM>% of the blade span <NUM> from the root radius. In other words, the second average value T2avg may be the average value of the leading edge thickness TLE in the region <NUM>.

The second average value T2avg may be between <NUM>% and <NUM>% of the first average value T1avg. The second average value T2avg being between <NUM>% and <NUM>% of the first average value T1avg may significantly improve bird strike capability of the fan blade <NUM>, with a negligible decrease in an aerodynamic efficiency of the fan blade <NUM>. Therefore, the fan blade <NUM> having the second average value T2avg being between <NUM>% and <NUM>% 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 portion <NUM> at radii between <NUM>% (indicated by the letter E in <FIG>, <FIG>, and <FIG>) and <NUM>% (indicated by the letter F in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from 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 region <NUM>. In some embodiments, as shown in the graph <NUM> of <FIG>, the first constant value T1c 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 graph <NUM> of <FIG>, the first constant value T1c may be equal to the first average value T1avg.

Furthermore, for all cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% (indicated by the letter G in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from 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 region <NUM>. In some embodiments, as shown in the graph <NUM> of <FIG>, 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 graph <NUM> of <FIG>, the second constant value T2c may be equal to the second average value T2avg.

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

In some embodiments, as shown in <FIG>, for all cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% (indicated by the letter G in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from the root radius, the leading edge thickness TLE may be greater than <NUM>% and less than <NUM>% of the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> at radii between <NUM>% (indicated by the letter E in <FIG>, <FIG>, and <FIG>) and <NUM>% (indicated by the letter F in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from the root radius. In other words, the leading edge thickness TLE at each cross-section through the aerofoil portion <NUM> in the region <NUM> may be greater than <NUM>% and less than <NUM>% of the leading edge thickness TLE at each cross-section through the aerofoil portion <NUM> in the region <NUM>.

The leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% of the blade span <NUM> from the root radius being greater than <NUM>% and less than <NUM>% of the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius may significantly improve bird strike capability of the fan blade <NUM>, with a negligible decrease in an aerodynamic efficiency of the fan blade <NUM>. Therefore, the fan blade <NUM> having 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 portion <NUM> at radii between <NUM>% (indicated by the letter F in <FIG>, <FIG>, and <FIG>) and <NUM>% (indicated by the letter G in <FIG>, <FIG>, and <FIG>) of the blade span <NUM> from the root radius, the leading edge thickness TLE may increase linearly with respect to the blade span <NUM>. For example, as shown in <FIG>, the leading edge thickness TLE may increase linearly from <NUM>% to <NUM>% of the blade span <NUM> from the root radius.

The leading edge thickness TLE increasing linearly with respect to the blade span <NUM> for cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% from the root radius may improve the aerodynamic performance of the fan blade <NUM> as 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 portion <NUM> at radii between <NUM>% and <NUM>% of the blade span from the root radius may minimally impact the aerodynamic performance of the fan blade <NUM>.

In some embodiments, for cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from 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 portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius, the leading edge thickness TLE may increase linearly from the first average value T1avg to the second average value T2avg. In some embodiments, for cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius, the leading edge thickness TLE may increase linearly from the first constant value T1c 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 = T1avg = 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 graph <NUM> in <FIG> is exemplary only, and a great many other geometries are possible in accordance with the present disclosure.

As discussed above with reference to <FIG>, the gas turbine engine <NUM> includes the fan <NUM> including the plurality of fan blades. The plurality of fan blades may include the fan blade <NUM>.

The fan blade <NUM> may be attached to a hub in any desired manner. For example, the fan blade <NUM> may include a fixture <NUM>, such as that shown in <FIG>, 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 blade <NUM> to the hub/disc.

Alternatively, the fan blade <NUM> and 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 blades <NUM> to the hub, or at least linear friction welding the aerofoil portions <NUM> to a hub that includes radially inner stub portions of the fan blades <NUM>.

<FIG> shows a flowchart depicting various steps of a method <NUM> of minimising an impact of a bird strike on a fan blade (e.g., the fan blade <NUM> of <FIG> and <FIG>) of a gas turbine engine (e.g., the gas turbine engine <NUM> of <FIG>). The method <NUM> may also be a method of designing the fan blade <NUM>. The method <NUM> will be described with additional reference to <FIG>.

