Gas turbine engine with optimized fan blade geometry

Abstract

A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Claims

1. A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: each fan blade comprises cross-sections, including a first cross-section, the cross-sections comprising: a camber line defined by a midpoint between its pressure surface and its suction surface, with a true chord (C) being defined as a distance along the camber line between a leading edge and a trailing edge of the fan blade; a covered passage defined as a portion of the cross-section between the trailing edge and a line (J) passing through a point (K) on the suction surface that is closest to a leading edge of a neighbouring fan blade; a covered passage length (P) defined as a distance along the camber line that is in the covered passage; and a covered passage percentage defined as the covered passage length (P) as a percentage of the true chord (C); the first cross-section of each fan blade having the covered passage percentage between 40 percent and 70 percent, wherein the first cross-section defines a change in angle in degrees between a line tangent to the camber line at the leading edge and a reference line (α1) and a line tangent to a point on the camber line that corresponds to a start of the covered passage and the reference line (α2) that satisfies: 20>|α2−α1| and wherein: each fan blade has a radial span extending from a root at a 0 percent span position to a tip at a 100 percent span position, and a ratio of a radius of the fan blade at the root (r.sub.root) to the radius of the fan blade at the tip (r.sub.tip) is less than 0.3; and a specific thrust at a forward Mach number of 0.8 is less than 90 NKg.sup.−1s.

2. The gas turbine engine of claim 1, wherein the change in angle in degrees of the camber line of the first section between the line tangent to the camber line at the leading edge and the reference line (α.sub.1) and the point on the camber line that corresponds to the start of the covered passage and the reference line (α.sub.2) satisfies: $.Math.\mathrm{\alpha 2}-\mathrm{\alpha 1}.Math.\ge 6+\frac{\left(\mathrm{covered}\mathrm{passage}\mathrm{percentage}-35\right)}{9}.$

3. The gas turbine engine of claim 1, wherein the cross-sections through each fan blade for which the covered passage percentage is between 40 percent and 70 percent, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies: $.Math.\mathrm{\alpha 2}-\mathrm{\alpha 1}.Math.\ge 7+\frac{\left(\mathrm{covered}\mathrm{passage}\mathrm{percentage}-35\right)}{9}.$

4. The gas turbine engine of claim 1, wherein the cross-sections through each fan blade for which the covered passage percentage is between 35 percent and 80 percent, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies: $.Math.\mathrm{\alpha 2}-\mathrm{\alpha 1}.Math.\ge 6+\frac{\left(\mathrm{covered}\mathrm{passage}\mathrm{percentage}-35\right)}{9}.$

5. The gas turbine engine of claim 1, wherein the cross-sections through each fan blade for which the covered passage percentage is between 35 percent and 80 percent, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies: $.Math.\mathrm{\alpha 2}-\mathrm{\alpha 1}.Math.\ge 7+\frac{\left(\mathrm{covered}\mathrm{passage}\mathrm{percentage}-35\right)}{9}.$

6. The gas turbine engine of claim 1, wherein the cross-sections through each fan blade for which the covered passage percentage is between 60 percent and 62 percent, the change in angle in degrees of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
|α2−α1|≥9.

7. The gas turbine engine of claim 1, wherein: the ratio of the radius of the fan blade at the root (r.sub.root) to the radius of the fan blade at the tip (r.sub.tip) is less than 0.29.

8. The gas turbine engine of claim 1, wherein: the ratio of the radius of the fan blade at the root (r.sub.root) to the radius of the fan blade at the tip (r.sub.tip) is less than 0.28.

9. The gas turbine engine of claim 1, wherein: 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 comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

10. The gas turbine engine of claim 1, wherein the fan diameter is greater than 250 cm.

11. The gas turbine engine of claim 1, wherein the fan diameter is greater than 330 cm.

12. The gas turbine engine of claim 1, wherein: at cruise conditions, each cross-section through each fan blade experiences an inlet relative Mach number M1rel; and for the cross-sections through each fan blade for which a value of M1rel at cruise is less than 0.8 and the covered passage percentage is between 40 percent and 62 percent, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
|α.sub.2−α1|≥9 where the cruise conditions correspond to a forward Mach number of 0.8, a pressure of 23000 Pa; and a temperature of −55 degrees Celsius.

13. The gas turbine engine of claim 1, wherein: at cruise conditions, each cross-section through each fan blade experiences an inlet relative Mach number M1rel; and for the cross-sections through each fan blade for which a value of M1rel at cruise is less than 0.8, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
|α2−α1|≥33−(M1rel*30).

