VANE JOINT

Abstract

A vane for a gas turbine engine, the vane including a platform with an airfoil extending radially from the upper surface of the platform. The platform includes a joint portion which includes a circumferentially extending flange and a recessed surface both formed on either the upper or lower surface of the platform. The flange and the recessed surface extend from opposing circumferential edges of the joint portion and each include a substantially radially-extending through hole.

Claims

1. A vane for a gas turbine engine, the vane comprising a platform having an upper surface and a lower surface, the platform comprising a mounting portion on which an airfoil is mounted extending radially from the upper surface and a joint portion at a first axial end, one of the upper or lower surfaces of the joint portion comprising a circumferentially-extending flange and a recessed surface, wherein the flange and recessed surface extend from opposing circumferential edges of the joint portion and each comprise a substantially radially-extending through hole.

2. The vane according to claim 1 wherein the flange and recessed surface are provided on the upper surface of the joint portion.

3. The vane according to claim 2 wherein the flange comprises a lower surface which is substantially aligned with the recessed surface in the circumferential direction.

4. The vane according to claim 1 wherein the recessed surface extends from a first axial end face of the platform.

5. A vane/stator assembly comprising first vane and circumferentially-adjacent second vane, each according to claim 1 wherein the through hole in the flange on the first vane is aligned with the through hole in the recessed surface of the second vane, the assembly further comprising a fixing element extending through the aligned through holes.

6. The assembly according to claim 5 wherein the flange of the first vane overlies the recessed surface of the second vane with the axes of the through holes aligned.

7. The assembly according to claim 6 wherein the flange of the first vane has a lower surface which abuts the recessed surface of the second vane, the lower surface of the flange of the first vane being substantially aligned in the circumferential direction with the recessed surface of the second vane.

8. The assembly according to claim 5 further comprising a third vane according to the platform having the upper surface and the lower surface, the platform comprising the mounting portion on which the airfoil is mounted extending radially from the upper surface and the joint portion at the first axial end, one of the upper or lower surfaces of the joint portion comprising the circumferentially-extending flange and the recessed surface, wherein the flange and recessed surface extend from opposing circumferential edges of the joint portion and each comprise a substantially radially-extending through hole, wherein the through hole in the recessed surface on the first vane is aligned with the through hole in the flange of the third vane, the assembly further comprising a second fixing element extending through the aligned through holes.

9. The assembly according to claim 5 wherein the assembly is a fan outlet guide vane assembly.

10. A gas turbine engine having a vane/stator assembly according to claim 5.

11. The gas turbine engine according to claim 10 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.

12. The gas turbine engine according to claim 11, 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.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0087] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0088] FIG. 1 is a sectional side view of a gas turbine engine;

[0089] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0090] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0091] FIG. 4 shows a perspective view of an exemplary embodiment of an OGV assembly 80;

[0092] FIG. 5 is an axial view of an OGV assembly illustrating the balancing of forces in the fixing elements; and

[0093] FIG. 6 shows an alternative view of FIG. 4 with the first OGV omitted.

DETAILED DESCRIPTION

[0094] 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.

[0095] 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.

[0096] 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.

[0097] 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.

[0098] 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.

[0099] 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.

[0100] 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.

[0101] 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.

[0102] 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.

[0103] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0104] 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.

[0105] 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.

[0106] FIG. 1 and FIG. 2 further show the location in the bypass duct 22 of an outlet guide vane (OGV) assembly 80 comprising a plurality of circumferentially-spaced outlet guide vanes 50.

[0107] With reference to FIG. 4, there is an exemplary embodiment of a first OGV 50 for a gas turbine engine 10 which is circumferentially adjacent to a second OGV 60 and a third OGV 70. The first OGV 50 comprises a platform 52 having an upper surface 53 and a lower surface 54 with an airfoil 55 extending radially from the upper surface 53 of a mounting portion 52d of the platform 52. The upper surface 53 is the radially outer surface and the lower surface 54 is the radially inner surface (i.e. closest to the engine axis).

[0108] The platform 52 further comprises a joint portion 51 at its upstream axial end i.e. at the axial end closest to the fan.

[0109] The upper surface 51a of the joint portion 51 comprises a circumferentially-extending flange 56 and a recessed surface 57. The flange 56 and recessed surface 57 extend from opposing circumferential edges of the joint portion 51. The flange 56 comprises a lower surface 56a and a substantially radially-extending through hole 58 (extending from the upper surface 53 of the platform 52 to the lower surface 56a of the flange 56). The through hole 58 is a counter-bored through hole.

[0110] The recessed surface also comprises a through hole 59 (visible in FIG. 6 for identical OGV 60) extending from the recessed surface 57 to the lower surface 54 of the platform.

[0111] The second and third OGVs 60, 70 have platforms 52′, 52″ that are identical to the platform 52 of the first OGV 50. Accordingly, the second OGV 60 has a second joint portion 51′ with an upper surface 53′ from which a second flange 56′ and a second recessed surface 57′ extend and the third OGV 70 has a third joint portion 51″ with an upper surface 53″ from which a third flange 56″ and a third recessed surface (not visible) extend.

