An Aerofoil Structure and a Method of Manufacturing an Aerofoil Structure for a Gas Turbine Engine

Abstract

The present disclosure relates to a method of manufacturing an aerofoil structure for a gas turbine engine, wherein the aerofoil structure includes a root configured to be received in a rotor disc of the gas turbine engine, the method including: providing a pre-moulded polymer insert; adding the polymer insert into a mould for forming the aerofoil structure; adding a composite constituent into the mould; and heating the composite constituent in the mould to bond the polymer insert to the composite constituent, the polymer insert being provided at a shoulder of the aerofoil structure root. The composite constituent is pre-impregnated with a resin and heating the composite constituent in the mould thermosets the resin. Heating the composite constituent in the mould to bond the polymer insert to the composite constituent forms an intermediate part and the method further comprises machining the intermediate part to remove excess material and form the aerofoil structure.

Claims

1. A method of manufacturing an aerofoil structure for a gas turbine engine, wherein the aerofoil structure comprises a root configured to be received in a rotor disc (240) of the gas turbine engine, wherein the method comprises: providing a pre-formed insert; adding the insert into a mould for forming the aerofoil structure; adding a composite constituent into the mould; and heating the composite constituent in the mould to bond the insert to the composite constituent, the insert being provided at a flank of the aerofoil structure root that faces a shoulder of the rotor disc; wherein the melting temperature of the insert is higher than the melting temperature of the resin.

2. The method of claim 1, wherein heating the composite constituent in the mould to bond the insert to the composite constituent forms an intermediate part, and the method further comprises machining the intermediate part to remove excess material and form the aerofoil structure.

3. The method of claim 1, wherein the composite constituent is pre-impregnated with a resin, and heating the composite constituent in the mould thermosets the resin.

4. The method of claim 1, wherein the composite constituent comprises carbon fibre pre-impregnated with a resin.

5. The method of claim 1, wherein the method further comprises: adding a further insert into the mould, the insert and the further insert being provided on either side of the composite constituent.

6. The method of claim 1, wherein the method further comprises: machining the insert prior to adding the insert into the mould.

7. The method of claim 1, wherein the method further comprises: adding an adhesive layer between the insert and the composite constituent.

8. The method of claim 1, wherein the insert is formed from a polymer.

9. (canceled)

10. An aerofoil structure for a gas turbine engine, wherein the aerofoil structure comprises a root configured to be received in a rotor disc of the gas turbine engine, wherein the aerofoil structure is formed from a composite constituent and an insert moulded together, the insert being provided at a flank of the root of the aerofoil structure that faces a shoulder of the rotor disc.

11. The aerofoil structure claim 10, comprising a further insert being provided at a further flank of the root of the aerofoil structure that faces a further shoulder of the rotor disc.

12. The aerofoil structure claim 10, wherein the insert is made of a polymer.

13. The aerofoil structure of claim 10, comprising an adhesive layer between the insert and the composite constituent to enhance adhesion between the insert and the composite constituent.

14. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of the aerofoil structures according to claim 10.

15. The gas turbine engine of claim 14, wherein the gas turbine engine further comprises: 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.

16. The gas turbine engine of claim 14, 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.

17. (canceled)

18. The method of claim 8, wherein the polymer comprises polyimide.

19. The aerofoil structure claim 12, wherein the polymer comprises polyimide.

20. The aerofoil structure claim 11, wherein at least one of the insert and the further insert is made of a polymer.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0056] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0057] FIG. 1 is a sectional side view of a gas turbine engine;

[0058] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0059] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0060] FIG. 4 is a flowchart depicting a method of manufacturing an aerofoil structure for the gas turbine engine;

[0061] FIG. 5 is a partial schematic view of a mould for manufacturing the aerofoil structure;

[0062] FIG. 6 is a schematic view of a root of the intermediate part from the mould; and

[0063] FIG. 7 is a schematic view of the aerofoil structure when installed in a rotor.

DETAILED DESCRIPTION

[0064] 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.

[0065] 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.

[0066] 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.

[0067] 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.

[0068] 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.

[0069] 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.

[0070] 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.

[0071] 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.

[0072] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0073] 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 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.

[0074] 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.

[0075] With reference to FIG. 4 the present disclosure relates to a method 100 of manufacturing an aerofoil structure 200 (partially shown in FIG. 7), such as a blade of fan 23. The method 100 comprises a first step 110 in which a pre-formed insert 210a (shown in FIG. 5) is provided. In a second step 120, the insert 210a is added into a mould 220 (shown in FIG. 5) for forming the aerofoil structure 200. In a third step 130, a composite constituent 230 (shown in FIG. 5) is added into the mould 220. Finally, in a fourth step 140 the composite constituent 230 is heated in the mould 220 to bond the insert 210a to the composite constituent 230.