14. The gas turbine engine of claim 1, wherein: at cruise conditions, each cross-section through each fan blade experiences an inlet relative Mach number M1rel; and no cross-section through each fan blade has a value of M1rel at cruise that is greater than 1.15, wherein the forward speed of the gas turbine engine at the cruise conditions is in the range of from Mach 0.75 to Mach 0.85.

15. The gas turbine engine of claim 1, wherein a fan tip loading is defined as dH/U.sub.tip.sup.2, where dH is the enthalpy rise across the fan and U.sub.tip is the velocity of the fan tip, and the fan tip loading at a forward Mach number of 0.8 is greater than 0.3 JKg.sup.−1K.sup.−1/(ms.sup.−1).sup.2.

16. The gas turbine engine of claim 15, wherein the fan tip loading at the forward Mach number of 0.8 is in the range of from 0.3 to 0.4 JKg.sup.−1K.sup.−1/(ms.sup.−1).sup.2.

17. The gas turbine engine of claim 1, wherein a bypass ratio defined as the ratio of the mass flow rate of a bypass flow (B) that flows along a bypass duct to the mass flow rate of a core flow (A) at cruise conditions is greater than 12, where the cruise conditions correspond to a forward Mach number of 0.8, a pressure of 23000 Pa; and a temperature of −55 degrees Celsius.

18. The gas turbine engine of claim 17, wherein the bypass ratio is greater than 13.

19. The gas turbine engine of claim 1, wherein a ratio of a radius of the fan blade at the root (r.sub.root) to the radius of the fan blade at the tip (r.sub.tip) is less than 0.29.

Description

DESCRIPTION OF THE DRAWINGS

(1) Embodiments will now be described by way of example only, with reference to the Figures, in which:

(2) FIG. 1 is a sectional side view of a gas turbine engine in accordance with the present disclosure;

(3) FIG. 2 is a side view of a fan blade for use with examples of the present disclosure;

(4) FIG. 3 is a schematic view of two neighbouring blades taken through the cross-section A-A in FIG. 2;

(5) FIG. 4 is a schematic view of two neighbouring blades taken through the cross-section A-A in FIG. 2;

(6) FIG. 5 is a graph showing, by way of example only, covered passage against camber change before the start of the covered passage; and

(7) FIG. 6 is a graph showing, by way of example only, camber change before the start of the covered passage against relative Mach number at the leading edge.

DETAILED DESCRIPTION

(8) With reference to FIG. 1, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, a gearbox 14, an intermediate pressure compressor 15, a high-pressure compressor 16, combustion equipment 17, a high-pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines the intake 12. The nacelle 21 may be, or may comprise, a fan containment case 23. The nacelle 21 and the fan case 23 may be separate components.

(9) The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated and compressed by the fan 13 to produce two air flows: a first air flow A into the engine core and a second air flow B which passes through a bypass duct 22 to provide propulsive thrust. The first and second airflows A, B split at a generally annular splitter 40, for example at the leading edge of the generally annular splitter 40 at a generally circular stagnation line. In use (for example, at cruise conditions, which may be as defined elsewhere herein), the ratio of the mass flow rate of the bypass flow B to the core flow A may be as described and/or claimed herein, for example at least 10.

(10) The engine core includes the intermediate pressure compressor 15 (which may be referred to herein as a first compressor 15) which compresses the air flow directed into it before delivering that air to the high pressure compressor 16 (which may be referred to herein as a second compressor 16) where further compression takes place.

(11) The compressed air exhausted from the high-pressure compressor 16 is directed into the combustion equipment 17 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high pressure turbine 18 (which may be referred to as a second turbine 18) and the low pressure turbine 19 (which may be referred to as a first turbine 19) before being exhausted through the nozzle 20 to provide additional propulsive thrust. The intermediate pressure compressor 15 is driven by the low pressure turbine 19 by a first (or low pressure) shaft 32. The high pressure compressor 16 is driven by the low pressure turbine 18 by a second (or high pressure) shaft 34. The first shaft 32 also drives the fan 13 via the gearbox 14. The gearbox 14 is a reduction gearbox in that it gears down the rate of rotation of the fan 13 by comparison with the intermediate pressure compressor 15 and low pressure turbine 19. The gearbox 14 may be any suitable type of gearbox, such as an epicyclic planetary gearbox (having a static ring gear, rotating and orbiting planet gears supported by a planet carrier and a rotating sun gear) or a star gearbox. Additionally or alternatively the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

(12) The first and second compressors 15, 16, first and second turbines 19, 18, first and second shafts 32, 34, and the combustor 17 may all be said to be part of the engine core.

