Low speed fan up camber

Abstract

A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (α3−α1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

Claims

1. A fan blade for a gas turbine engine, the fan blade comprises: an airfoil portion having a leading edge extending from a root to a tip, the fan blade has a radial span (m) extending between the leading edge at the root at a 0% span position and to the leading edge at the tip at a 100% span position; a camber line defined by the midpoint between its pressure surface and its suction surface, with a true chord (C) being defined as the distance along the camber line between the leading edge and a trailing edge of the fan blade; and a covered passage defined as the 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 the leading edge of a neighbouring fan blade and the leading edge of that neighbouring fan blade; wherein: at the cross section through the fan blade at the point 80% of the blade span of the airfoil portion the change in angle (α3−α1) of the camber line between the leading edge and the trailing edge satisfies:
|α3−α1|>20° and at a cross section through the fan blade at a point 80% of the blade span of the airfoil portion the change in angle (α2−α1) of the camber line between the leading edge and the point on the camber line that corresponds to the start of the covered passage satisfies:
|α2−α1|>7° and at a cross section through the fan blade at a height greater than 80% of the blade span of the airfoil portion the angle (α3−α1) of the camber line between the leading edge and the trailing edge satisfies:
−0.4°*cross section percentage height+64°≥|α3−α1|≥−0.3°*cross section percentage height+44°.

2. The fan blade of claim 1, wherein at the cross section through the fan blade at the point 80% of the blade span of the airfoil portion the change in angle (α3−α1) of the camber line between the leading edge and the trailing edge satisfies: 32°>|α3−α1|>20°.

3. The fan blade of claim 1, wherein at the cross section through the fan blade at the point 80% of the blade span of the airfoil portion the change in angle (α3−α1) of the camber line between the leading edge and the trailing edge satisfies: 29°>|α3−α1|>20°.

4. The fan blade of claim 1, wherein at the cross section through the fan blade at the point 80% of the blade span of the airfoil portion the change in angle (α2−α1) of the camber line between the leading edge and the point on the camber line that corresponds to the start of the covered passage satisfies: 15°>|α2−α1|>7°.

5. The fan bade of claim 1, wherein at the cross section through the fan blade at the point 80% of the blade span of the airfoil portion the change in angle (α2−α1) of the camber line between the leading edge and the point on the camber line that corresponds to the start of the covered passage satisfies: 10°>|α2−α1|>7°.

6. 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 according to claim 1; 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.

7. The gas turbine engine of claim 6, 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 0.33.

8. The gas turbine engine of claim 6, 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.

9. The gas turbine engine of claim 6, wherein the fan diameter is greater than 250 cm.

10. The gas turbine engine of claim 6, wherein a bypass ratio is 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 the core flow (A) at cruise conditions, and the bypass ratio is greater than 10.

11. The gas turbine engine of claim 6, wherein the specific thrust at cruise conditions is less than 100 NKg.sup.−1s.sup.−1.

Description

BRIEF 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;

(3) FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

(4) FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

(5) FIG. 4 is a side view of a fan blade for use with examples of the present disclosure;

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

(7) FIG. 6 is a graph showing, by way of example only, camber change between the leading edge and the trailing edge (α3−α1).

DETAILED DESCRIPTION OF THE DISCLOSURE

(8) 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.

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

(10) In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

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

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

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

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

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

(16) 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.

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

(18) 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 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

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

(20) FIG. 4 shows a fan blade 130 of the fan 23 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.

(21) The radius of the leading edge 136 of the fan blade 130 at its root 132 is designated in FIG. 4 as r.sub.root. The radius of the leading edge 136 of the fan blade 130 at its tip 134 is designated in FIG. 4 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.

(22) 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).

(23) In use of the gas turbine engine 10, the fan 23 (with associated fan blades 130) rotates about the rotational axis 9. 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 130 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). The fan tip loading at cruise conditions may be in the ranges described and/or claimed elsewhere herein.

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

(25) 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.

(26) FIG. 5 is a schematic showing a cross-section A-A (indicated in FIG. 4) through two neighbouring fan blades 130.

(27) The neighbouring fan blades 130 are both part of the fan 23. The neighbouring fan blades 130 may be substantially identical to each other, as in the example of FIG. 5. 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. This spacing S may be referred to as the pitch of the fan blades 130. Although indicated as a straight line, 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)).

(28) A camber line 142 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, between any two points is simply the angle between the tangent to the camber line 142, at each of those two points.

(29) A true chord for a given cross-section, CA, 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 CA 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 FIG. 5).

(30) 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.

(31) 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.

(32) The angle of the camber line 142, (that is, the tangent to the angle of the camber line 142,) for a given cross-section A-A, changes between the leading edge 136 of the blade 130 and the trailing edge 138 of the blade 130, and also between the leading edge 136 of the blade and the point on the camber line 142 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 passes, that is the point at which the line J crosses the camber line 142. 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 that is importance.

(33) In the example of FIG. 5, the angle of the camber line 142 is measured relative to a line that is parallel with the axial direction 300. The difference between the angle α1 of the camber line 142 at the leading edge 136 and the angle α.sub.3 of the trailing edge 138 is simply given |α.sub.3−α.sub.1|. The difference between the angle α.sub.1 of the camber line 142 at the leading edge 136 and the angle α2 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.

(34) It will be appreciated that the change in angle (|α.sub.3−α.sub.1|) of the camber line 142 between the leading edge 136 and the trailing edge 138, and that the change in angle (|α.sub.2−α.sub.1|) of the camber line 142 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.

(35) FIG. 6 is a graph showing three examples (lines D, E, F) of how the change in angle (|α.sub.3−α.sub.1|) of the camber line 142 between the leading edge 136 and the trailing edge 138 may vary with a particular cross section of the blade span through the respective blade 130. 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. 6 represents the change in angle (|α.sub.3−α.sub.1|) for a particular cross-section through the respective blade 130.

(36) The relationships D, E, F plotted in FIG. 6 are examples of fan blades 130 that satisfy the relationships between the cross section blade height and the change in angle of the camber line between the leading edge (α.sub.1) and the trailing edge (α.sub.3), For example at a cross section through the fan blade at a height greater than 80% of the blade span of the airfoil portion (130) the angle (α.sub.3−α.sub.1) of the camber line (142) between the leading edge (136) and the trailing edge (138) satisfies:
|α.sub.3−α.sub.1|≥−0.5*cross section percentage height+60

(37) 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.

(38) 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.