[0076] FIG. 5 depicts at least an end of the mould 220 and a root end of the aerofoil structure 200 during manufacture. The mould 220 may split along axis 222 or along any other line to provide first and second mould parts 220a, 220b. The insert 210a may be placed into the mould 220, in particular the first mould part 220a. A further insert 210b may be placed into the mould 220, in particular the second mould part 220b. The inserts 210a, 210b may then form integral parts of the mould 220 ready to receive the composite constituent 230. The inserts 210, 210b may be pre-moulded to substantially the required shape before placement in the mould 220. The inserts 210, 210b may have been moulded in a different mould to mould 220 prior to insertion into mould 220.

[0077] The composite constituent 230 may then be placed on either the insert 210a or further insert 210b. For example, the method may comprise laying up the composite constituent into either or both of the mould parts 220a, 220b. The two mould parts 220a, 220b may be brought together. Accordingly, the insert 210a and further insert 210b may be provided either side of the composite constituent 230. The composite constituent 230 may form an elongate structure extending from the root to a tip of the aerofoil structure 200.

[0078] The inserts 210a, 210b may be machined prior to adding the inserts into the mould 220. For example, the inserts 210, 210b may be roughened, e.g. grit-blasted, prior to adding the inserts into the mould. In particular, the surfaces of the inserts 210a, 210b that face the composite constituent 230 in the mould 220 may be roughened. The inserts 210, 210b may also be degreased prior to placement in the mould 220. Such treatments may improve the subsequent adhesion to the composite constituent 230. However, to further enhance adhesion, the method may further comprise adding an adhesive layer between each of the inserts 210a, 210b and the composite constituent 230.

[0079] Once the inserts 210a, 210b and composite constituent 230 are in place in the mould 220, the components are heated, e.g. as part of an autoclave cure process. The composite constituent 230 may be pre-impregnated with a resin and heating the composite constituent in the mould 220 may thermoset the resin. In particular, the composite constituent 230 may comprise carbon fibre pre-impregnated with the resin. The resin (together with the adhesive layer if provided) may bond the inserts 210a, 210b to the composite constituent 230. Conveniently, the inserts 210a, 210b may be bonded to the composite constituent 230 in the same process as in which the composite constituent 230 is thermoset.

[0080] To avoid the inserts 210a, 210b deforming, the melting temperature of the inserts may be higher than the temperatures encountered in the mould 220 and higher than the melting temperature of the resin. The inserts 210a, 210b may be formed from a wear resistant material. By way of example, the inserts 210a, 210b may be formed from a polymer, such as Polyimide. In particular, the inserts 210a, 210b may be formed from 420X Polyimide provided by Icon Polymer. It is also envisaged that the inserts 210a, 210b may be formed from a non-polymer material, such as a metal or a ceramic or any other suitable material.

[0081] Referring to FIG. 6, once the heating and bonding has completed, an intermediate part 232 may be removed from the mould 220. The method 100 may further comprise machining the intermediate part 232 to remove excess material and form the aerofoil structure 200. For example, the intermediate part 232 may be machined along dotted line 234 to provide the required shape of the root of the aerofoil structure 200. The machining may remove portions of the inserts 210a, 210b and/or the composite constituent 230.

[0082] As shown, the inserts 210a, 210b may be elongate. The inserts may be curved and a midline (or either surface) of the inserts 210a, 210b may have a point of inflection between ends of the inserts 210a, 210b. Furthermore, the inserts 210a, 210b may taper at an end furthest from the root of the aerofoil structure 200 such that the inserts blend into the composite constituent 230. The dotted line 234 (along which the intermediate part 232 may be machined) may extend lengthwise along at least a portion of each insert 210a, 210b. As such, the machining process may reduce the thickness of the inserts 210a, 210b along at least a portion of their length.

[0083] FIG. 7 depicts the resulting aerofoil structure 200 with its root 202 received in a rotor disc 240, which may be made from titanium or an alloy thereof. As shown, the insert 210a and further insert, 210b are provided on a respective flank and further flank of the root 202 and facing a respective shoulder 242a and further shoulder 242b of the rotor disc 240. In particular, the inserts 210a, 210b may contact the shoulders 242a, 242b where the machining line 234 extends lengthwise along at least a portion of each insert 210a, 210b (i.e. where the thickness of the inserts 210a, 210b has been reduced). Machining in the area where the inserts 210a, 210b contact the shoulders 242a, 242b may help to provide the desired shape at the interface between the aerofoil structure 200 and the rotor disc 240.

[0084] The integrally formed inserts 210a, 210b advantageously provide an interface between the composite constituent 230 and the rotor disc 240. The inserts are conveniently provided during the manufacturing process of the composite constituent 230 and a separate layer is not required between the aerofoil structure 200 and the rotor disc 240. This reduces the cost and speeds up the manufacturing process. The inserts 210a, 210b may also be more readily repaired, e.g. by machining a worn insert to conform to a desired shape.

[0085] 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.