(13) 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 24 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.

(14) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction 300 (which is aligned with the rotational axis 11), a radial direction 400, and a circumferential direction 500 (shown perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions 300, 400, 500 are mutually perpendicular.

(15) FIG. 2 shows a fan blade 130 of the fan 13 in the gas turbine engine 10 in greater detail. The fan blade 130 extends from a root 132 to a tip 134 in a substantially radial spanwise direction 400. The root 132 may be defined by the radially innermost gas-washed points of the blade 130 and/or may be defined as an intersection between the fan blade 130 and a surface (for example a conical and/or cylindrical surface and/or an otherwise profiled endwall) from which the fan blades 13 extend. The fan blade 130 has a leading edge 136 and a trailing edge 138. The leading edge 136 may be defined as the line defined by the axially forwardmost points of the fan blade 130 from its root 132 to its tip 134. The fan blade 130 may (or may not) have a fixture portion (not shown) radially inboard of the root, which may be used to fix the fan blade 130 to the rest of the engine.

(16) The radius of the leading edge 136 of the fan blade 130 at its root 132 is designated in FIG. 2 as r.sub.root. The radius of the leading edge 136 of the fan blade 130 at its tip 134 is designated in FIG. 2 as r.sub.tip. The ratio of the radius of the leading edge 136 of the fan blade 130 at its root 132 to the radius of the leading edge 136 of the fan blade 130 at its tip 134 (r.sub.root/r.sub.tip) may be as described and/or claimed herein, for example less than 0.35 and/or less than 0.33 and/or less than 0.28.

(17) The span m of the blade 130 is defined as the difference in the radius of the leading edge 136 at the tip and the radius of the leading edge 136 at the root (r.sub.tip−r.sub.root).

(18) In use of the gas turbine engine 10, the fan 13 (with associated fan blades 130) rotates about the rotational axis 11. This rotation results in the tip 134 of the fan blade 130 moving with a velocity U.sub.tip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. Accordingly, a fan tip loading may be defined as dH/U.sub.tip.sup.2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan (or in the bypass stream) and U.sub.tip is the velocity of the fan tip (which may be defined as fan tip radius at leading edge multiplied by rotational speed). As noted elsewhere herein, the fan tip loading at cruise conditions may be greater than (or on the order of) 0.3, for example greater than (or on the order of) 0.31, for example greater than (or on the order of) 0.32, for example greater than (or on the order of) 0.33, for example greater than (or on the order of) 0.34, for example greater than (or on the order of) 0.35, for example greater than (or on the order of) 0.36, for example in the range of from 0.3 to 0.4 (all figures having units JKg.sup.−1K.sup.−1/(ms.sup.−1).sup.2).

(19) The specific thrust of the gas turbine engine 10 may be in the ranges described and/or claimed herein.

(20) A cross-sectional plane A-A or B-B through the blade 130 may be defined by an extrusion in the circumferential direction of a straight line formed between a point on the leading edge 136 that is at a given percentage X of the span m from the root 132 (i.e. at a radius of (r.sub.root+X/100*(r.sub.tip−r.sub.root))), and a point on the trailing edge that is at the same radial percentage X of a trailing edge radial extent t along the trailing edge 138 from the root 132 at the trailing edge 138. The circumferential direction of the extrusion may be taken at the leading edge position of the plane A-A, B-B. In other words, reference to a cross-section through the fan blade 130 may mean a section through the aerofoil in a plane defined by: a line that passes through the point on the leading edge that is at a given percentage of the span m along the leading edge from the leading edge root and points in the direction of the tangent to the circumferential direction at that point on the leading edge; and a point on the trailing edge that is at that same percentage along the trailing edge 138 from the trailing edge root.

(21) FIG. 3 is a schematic showing a cross-section A-A (indicated in FIG. 2) through two neighbouring fan blades 130. FIG. 4 is a similar schematic to that shown in FIG. 3, but taken through a cross-section B-B at a different spanwise position of the blades 130.

