Aerofoil assembly and method

Abstract

An aerofoil assembly includes a platform and a plurality of aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. Each aerofoil has a leading edge proximal to the first edge and a trailing edge distal to the first edge. A pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface. A mid-pitch location is defined midway along the pitch spacing. The platform defines one or more recesses disposed between the leading edges of the plurality of aerofoils and the first edge. Each of the one or more recesses is disposed proximal to the mid-pitch location between adjacent aerofoils.

Claims

1. An aerofoil assembly comprising: a platform having a first edge, a second edge and a platform surface disposed between the first edge and the second edge; and a plurality of aerofoils extending radially outward from the platform and disposed between the first edge and the second edge, each aerofoil having a leading edge proximal to the first edge and a trailing edge distal to the first edge, wherein a pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface, and wherein a mid-pitch location is defined midway along the pitch spacing; wherein the platform defines one or more recesses entirely disposed between the leading edges of the plurality of aerofoils and the first edge, wherein each of the one or more recesses is disposed between adjacent aerofoils, wherein each of the one or more recesses comprises a first lobe and a second lobe adjoining the first lobe, and wherein the first lobe is immediately upstream of the second lobe.

2. The aerofoil assembly of claim 1, wherein each of the one or more recesses has a maximum depth of between 0.1% to 6% of a maximum height of each aerofoil relative to the platform surface.

3. The aerofoil assembly of claim 1, further comprising a plurality of blade segments disposed adjacent to each other, each blade segment comprising a pair of adjacent aerofoils from the plurality of aerofoils and a platform portion that forms part of the platform.

4. The aerofoil assembly of claim 3, wherein the platform portion of each blade segment defines a recess from the one or more recesses between the pair of adjacent aerofoils.

5. The aerofoil assembly of claim 1, wherein each of the one or more recesses is disposed between the mid-pitch location and the first edge.

6. The aerofoil assembly of claim 1, wherein a minimum distance between the mid-pitch location and each of the one or more recesses is between 0% to 70% of a distance between the first edge and the leading edge of each aerofoil.

7. The aerofoil assembly of claim 1, wherein a minimum distance between the first edge and each of the one or more recesses is between 0% to 70% of a distance between the first edge and the leading edge of each aerofoil.

8. The aerofoil assembly of claim 1, wherein a minimum distance between each of the one or more recesses and the leading edge of each of the adjacent aerofoils is between 10% to 60% of the pitch spacing between the leading edges of the adjacent aerofoils.

9. The aerofoil assembly of claim 1, wherein the aerofoil assembly is a turbine blade assembly.

10. A gas turbine engine comprising the aerofoil assembly of claim 1.

11. A method of reducing losses in an aerofoil assembly, comprising: providing a platform having a first edge, a second edge and a platform surface disposed between the first edge and the second edge; providing a plurality of aerofoils extending radially outward from the platform and disposed between the first edge and the second edge, each aerofoil having a leading edge proximal to the first edge and a trailing edge distal to the first edge, wherein a pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface, and wherein a mid-pitch location is defined midway along the pitch spacing; and forming one or more recesses entirely disposed between the leading edges of the plurality of aerofoils and the first edge, wherein each of the one or more recesses is disposed between adjacent aerofoils, wherein each of the one or more recesses comprises a first lobe and a second lobe adjoining the first lobe, and wherein the first lobe is immediately upstream of the second lobe.

12. The method of claim 11, wherein each of the one or more recesses has a maximum depth of between 0.1% to 6% of a maximum height of each aerofoil relative to the platform surface.

13. The method of claim 11, wherein each of the one or more recesses is disposed between the mid-pitch location and the first edge.

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 schematic side view of an aerofoil assembly of the gas turbine engine;

(6) FIG. 4A is a detailed view of a region R of FIG. 4;

(7) FIG. 5 is a partial schematic perspective view of the aerofoil assembly of FIG. 4;

(8) FIG. 6 is a partial schematic plan view of the aerofoil assembly of FIG. 5; and

(9) FIG. 7 is a flowchart of a method of reducing losses in an aerofoil assembly.