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 <NUM>% of the total length of the camber line from the leading edge.

At step <NUM>, the method <NUM> includes providing a first maximum value of the leading edge thickness for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from a root radius. For example, the method <NUM> may include providing the first maximum value T1max of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius. In other words, the method <NUM> may include providing the first maximum value T1max of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> in the region <NUM>.

At step <NUM>, the method <NUM> further includes providing a second maximum value of the leading edge thickness for cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius. The second maximum value is between <NUM>% and <NUM>% of the first maximum value. For example, the method <NUM> may include providing the second maximum value T2max of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% of the blade span <NUM> from the root radius. In other words, the method <NUM> may include providing the second maximum value T2max of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> in the region <NUM>.

In some embodiments, the method <NUM> may further include providing a first average value of the leading edge thickness for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius. For example, the method <NUM> may include providing the first average value T1avg of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius. In other words, the method <NUM> may include providing the first average value T1avg of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> in the region <NUM>.

In some embodiments, the method <NUM> may further include providing a second average value of the leading edge thickness for cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius. The second average value is between <NUM>% and <NUM>% of the first average value. For example, the method <NUM> may further include providing the second average value T2avg of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% of the blade span <NUM> from the root radius. In other words, the method <NUM> may further include providing the second average value T2avg of the leading edge thickness TLE for cross-sections through the aerofoil portion <NUM> in the region <NUM>.

In some embodiments, the method <NUM> may further include providing a first constant value of the leading edge thickness for all cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius. For example, the method <NUM> may include providing the first constant value T1c of the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius. In other words, the method <NUM> may include providing the first constant value T1c of the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> in the region <NUM>.

In some embodiments, the method <NUM> may further include providing a second constant value for all cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius. The second constant value is between <NUM>% and <NUM>% of the first constant value. For example, the method <NUM> may include providing the second constant value T2c for all cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% of the blade span <NUM> from the root radius. In other words, the method <NUM> may include providing the second constant value T2c for all cross-sections through the aerofoil portion <NUM> in the region <NUM>.

In some embodiments, the method <NUM> may further include providing the leading edge thickness for all cross-sections through the aerofoil portion at radii greater than <NUM>% of the blade span from the root radius greater than <NUM>% and less than <NUM>% of the leading edge thickness for all cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius. For example, the method <NUM> may include providing the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> at radii greater than <NUM>% of the blade span <NUM> from the root radius greater than <NUM>% and less than <NUM>% of the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius. As a result, the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> in the region <NUM> may be greater than <NUM>% and less than <NUM>% of the leading edge thickness TLE for all cross-sections through the aerofoil portion <NUM> in the region <NUM>.

In some embodiments, the method <NUM> may further include increasing the leading edge thickness linearly with respect to the blade span for cross-sections through the aerofoil portion at radii between <NUM>% and <NUM>% of the blade span from the root radius. For example, the method <NUM> may include increasing the leading edge thickness TLE linearly with respect to the blade span <NUM> for cross-sections through the aerofoil portion <NUM> at radii between <NUM>% and <NUM>% of the blade span <NUM> from the root radius.

Claim 1:
A fan blade (<NUM>) for a gas turbine engine (<NUM>), the fan blade (<NUM>) comprising:
an aerofoil portion (<NUM>) comprising a leading edge (<NUM>) extending from a root (<NUM>) to a tip (<NUM>), a distance between the leading edge (<NUM>) at the root (<NUM>) and the leading edge (<NUM>) at the tip (<NUM>) defining a blade span (<NUM>), wherein:
a leading edge thickness (TLE) is defined as a thickness of a cross-section at a given radius at a location along a camber line (C) that is <NUM>% of the total length of the camber line (C) from the leading edge (<NUM>);
for cross-sections through the aerofoil portion (<NUM>) at radii between <NUM>% and <NUM>% of the blade span (<NUM>) from a root radius, the leading edge thickness (TLE) comprises a first maximum value (T1max);
for cross-sections through the aerofoil portion (<NUM>) at radii greater than <NUM>% of the blade span (<NUM>) from the root radius, the leading edge thickness (TLE) comprises a second maximum value (T2max); and
the second maximum value (T2max) is between <NUM>% and <NUM>% of the first maximum value (T1max).