[0112] In the assembly 80, the flange 56 of the first OGV 50 overlies the recessed surface 57′ of the joint portion 51′ of the second OGV 60 i.e. the lower surface of the flange 56a of the first OGV 50 abuts the recessed surface 57′ of the second OGV 60. To facilitate this abutment, the lower surface 56a of the flange of the first OGV 50 is aligned in the circumferential direction with the recessed surface 57′ of the second OGV 60.

[0113] The through hole 58 of the flange 56 of the first OGV 50 is axially and circumferentially aligned with the axis of the through hole 59 (visible in FIG. 6) of the recessed surface 57′ of the second OGV 60. A bolt 82 is affixed through the aligned through holes to secure the platform 52 of the first OGV 50 to the platform 52′ of the second OGV 60 together with their circumferential edges abutting one another. The bolt 82 is flush with the upper surface of the flange 56.

[0114] The bolt 82 further extends into the front structure 81 of the gas turbine engine.

[0115] Likewise, the flange 56″ of the third OGV 70 overlies the recessed surface 57 of the joint portion 51 of the first OGV 50 with the axis of the through hole 58′ of the third OGV 70 aligned in the radial direction with the axis of the through hole (not visible) of the first OGV 50. A second bolt 82′ passes through the aligned through holes (flush with the upper surface of the flange 56″ of the third vane 70) to secure the platforms 51, 51″ of the first and third OGV 50, 70 together as well as to the front structure 81 of the gas turbine engine 10. The circumferential edges of the first and third OGV 50, 70 are in abutment.

[0116] Although not shown, the OGV assembly 80 comprises further OGVs similar to the first, second third OGVs 50, 60, 70 so that the assembly 80 completely circumscribes the front structure 81.

[0117] This arrangement ensures that the forces on the fixing bolts 82, 82′ are reduced as substantially equal and opposite forces arising from adjacent airfoils 55, 55′, 55″ will be reacted at a single fixing bolt 82, 82′ between adjacent OVGs, 50, 60, 70 as shown in FIG. 5. The tangential lean of the aerofoils 55, 55′, 55″ ensures that the direction of loading is maintained as load reversal would double the loads for the given joint configuration.

[0118] Each of the flanges 56, 56′, 56″ has a convex curved circumferential end face 56b, 56b′. 56b″.

[0119] Each of the recessed surfaces 57, 57′, 57″ (57″ not visible) is defined by a respective first radially-extending wall 57a, 57a′, 57a″ which is most clearly seen in FIG. 6 (showing 57a′) where the first OGV 50 has been omitted. Each first radially-extending end wall 57a, 57a′, 57a″ joins the circumferential edge of the respective platform 52, 52′, 52″ with the respective upstream axial end face of the platform 52, 52′, 52″. Each of the radially-extending walls 57a, 57a′, 57a″ has a concave curved profile in a plane perpendicular to the respective recessed surface 57, 57′, 57″ (57″ not visible). This profile is complementary to the convex profile of the circumferential end face 56b, 56b′, 56b″ of the flange 56, 56′, 56″.

[0120] Accordingly, when the lower surface 56a of the flange 56 of the first OGV 50 overlies the recessed surface 57′ of the second OGV 60, the curved circumferential end face 56b of the flange 56 of the first OGV 50 can abut against curved first radially-extending wall 57a′ of the second OGV 60.

[0121] Each platform 52, 52′, 52″ has a respective second radially-extending wall 52a, 52a′, 52a″ extending radially inwards from the respective lower surface 56a, 56a′, 56a″ (56a″ not visible) of the flange 56, 56′, 56″ to the lower surface 54, 54′, 54″ of the platform 52, 52′, 52″. The second radially-extending wall 52a has a concave curved profile perpendicular to the plane of the flange 56.

[0122] Each platform 52, 52′, 52″ also has a respective third radially-extending wall 52b, 52b′, 52b″ (52b″ not visible) extending radially inwards from the respective recessed surface 57, 57′, 57″ (57″ not visible) to the respective lower surface 54, 54′, 54″ of the platform 52, 52′, 52″. The third radially-extending wall 52b, 52b′, 52b″ (52b″ not visible) has a convex curved profile perpendicular to the plane of the recessed surface 57, 57′, 57″ (57″ not visible). This profile is complementary to the convex curved profile of the radially-extending first wall 57, 57′, 57″ (57″ not visible).

[0123] Each upper surface 51a, 51a′, 51a″ of the joint portion 51, 51′, 51″ is radially recessed from the respective upper surface 53, 53′, 53″ of the platform 52, 52′, 52″. This is to allow a circumferentially-extending panel (not shown) to overlie the joint portions 51, 51′, 51″ and to be flush with the upper surfaces 53, 53′, 53″ of the platforms 52, 52′, 52″.

[0124] 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.