DETAILED DESCRIPTION OF THE DISCLOSURE

(10) 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.

(11) 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.

(12) 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 core exhaust 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.

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

(14) 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.

(15) 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.

(16) 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.

(17) 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.

(18) 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.

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

(20) 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 exhaust 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. 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 10 may not comprise a gearbox 30.

(21) 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.

(22) FIG. 4 illustrates an aerofoil assembly 200 in accordance with an embodiment of the present disclosure. The gas turbine engine 10 (shown in FIG. 1) includes the aerofoil assembly 200. In an embodiment, the aerofoil assembly 200 is a turbine blade assembly of the gas turbine engine 10. The aerofoil assembly 200 may be part of at least one of the high pressure turbine 17 and the low pressure turbine 19. FIG. 4A is a detailed view of a region R of FIG. 4.

(23) Referring to FIGS. 1, 4 and 4A, the aerofoil assembly 200 includes a row of stator vanes 202 (only one shown in FIG. 4) and a row of blades 204 (only one shown in FIG. 4) located downstream of the row of stator vanes 202. The blades 204 may be mounted on a rotor disc (not shown). The stator vanes 202 and the blades 204 may form a single stage of the aerofoil assembly 200. In some embodiments, the aerofoil assembly 200 may include multiple stages.

(24) The stator vanes 202 extend from a static wall 206. The blades 204 extend from a rotating wall 208. A wheel space 210 is defined between the static wall 206 and the rotating wall 208. In operation, cooling air or purge air 212 is introduced into the wheel space 210. Purge air 212 may cool components and spaces within the wheel space 210. Purge air 212 may be tapped from a compressor, for example, the low pressure compressor 14 and/or the high pressure compressor 15.

(25) Further, a hot gas region 213 is defined between the stator vanes 202 and the blades 204. The hot gas region 213 receives hot gas 214. Purge air 212 may restrict incursion of hot gas 214 into the wheel space 210. Specifically, a flow of purge air 212 may be used to purge the wheel space 210 into the hot gas region 213 such that purge air 212 restricts hot gas 214 from flowing into the wheel space 210. Purge air 212 may therefore provide a rim seal flow in the aerofoil assembly 200.

(26) An ejection of purge air 212 out of the wheel space 210 and interaction with hot gas 214 may result in a reduction of efficiency of the aerofoil assembly 200. The reduction of efficiency may be due to various types of losses, for example, mixing losses, penetration losses, secondary vortices, etc.

(27) The aerofoil assembly 200 further includes a platform 302 having a first edge 304, a second edge 306, and a platform surface 308 disposed between the first edge 304 and the second edge 306. The first edge 304 faces the stator vanes 202. The platform surface 308 is the radially outward surface of the platform 302. Each blade 204 includes an aerofoil 310 extending radially outward from the platform 302 and disposed between the first edge 304 and the second edge 306. The aerofoil assembly 200 includes a plurality of such aerofoils 310 arranged in an array. Each aerofoil 310 includes a leading edge 312 proximal to the first edge 304 of the platform 302 and a trailing edge 314 distal to the first edge 304. Each aerofoil 310 defines a maximum height “HM” relative to the platform surface 308. The maximum height “HM” is the maximum radial height between the platform surface 308 and a tip 315 of the aerofoil 310. The maximum height “HM” may be defined between the platform surface 308 and the tip 315 of the aerofoil 310 adjacent to the trailing edge 314.