(22) The neighbouring fan blades 130 are both part of the fan 13. The neighbouring fan blades 130 may be substantially identical to each other, as in the example of FIGS. 3 and 4. The spacing between the blades 130 (for example between any two equivalent points on the fan blades 130, for example a given spanwise position on their leading edges 135 and/or trailing edges 138) is indicated by the letter S in FIGS. 3 and 4. This spacing S may be referred to as the pitch of the fan blades 130. Although indicated as a straight line in FIGS. 3 and 4, in the illustrated example the spacing S is actually the circumferential distance between the two neighbouring fan blades 130, and as such may be different depending on the spanwise position of the cross-section (in particular, the spacing S typically increases with increasing spanwise position (or increasing radius)).

(23) A camber line (142 in FIG. 3, 242 in FIG. 4) is defined for a given cross-section as the line formed by the points in that cross-section that are equidistant from a pressure surface 139 and a suction surface 137 of the blade 130 for that cross-section. The change in the angle of the camber line 142, 242 between any two points is simply the angle between the tangent to the camber line 142, 242 at each of those two points.

(24) A true chord for a given cross-section (C.sub.A in FIG. 3, C.sub.B in FIG. 4) is the distance along the camber line (which would typically be a curved line) between the leading edge 136 and the trailing edge 138 of the aerofoil 130 in that cross-section. Accordingly, the true chord (C.sub.A, C.sub.B) would typically be the length of a curved line. Note that this is different to what might conventionally be referred to as the chord length, which would be the length of a straight line drawn between the leading edge 136 and the trailing edge 138 of the aerofoil 130 in that cross-section (and is not shown in FIGS. 3 and 4).

(25) A covered passage is defined as the part of the blade (or part of the passage between the blades) for a given cross-section that is between a line J that passes through the point K on the suction surface 137 of the blade 130 and the leading edge 136 of the neighbouring blade 130 that is adjacent the suction surface 137. The point K is defined as the point K on the suction surface that is closest to the leading edge 136 of a neighbouring blade. The line J may pass entirely through the cross-section of the blade, so as to separate the cross-section into two parts: a covered passage part that is between the trailing edge 138 and the line J, and a non-covered passage part that is between the leading edge 136 and the line J. The line J may be described as being a straight line when viewed from a radial direction.

(26) A covered passage length P (P.sub.A in FIG. 3 and P.sub.B in FIG. 4) is then defined as the distance along the portion of the camber line C that is within the covered passage portion of the blade. Again, the covered passage length P would therefore typically be the length of a curved line.

(27) A covered passage percentage is then defined as the covered passage length (P) as a percentage of the true chord (C), that is ((P.sub.A/C.sub.A)*100) for the cross-section A-A and ((P.sub.B/C.sub.B)*100) for the cross-section B-B.

(28) Note that one or both of the true chord length C and local pitch S may change depending on the spanwise position of the cross-section.

(29) The angle of the camber line 142, 242 (that is, the tangent to the angle of the camber line 142, 242) for a given cross-section A-A, B-B changes between the leading edge 136 of the blade 130 and the point on the camber line 142, 242 that is at the start of the covered passage P. In this regard, the start of the covered passage P may be the axially forwardmost point of the covered passage P through which the camber line 142, 242 passes, that is the point at which the line J crosses the camber line 142, 242. The angle of the camber line may be measured relative to any other line in the plane of the cross-section, because it is change in angle of the camber line 142, 242 that is importance.

(30) In the example of FIGS. 3 and 4, the angle of the camber line 142, 242 is measured relative to a line that is parallel with the axial direction 300. The difference between the angle α.sub.1 of the camber line 142, 242 at the leading edge 136 and the angle α.sub.2 at the start of the covered passage P is simply given by |α.sub.2−α.sub.1|. Where the term change (or difference) in angle of the camber line between two points is used herein, this means the magnitude of the change (or difference) in the angle of the camber line between those two points.

(31) It will be appreciated that the length of the covered passage P and the change in angle (|α.sub.2−α.sub.1|) of the camber line 142, 242 between the leading edge 136 and the start of the covered passage P are different for at least some cross-sections taken through the blade 130. This is illustrated by way of example only by the difference between the cross-sections A-A and B-B shown in FIGS. 2 to 4.

