PROCESS OF FORMING AN AEROFOIL

Abstract

A process of forming an aerofoil is provided. The process includes: providing a layered, planar pre-form; inflating and hot creep forming the pre-form to form an intermediate structure having aerofoil pressure and suction surfaces, and a front, edge-receiving portion joining front edges of the pressure and suction surfaces; providing a leading edge piece; and bonding the leading edge piece to the front, edge-receiving portion of the intermediate structure to form an aerofoil in which the leading edge piece forms an aerofoil leading edge.

Claims

1. A process of forming an aerofoil, the process including: providing a layered, planar pre-form; inflating and hot creep forming the pre-form to form an intermediate structure having aerofoil pressure and suction surfaces, and a front, edge-receiving portion joining front edges of the pressure and suction surfaces; providing a leading edge piece; and bonding the leading edge piece to the front, edge-receiving portion of the intermediate structure to form an aerofoil in which the leading edge piece forms an aerofoil leading edge.

2. The process of forming an aerofoil according to claim 1 which further includes: machining the front edge-receiving portion preparatory to bonding the leading edge piece thereto.

3. The process of forming an aerofoil according to claim 1, wherein the intermediate structure from the hot creep forming and inflating further has an aerofoil trailing edge.

4. The process of forming an aerofoil according to claim 1, wherein the intermediate structure from the hot creep forming and inflating further has a rear, edge-receiving portion joining rear edges of the pressure and suction surfaces, and wherein the process further includes: providing a trailing edge piece; and bonding the trailing edge piece to the rear, edge-receiving portion of the intermediate structure to form an aerofoil in which the trailing edge piece forms an aerofoil trailing edge.

5. A process of forming an aerofoil, the process including: providing a layered, planar pre-form; inflating and hot creep forming the pre-form to form an intermediate structure having aerofoil pressure and suction surfaces, and a rear, edge-receiving portion joining rear edges of the pressure and suction surfaces; providing a trailing edge piece; and bonding the trailing edge piece to the rear, edge-receiving portion of the intermediate structure to form an aerofoil in which the trailing edge piece forms an aerofoil trailing edge.

6. The process of forming an aerofoil according to claim 4 which further includes: machining the rear, edge-receiving portion preparatory to bonding the trailing edge piece thereto.

7. The process of forming an aerofoil according to claim 5, wherein the intermediate structure from the hot creep forming and inflating further has an aerofoil leading edge.

8. The process of forming an aerofoil according to claim 1, wherein the leading edge piece and/or the trailing edge piece are bonded to the respective edge-receiving portion by fusion welding or diffusion bonding.

9. The process of forming an aerofoil according to claim 1, wherein the layered, planar pre-form, and the leading edge piece and/or the trailing edge piece are formed of titanium alloy.

10. The process of forming an aerofoil according to claim 1, wherein the aerofoil is an outlet guide vane of a fan of a ducted fan gas turbine engine.

11. An aerofoil formed by the process according to claim 1.

12. A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: the engine has a row of aerofoils according to claim 11.

13. The gas turbine engine according to claim 12, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); the engine core further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) 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

[0057] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0058] FIG. 1 is a schematic, part-sectional side view of a gas turbine engine;

[0059] FIG. 2 is a close up schematic, part-sectional side view of an upstream portion of a gas turbine engine;

[0060] FIG. 3 is a partially cut-away axial view of a gearbox for a gas turbine engine;

[0061] FIGS. 4A to 4E schematically illustrate stages in a conventional superplastic inflating and hot creep forming process for forming an outlet guide vane; and

[0062] FIGS. 5A to 5F schematically illustrate stages in a process for forming an outlet guide vane according to the present disclosure.

[0063] 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, containing fan OGVs 50. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0064] 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.

[0065] 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.

[0066] 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.

[0067] 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.

[0068] 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.

[0069] 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.

[0070] 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.

[0071] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0072] 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.

[0073] 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.

[0074] FIGS. 4A to 4E schematically illustrate stages in a conventional superplastic inflating and hot creep forming process for forming fan OGVs. In FIG. 4A two flat titanium alloy plates are laid on top of each other. In FIG. 4B, front and rear portions joining respectively front and rear edges of the plates are fused together, excess material may then be removed e.g. by waterjet cutting. In FIG. 4C the fused plates are placed over a cambered former and subjected to SPF inflation and HCF in suitable pressing equipment to form an intermediate structure in which one of the plates provides an aerofoil pressure surface and the other an aerofoil suction surfaces. An elliptic leading edge (dashed line in FIG. 4D) is then formed directly on the fused front edges of the plates by adaptive machining to complete the aerofoil shape of the OGV (FIG. 4E). The aerofoil trailing edge may be similarly machined from the fused rear edges, or the fused rear edges of the intermediate structure may be in substantially final form after the SPF and HCF stage.

[0075] As previously mentioned however, this conventional process provides relatively poor control of vane inlet angle at the aerodynamically sensitive leading edge and, despite the machining stage, relies substantially on the cambered former and the pressing equipment for the shape of the leading edge such that modifications and variations to the vane are expensive to make.

[0076] Accordingly, FIGS. 5A to 5F schematically illustrate stages in an improved process for forming the fan OGVs 50 of the engine of FIGS. 1 to 3. In this process, the OGVs are not solely manufactured using an SPF/HCF process with a machined leading edge ellipse. Instead the aerofoil is manufactured at a reduced chord (using SPF/HCF as before), but with an edge-receiving portion which can conveniently be machined into the front fused edges of the plates to receive a separate leading edge piece manufactured under a different process with greater accuracy than can be achieved by the adaptive machining of the conventional process.

[0077] Thus FIG. 5A shows two flat titanium alloy plates laid on top of each other, but with reduced chordal length relative to the plates of FIG. 4A. Similarly to the stages of FIGS. 4B and 4C, in FIG. 5B, front and rear portions joining respectively front and rear edges of the plates are fused together, and in FIG. 5C the fused plates are placed over a cambered former and subjected to SPF inflation and HCF in the pressing equipment to form an intermediate structure in which one of the plates provides an aerofoil pressure surface and the other an aerofoil suction surfaces. As shown in FIG. 5D, a front edge-receiving portion (indicated by dashed lines) is then machined from the fused front edges of the plates. Referring to a FIG. 5E, a separately manufactured leading edge piece is bonded to the front edge-receiving portion, e.g. by fusion welding or diffusion bonding. The finished OGV is shown in FIG. 5F.

[0078] The greater accuracy and consistency of leading edge shape which this process provides can improve overall engine performance and reduce fuel burn. In addition, the leading edge may be more susceptible to in-service repair. A further advantage of the process is that it is relatively easy to change the shape of the leading edge, providing vane design flexibility without a need for expensive new camber standards. For example, complex geometric features can be relatively easily included at the leading edge, e.g. for noise mitigation.

[0079] Although described above in relation to shaping of the leading edge, the same approach (i.e. providing an edge-receiving portion and bonding a separately manufactured edge piece thereto) can applied to the trailing edge of the OGV. Also, although described above in relation to a fan OGV, the process can be applied to other aerofoil structures such as compressor stator vanes and compressor OGVs.

[0080] 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.