(28) The platform 302 defines one or more recesses 316 (only one shown in FIG. 4A) disposed between the leading edges 312 of the plurality of aerofoils 310 and the first edge 304 of the platform 302. In some embodiments, each of the one or more recesses 316 is formed by removing material from the platform surface 308. In some other embodiments, each of the one or more recesses 316 is formed by casting. In other words, each of the one or more recesses 316 is a cast-in feature. Each of the one or more recesses 316 has a maximum depth “DM” of between about 0.1% to about 6% of the maximum height “HM” of each aerofoil 310 relative to the platform surface 308. A depth of the recess 316 may increase from a boundary 317 (shown in FIG. 5) of the recess 316 to the maximum depth “DM”. Further, the depth of the recess 316 may be defined with respect to a baseline “BL” (shown by a dashed line) of the platform surface 308 without any recess. The baseline “BL” is a normal profile of the platform surface 308 without any recess or removal of material.

(29) The aerofoil 310 further defines a chord length “CL” between the leading edge 312 and the trailing edge 314. The chord length “CL” is a length of a straight line connecting the leading and trailing edges 312, 314. In some embodiments, a minimum distance “D2” between the recess 316 and the leading edge 312 is 0 percent (%) to 5% of the chord length “CL”. The minimum distance “D2” may be a minimum distance between the boundary 317 of the recess 316 and the leading edge 312.

(30) The recesses 316 may reduce losses due to purge air 212 and improve the efficiency of the aerofoil assembly 200, and hence the gas turbine engine 10.

(31) FIG. 5 illustrates a partial perspective view of the aerofoil assembly 200 in accordance with an embodiment of the present disclosure. FIG. 6 illustrates a partial plan view of the aerofoil assembly 200 in accordance with an embodiment of the present disclosure. As shown in FIG. 5, the aerofoil assembly 200 includes a plurality of blade segments 402 disposed adjacent to each other. Each blade segment 402 includes a pair of adjacent aerofoils 310 from the plurality of aerofoils 310 and a platform portion 404 that forms part of the platform 302. The pair of adjacent aerofoils 310 extends outwardly from the platform portion 404 of the platform 302. Each blade segment 402 further defines a recess 316 from the one or more recesses 316 between the pair of adjacent aerofoils 310. In other embodiments, each blade segment 402 may include more than two adjacent aerofoils 310. Further, in some embodiments, each platform portion 404 may define two or more recesses 316 between adjacent aerofoils 310.

(32) The platform portion 404 of each blade segment 402 includes a pair of longitudinal edges 406. The longitudinal edge 406 of the platform portion 404 of each blade segment 402 is aligned with the longitudinal edge 406 of the platform portion 404 of the adjacent blade segment 402. In the illustrated embodiment of FIG. 5, one blade segment 402 is shown and an adjacent blade segment 402 is partially shown without the aerofoils 310. However, multiple such blade segments 402 may be aligned to form a circumferential array of the aerofoils 310. Each blade segment 402 may further include a blade root (not shown). Each blade root may extend radially inward from the corresponding platform portion 404.

(33) The platform portions 404 together form the platform 302. The first edge 304 of the platform 302 may be formed together by first edge segments (not shown) of the blade segments 402. Similarly, the second edge 306 may be formed together by second edge segments (not shown) of the blade segments 402. The longitudinal edge 406 of one blade segment 402 may be joined to the longitudinal edge 406 of the adjacent blade segment 402 by various methods, for example, but not limited to, welding, brazing, mechanical fasteners, mechanical joints, or combinations thereof.