(32) FIG. 5 is a graph showing three examples (lines D, E, F) of how the change in angle (|α.sub.2−α.sub.1|) of the camber line 142, 242 between the leading edge 136 and the start of the covered passage P may vary with the covered passage percentage for examples of fan blades 130 in accordance with the present disclosure. Note that the x-axis in FIG. 5 represents (100-covered passage percentage), rather than simply the covered passage percentage. Each of the lines D, E, F represents a different fan blade 130 that may be in accordance with aspects of the present disclosure. Each point on one of the lines D, E, F shown in FIG. 5 represents the change in angle (|α.sub.2−α.sub.1|) and the covered passage percentage for a particular cross-section through the respective blade 130. However, each line does not necessarily show all spanwise cross-sections for the fan blade 130. Purely by way of example, the lines D, E, F shown in FIG. 5 may represent spanwise cross-sections extending from 35% to 100% of the blade span, although of course this is in no way limitative, and the curves may represent cross-sections taken over other spanwise extents.

(33) The relationships D, E, F plotted in FIG. 5 are examples of fan blades 130 that satisfy the relationships between the covered passage percentage and the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) described and/or claimed herein. For example, in the FIG. 5 examples, for all cross-sections through each fan blade 130 for which the covered passage percentage is between 40% and 70%, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:

(34) $.Math.\mathrm{\alpha 2}-\mathrm{\alpha 1}.Math.\ge 6+\frac{\left(\mathrm{covered}\mathrm{passage}\mathrm{percentage}-35\right)}{9}$
and/or
for all cross-sections through each fan blade 130 for which the covered passage percentage is greater than 60%, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
|α2−α1|≥9.

(35) In FIG. 5, a point that is more towards the bottom right of the line than another point represents a cross-section that is closer to the tip 134 than that other point. A point on a line that has a higher covered passage percentage relative to another point (i.e. more towards the left of the curves) is generally closer to the tip 134, although this is not essential.

(36) FIG. 6 is a graph showing three examples (lines G, H, I) of how the change in angle (|α.sub.2−α.sub.1|) of the camber line 142, 242 between the leading edge 136 and the start of the covered passage P may vary with the inlet relative Mach number M1rel for different cross-sections through the blade 130. Each of the lines G, H, I represents a different fan blade 130 that may be in accordance with aspects of the present disclosure, which may be the same or different to the blades 130 represented by lines D, E, F in FIG. 5. Purely by way of example, the line G in FIG. 6 may represent the same blade 130 as the line D in FIG. 6. Each point on one of the lines G, H, I shown in FIG. 6 represents the change in angle (|α.sub.2−α.sub.1|) and the inlet relative Mach number for a particular cross-section through the blade 130. However, each line does not necessarily show all spanwise cross-sections for the fan blade 130. Purely by way of example, the lines G, H, I shown in FIG. 6 may represent spanwise cross-sections extending from 35% to 100% of the blade span, although of course this is in no way limitative, and the curves may represent cross-sections taken over other spanwise extents.

(37) The inlet relative Mach number may be calculated using the vector sum of the blade forward speed (which may be taken as the forward speed of an aircraft to which a gas turbine engine 10 is attached) and the linear blade speed at the radial position of the leading edge 136 of the cross-section due to the rotation of the fan blades 130, at cruise conditions (which may be as defined elsewhere herein). This is illustrated schematically in FIGS. 3 and 4, which show a velocity triangle (towards the bottom left of the Figures) in which the forward (axial) velocity of the blade 130 (or engine 10) is shown as “U”, the linear circumferential velocity at the respective radius r.sub.A, r.sub.B due to the rotation Ω of the fan 13 is indicated as rΩ, with the vector sum being shown as V1rel. From V1rel, the inlet relative Mach number can be calculated in the conventional manner, using the local speed of sound. The inlet relative Mach number generally increases with increasing radial (or spanwise) position of the cross-section.

(38) The relationships plotted in FIG. 6 are examples of fan blades 130 that satisfy the relationships between the inlet relative Mach number and the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) described and/or claimed herein. For example, in the FIG. 6 examples, for all cross-sections through each fan blade 130 for which a value of M1rel at cruise is less than 0.8, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
|α2−α1|≥9
and/or
for all cross-sections through each fan blade for which a value of M1rel at cruise is less than 0.75, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
|α2−α1|≥10
and/or
for all cross-sections through each fan blade for which a value of M1rel at cruise is less than 0.8, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
|α2−α1|≥33−(M1rel*30).

(39) In use, the gas turbine engine 10 may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine 10 may be mounted in order to provide propulsive thrust. Parameters such as pressure ratios referred to herein may be taken at such a cruise condition.

(40) 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.