(34) Referring to FIGS. 4, 4A, 5 and 6, a pitch spacing “PD” is defined between the leading edges 312 of the adjacent aerofoils 310 along the platform surface 308. A mid-pitch location “PL” is defined midway along the pitch spacing “PD”. The mid-pitch location “PL” may be a point defined midway (i.e., mid-point) on a straight line connecting the leading edges 312 of the adjacent aerofoils 310. A length of the straight line connecting the leading edges 312 is the pitch spacing “PD”. The recess 316 is disposed proximal to the mid-pitch location “PL” between the adjacent aerofoils 310. A distance “AL” is defined between the first edge 304 and the leading edge 312 of each aerofoil 310. The distance “AL” is an axial distance between the first edge 304 of the platform 302 and an aerofoil leading edge plane. In some embodiments, a minimum distance “D1” between the first edge 304 and each of the one or more recesses 316 is between about 0% to about 70% of the distance “AL” between the first edge 304 and the leading edge 312 of each aerofoil 310. The minimum distance “D1” may be a minimum axial distance between the boundary 317 of the recess 316 and the first edge 304. In some embodiments, a minimum distance “D3” between the mid-pitch location “PL” and each of the one or more recesses 316 is between about 0% to about 70% of the distance “AL” between the first edge 304 and the leading edge 312 of each aerofoil 310. The minimum distance “D3” is a minimum axial distance between the boundary 317 of the recess 316 and the mid-pitch location “PL”. The depth of the recess 316 increases from the boundary 317. A variation of the depth of the recess 316 from the boundary 317 may be uniform or non-uniform along a length of the boundary 317. As illustrated in FIG. 4A, the variation of the depth of the recess 316 may be non-linear and varies along the length of the boundary 317.

(35) In an embodiment, each of the one or more recesses 316 extends from the first edge 304 of the platform 302 to the mid-pitch location “PL”. In such a case, the first edge 304 includes a portion of the recess 316. Further, each of the minimum distances “D1”, “D2” and “D3” is zero.

(36) In some embodiments, a minimum distance “D4”, “D5” between each of the one or more recesses 316 and the leading edge 312 of each of the adjacent aerofoils 310 is between about 10% to about 60% of the pitch spacing “PD” between the leading edges 312 of the adjacent aerofoils 310. The minimum distance “D4” may be defined between the recess 316 and one of the adjacent aerofoils 310 on one side of the recess 316. The minimum distance “D5” may be defined between the recess 316 and the other of the adjacent aerofoils 310 on another side of the recess 316. In some embodiments, the minimum distance “D4” is equal to the minimum distance “D5”. In some other embodiments, the minimum distance “D4” is different from the minimum distance “D5”. The minimum distance “D4” may be a minimum circumferentially projected distance from the boundary 317 of the recess 316 to the leading 312 of one of the adjacent aerofoils 310. Similarly, the minimum distance “D5” may be a minimum circumferentially projected distance from the boundary 317 of the recess 316 to the leading edge 312 of the other adjacent aerofoil 310.

(37) Further, a mid-pitch region “PR” is at least partly defined between the adjacent aerofoils 310. The mid-pitch region “PR” extends from the first edge 304 to the second edge 306 of the platform 302. Further, the mid-pitch region “PR” extends through the mid-pitch location “PL”. The mid-pitch region “PR” may be a line that is a locus of mid-points between the adjacent aerofoils 310 on the platform surface 308. Specifically, the mid-pitch region “PR” may be a line that joins all mid-points between a pressure surface 318 of one aerofoil 310 and a suction surface 320 of the adjacent aerofoil 310 along the platform surface 308. The line may be straight, curved or a combination of both. Further, the mid-pitch region “PR” intersects the mid-pitch location “PL”.

(38) The recess 316 includes a first lobe 322 and a second lobe 324 disposed adjoining the first lobe 322. The boundary 317 of the recess 316 may therefore define two curved regions that are joined by a pair of rounded regions. In some embodiments, an area of the first lobe 322 may be substantially equal to an area of the second lobe 324. In alternative embodiments, the area of the first lobe 322 may be different from the area of the second lobe 324. Each of the first lobe 322 and the second lobe 324 may have any suitable shape, for example, but not limited to, circular, elliptical, oval or any curved shape.

(39) The shape of the recess 316 is exemplary in nature. In alternative embodiments, the shape of the recess 316 can be, for example, but not limited to, polygonal, oval, circular, elliptical, or any other suitable shape.

(40) Each aerofoil 310 may be made of any suitable material such as a metal, a metal alloy, a ceramic, a composite, or combinations thereof. Each aerofoil 310 may include one or more channels for allowing flow of a cooling fluid.

(41) The platform 302 may be made of any suitable material such as a metal, a metal alloy, a ceramic, a composite, or combinations thereof. The platform 302 may include one or more channels for allowing flow of a cooling fluid.

(42) The recesses 316 described above may provide an easier escape path for purge air 212 to leak into the hot gas region 213 with minimal interaction with the respective leading edges 312. Maintaining a gap or distance between the flow of purge air 212 and the respective leading edges 312 may mitigate the formation of pressure side horseshoe vortices. Therefore, the recesses 316, may reduce secondary losses in a passage between respective adjacent aerofoils 310. Consequently, the recesses 316 may improve the efficiency of the aerofoil assembly 200. In some cases, the recesses 316 may improve a stage efficiency of a turbine by at least 0.1%, at least 0.2%, at least 0.5%, at least 1%, at least 2%, or at least 3%. The recesses 316 may also result in weight reduction of the aerofoil assembly 200.

(43) FIG. 7 illustrates a method 600 of reducing losses in an aerofoil assembly. The method 600 will be described with reference to the aerofoil assembly 200 described above with reference to FIGS. 4, 4A, 5 and 6.

(44) At step 602, the method 600 includes providing the platform 302 having the first edge 304, the second edge 306, and the platform surface 308 disposed between the first edge 304 and the second edge 306.

(45) At step 604, the method 600 includes providing the plurality of aerofoils 310 extending radially outward from the platform 302, and disposed between the first edge 304 and the second edge 306. Each aerofoil 310 has the leading edge 312 proximal to the first edge 304 and the trailing edge 314 distal to the first edge 304. The pitch spacing “PD” is defined between the leading edges 312 of adjacent aerofoils 310 along the platform surface 308. The mid-pitch location “PL” is defined midway along the pitch spacing “PD”.

(46) At step 606, the method 600 further includes forming the one or more recesses 316 disposed between the leading edges 312 of the plurality of aerofoils 310 and the first edge 304. Each of the one or more recesses 316 is disposed proximal to the mid-pitch location “PL” between adjacent aerofoils 310.

(47) In some embodiments, each of the one or more recesses 316 is formed by removing material from the platform surface 308. The material can be removed from the platform surface 308 by various material removal processes, for example, but not limited to, milling, drilling, grinding, electrical discharge machining, ultrasonic machining, abrasive jet machining, electron beam machining, or combinations thereof.

(48) In some other embodiments, each of the one or more recesses 316 is formed by casting. In other words, each of the one or more recesses 316 is a cast-in feature.

(49) In some embodiments, each of the one or more recesses 316 has the maximum depth “DM” of between about 0.1% to about 6% of the maximum height “HM” of each aerofoil 310 relative to the platform surface 308. In some embodiments, the mid-pitch region “PR” at least partly defined between adjacent aerofoils 310 extends from the first edge 304 to the second edge 306 through the mid-pitch location “PL”. Each of the one or more recesses 316 is disposed on the mid-pitch region “PR”.

(50) In some embodiments, each of the one or more recesses 316 includes the first lobe 322 and the second lobe 324 adjoining the first lobe 322.

(51) In some embodiments, the minimum distance “D3” between the mid-pitch location “PL” and the recess 316 is between about 0% to about 70% of the distance “AL” between the first edge 304 and the leading edge 312 of each aerofoil 310.

(52) In some embodiments, the minimum distance “D1” between the first edge 304 and the recess 316 is between about 0% to about 70% of the distance “AL” between the first edge 304 and the leading edge 312 of each aerofoil 310.

(53) In some embodiments, the minimum distance “D4”, “D5” between the recess 316 and the leading edge 312 of each of the adjacent aerofoils 310 is between about 10% to about 60% of the pitch spacing “PD” between the leading edges 312 of the adjacent aerofoils 310.

(54) In some embodiments, the recess 316 extends from the first edge 304 of the platform 302 to the mid-pitch location “PL”.

(55) The method 600 may improve the stage efficiency of a turbine by mitigating the formation of pressure side horseshoe vortices. The method 600 may also result in weight reduction of the turbine.

(56) 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 and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.